2010
NASA’S CONTRIBUTIONS TO
Copyright © 2010 by the National Aeronautics and Space Administration

The opinions expressed in this volume are those of the authors and do not 

necessarily reflect the official position of the United States Government or of 

the National Aeronautics and Space Administration.
NASA’S CONTRIBUTIONS TO

VOLUME 1

AERODYNAMICS

STRUCTURES

PROPULSION

CONTROLS

DR. RICHARD P. HALLION, EDITOR

National Aeronautics and Space Administration 

Headquarters 

300 E St SW

Washington, DC 20546

2010

NASA/SP-2010-570-Vol 1

www.nasa.gov
 Library of Congress Cataloging-in-Publication Data

NASA’s contributions to aeronautics : aerodynamics, structures, 

propulsion, and controls / Richard P. Hallion, editor.

 v. cm.

 Includes bibliographical references and index.

 “NASA/SP-2010-570.”

 1. United States. National Aeronautics and Space Administration--

History. 2. Aeronautics--Research--United States--History. I. Hallion, 

Richard.

 TL521.312.N3825 2010

 629.130973--dc22

 2009044645
Richard P. Hallion ............................................................................................ 2

Jeremy Kinney ............................................................................................... 88

John F. Ward ............................................................................................... 134

Lawrence R. Benson ..................................................................................... 180

T.A. Heppenheimer ...................................................................................... 276

Robert G. Hoey ........................................................................................... 360

John D. Anderson, Jr. ................................................................................... 426

David C. Aronstein ...................................................................................... 460

T.A. Heppenheimer ...................................................................................... 568

Contents

Foreword ...................................................................................................... vii

CASE

1

CASE

2

CASE

3

CASE

6

CASE

4

CASE

7

CASE

5

CASE

8

CASE

9

Sweep and Swing: Reshaping the Wing for the Jet and Rocket Age

Richard Whitcomb and the Quest for Aerodynamic Efficiency 

NACA–NASA and the Rotary Wing Revolution 

Softening the Sonic Boom: 50 Years of NASA Research 

Toward Transatmospheric Flight: From V-2 to the X-51 

Physical Problems, Challenges, and Pragmatic Solutions 

NASA and the Evolution of Computational Fluid Dynamics 

NASA and Computational Structural Analysis 

High-Temperature Structures and Materials 
Fly-by-Wire: Making the Electric Jet 

Albert C. Piccirillo ....................................................................................... 630

Advancing Propulsive Technology 

James Banke ............................................................................................... 734

Leaner and Greener: Fuel Efficiency Takes Flight 

Caitlin Harrington ...................................................................................... 784

Good Stewards: NASA’s Role in Alternative Energy

Bruce I. Larrimer ......................................................................................... 822

Index ......................................................................................................... 879

CASE

10

CASE

11

CASE

12

CASE

13
vii

Foreword

AS THIS BOOK GOES TO PRESS, the National Aeronautics and 

Space Administration (NASA) has passed beyond the half century mark, its longevity a tribute to how essential successive 

Presidential administrations—and the American people whom they 

serve—have come to regard its scientific and technological expertise. In 

that half century, flight has advanced from supersonic to orbital velocities, the jetliner has become the dominant means of intercontinental 

mobility, astronauts have landed on the Moon, and robotic spacecraft 

developed by the Agency have explored the remote corners of the solar 

system and even passed into interstellar space. 

Born of a crisis—the chaotic aftermath of the Soviet Union’s space 

triumph with Sputnik—NASA rose magnificently to the challenge of the 

emergent space age. Within a decade of NASA’s establishment, teams 

of astronauts would be planning for the lunar landings, first accomplished with Neil Armstrong’s “one small step” on July 20, 1969. Few 

events have been so emotionally charged, and none so publicly visible 

or fraught with import, as his cautious descent from the spindly little Lunar Module Eagle to leave his historic boot-print upon the dusty 

plain of Tranquillity Base.

In the wake of Apollo, NASA embarked on a series of space initiatives that, if they might have lacked the emotional and attention-getting 

impact of Apollo, were nevertheless remarkable for their accomplishment and daring. The Space Shuttle, the International Space Station, 

the Hubble Space Telescope, and various planetary probes, landers, rovers, and flybys speak to the creativity of the Agency, the excellence of its 

technical personnel, and its dedication to space science and exploration.

But there is another aspect to NASA, one that is too often hidden in 

an age when the Agency is popularly known as America’s space agency 

and when its most visible employees are the astronauts who courageously 
NASA’s Contributions to Aeronautics

viii

rocket into space, continuing humanity’s quest into the unknown. 

That hidden aspect is aeronautics: lift-borne flight within the atmosphere, 

as distinct from the ballistic flight of astronautics, out into space. 

It is the first “A” in the Agency’s name and the oldest-rooted of the 

Agency’s technical competencies, dating to the formation, in 1915, of 

NASA’s lineal predecessor, the National Advisory Committee for 

Aeronautics (NACA). It was the NACA that largely restored America’s 

aeronautical primacy in the interwar years after 1918, deriving the airfoil 

profiles and configuration concepts that defined successive generations 

of ever-more-capable aircraft as America progressed from the subsonic 

piston era into the transonic and supersonic jet age. NASA, succeeding the NACA after the shock of Sputnik, took American aeronautics 

across the hypersonic frontier and onward into the era of composite 

structures, electronic flight controls, and energy-efficient flight.

This volume, the first of a two-volume set, traces contributions by NASA and the post–Second World War NACA to the field of 

aeronautics. It was that work that enabled the exploitation of the 

turbojet and high-speed aerodynamic revolution that led to the gasturbine-powered jet age that followed, within which we still live. 

The subjects covered in this first volume are an eclectic mix of surveys, case studies, and biographical examinations ranging across 

multiple disciplines and technical competencies residing within the 

National Aeronautics and Space Administration. The topics are indicative of the range of Agency work and the capabilities of its staff. 

They include: 

• 

• 

• 

The advent of the sharply swept-back wing, which enabled 

taking fullest advantage of the turbojet revolution and 

thereby launched the era of high-speed global mass 

mobility, becoming itself the iconic symbol of the jet age.

The contributions and influence of Richard T. Whitcomb, 

a legendary NACA–NASA researcher who gave to aeronautics some of the key methods of reducing drag and 

improving flight efficiencies in the challenging transonic 

region, between subsonic and supersonic flight.

The work of the NACA and NASA in furthering the rotary 

wing revolution via research programs on a range of 

rotorcraft from autogiros through helicopters, convertiplanes, ducted fan, tilt wing, and tilt rotor craft.
Foreword

ix

• 

• 

• 

• 

• 

• 

• 

How NASA worked from the earliest days of the supersonic revolution to mitigate the shock and disturbing effects of the sonic boom, developing creative test 

approaches to evaluate boom noise and overpressures, and then methods to alleviate boom formation 

and impingement, leading to novel aircraft shaping 

and methods that are today promising to revolutionize 

the design of transonic and supersonic civil and 

military aircraft.

How the NACA and NASA, having mastered the transonic and supersonic regions, took on the challenge of 

extending lift-borne flight into the hypersonic region 

and thence into space, using exotic “transatmospheric” 

vehicles such as the legendary X-15, various lifting 

bodies, and the Space Shuttle, and extending the frontiers of air-breathing propulsion with the Mach 9+ 

scramjet-powered X-43.

The physical problems and challenges that forced NASA 

and other researchers to study and find pragmatic solutions for such thorny issues as aeroelasticity, oscillatory 

instabilities forcing development of increasingly sophisticated artificial stability systems, flight simulation for 

high-performance aerospace vehicles, and aerothermodynamic structural deformation and heating.

NASA’s role in advancing and maturing computational 

fluid dynamics (CFD) and applying this new tool to aeronautical research and aerospace vehicle design.

The exploitation of materials science and development 

of high-temperature structures to enable design of practical high-speed military and civil aircraft and spacecraft.

The advent of computerized structural loads prediction, 

modeling, and simulation, which, like CFD, revolutionized aerospace design practices, enhancing both safety 

and efficiency.

NASA’s pioneering of electronic flight control (“fly-bywire”), from rudimentary testbeds evolved from Apolloera computer architectures and software, to increasingly 

sophisticated systems integrating aerodynamic and propulsion controls.
NASA’s Contributions to Aeronautics

x

• 

• 

• 

How the NACA and NASA advanced the gas turbine 

revolution, producing more efficient engine concepts 

and technology for application to new generations of 

military and civilian aircraft.

How NASA has contributed to the quest for fuel-efficient 

and environmentally friendly aircraft technology, studying combustion processes, alternative fuels, and pollutant transfer into the upper atmosphere, searching for 

appropriate technological solutions, and resulting in less 

polluting, less wasteful, and more efficient aircraft designs.

The Agency’s work in promoting global environmental 

good stewardship by applying its scientific and technical 

competencies to wind and solar energy, resulting in more 

efficient energy-producing wind turbines and high-altitude 

solar-powered long-endurance unpiloted aerial vehicles.

The record of NACA–NASA accomplishments in aeronautics demonstrates the value of consistent investment in aeronautical research as 

a means of maintaining the health and stability of America’s aerospace 

industrial base. That base has generated an American predominance in 

both civil and military aeronautics, but one that is far from assured as 

the Nation enters the second century of winged flight. It is hoped that 

these studies, offering a glimpse at the inner workings of the Agency and 

its personnel, will prove of value to the men and women of NASA, to 

those who benefit across the United States and overseas from their dedicated work, and to students of aeronautics and members of the larger 

aerospace community. It is to the personnel of NASA, and the NACA 

before them, that this volume is dedicated, with affection and respect.

Dr. Richard P. Hallion

August 4, 2010
THIS PAGE INTENTIONALLY BLANK
2

The X-48B subscale demonstrator for the Blended Wing-Body (BWB). The BWB may represent 

the next extension of the swept and delta wing, to transform flight away from the rule of the 

“tube and wing” jetliner. NASA.
3

1 Sweep and Swing: 

Reshaping the Wing for 

the Jet and Rocket Age

Richard P. Hallion

T

HE PROGRESSIVE EVOLUTION OF AIRCRAFT DESIGN HAS WITNESSED 

continuous configuration changes, adaptations, and reinterpretations. The canard wood-and-fabric biplane launched the powered 

flight revolution and gave way to the tractor biplane and monoplane, and 

both gave way to the all-metal monoplane of the interwar era. The turbojet engine set aside the piston engine as the primary motive power for 

long-range commercial and military aircraft, and it has been continually 

refined to generate the sophisticated bypass turbofans of the present era, 

some with afterburning as well. The increasing airspeed of aircraft drove 

its own transformation of configuration, measurable in the changed relationship between aspect and fineness ratios. Across the primacy of the 

propeller-driven era, from the beginning of the 20th century to the end of 

the interwar era, wingspan generally far exceeded fuselage length. That 

changed early in the jet and rocket era. By the time military and test pilots 

from the National Advisory Committee for Aeronautics (NACA) first probed 

the speed of sound with the Bell XS-1 and Douglas D-558-1 Skystreak, 

wingspan and fuselage length were roughly equal. Within a decade, as 

aircraft speed extended into the supersonic regime, the ratio of wingspan 

to fuselage length dramatically reversed, evidenced by aircraft such as 

the Douglas X-3, the Lockheed F-104 Starfighter, and the Anglo-French 

Concorde Supersonic Transport (SST). Nicknames handily captured the 

CASE 1

1

The development of the swept and delta wing planform enabled 

practical attainment of the high speeds promised by the invention of 

the turbojet engine and the solid-and-liquid-fueled rocket. Refining the 

swept and delta planforms from theoretical constructs to practical realities involved many challenges and problems requiring creative analysis 

and study by NACA and NASA researchers. Their insight and perseverance led to the swept wing becoming the iconic symbol of the jet age.
NASA’s Contributions to Aeronautics

4

1 transformation: the rakish X-3 was known informally as the “Stiletto” and 

the only slightly less sleek F-104 as the “Missile with a Man in It.”

There was as well another manifestation of profound design transformation, one that gave to the airplane a new identity that swiftly became 

a global icon: the advent of the swept wing. If the biplane constituted the 

normative airplane of the first quarter century of flight and the straight 

wing cantilever monoplane that of the next quarter century, by the time of 

the golden anniversary of Kitty Hawk, the swept wing airplane had supplanted both, its futuristic predominance embodied by the elegant North 

American F-86 Sabre that did battle in “MiG Alley,” high over North Korea’s 

blue-gray hills bordering the Yalu River. In the post-Korean era, as swept 

wing Boeing 707 and Douglas DC-8 jet airliners replaced what historian 

Peter Brooks termed the “DC-4 generation” of straight wing propeller-driven 

transports, the swept wing became the iconic embodiment of the entire jet 

age.1 Today, 75 years since its enunciation at an international conference, 

the high-speed swept wing is the commonly accepted global highway symbol for airports, whether an intercontinental center such as Los Angeles, 

Frankfurt, or Heathrow; regional hubs such as Dallas, Copenhagen, or 

Charlotte; or any of the myriad general aviation and business aviation airfields around the world, even those still primarily populated, ironically, by 

small, straight wing propeller-and-piston-driven airplanes.

The Tailless Imperative: The Early History of Swept and Delta Wings

The high-speed swept wing first appeared in the mid-1930s and, like most 

elements in aircraft design, was European by birth. But this did not mark 

the swept wing’s first appearance in the world’s skies. The swept wing dated 

to before the First World War, when John Dunne had developed a series 

of tailless flying wing biplanes using the swept planform as a means of 

ensuring inherent longitudinal stability, imparting “self-correcting” restoration of any gust-induced pitching motions. Dunne’s aircraft, while 

freakish, did enjoy some commercial success. He sold manufacturing 

1. Peter W. Brooks, The Modern Airliner: Its Origins and Development (London: Putnam & Co., Ltd., 

1961), pp. 91–111. Brooks uses the term to describe a category of large airliner and transport aircraft 

defined by common shared design characteristics, including circular cross-section constant-diameter 

fuselages, four-engines, tricycle landing gear, and propeller-driven (piston and turbo-propeller), from the 

DC-4 through the Bristol Britannia, and predominant in the time period 1942 through 1958. Though 

some historians have quibbled with this, I find Brooks’s reasoning convincing and his concept of such a 

“generation” both historically valid and of enduring value.
Case 1 | Sweep and Swing: Reshaping the Wing for the Jet and Rocket Age

5

rights to the Burgess Company in the United States, which subsequently 

produced two “Burgess-Dunne” seaplanes for the U.S. Navy. Lt. Holden 

C. Richardson, subsequently one of the first members of the NACA, had 

urged their purchase “so that the[ir] advantages and limitations can be 

thoroughly determined . . . as it appears to be only the beginning of an 

important development in aeronautical design.”2

That it was, though not in the fashion Richardson expected. The swept 

wing remained an international staple of tailless self-stabilizing design, 

typified in the interwar years by the various Westland Pterodactyl aircraft 

designed by Britain’s G.T.R. Hill, the tailless aircraft of Boris Ivanovich 

Cheranovskiy, Waldo Waterman’s Arrowplane, and a series of increasingly sophisticated sailplanes and powered aircraft designed by Germany’s 

Alexander Lippisch. However, it would not become the “mainstream” element of aircraft design its proponents hoped until applied to a very different purpose: reducing transonic aerodynamic effects.3

 The transonic 

swept wing effectively increased a wing’s critical Mach number (the “drag 

divergence Mach number”), delaying the onset of transonic drag rise and 

enabling an airplane to fly at higher transonic and supersonic speeds for 

the same energy expenditure and drag penalty that a straight wing airplane 

would expend and experience at much lower subsonic speeds.

In 1935, leading aerodynamicists gathered in Rome for the Volta 

Congress on High Speeds in Aviation, held to coincide with the opening of 

Italy’s impressive new Guidonia laboratory complex. There, a young German 

fluid dynamicist, Adolf Busemann, unveiled the concept of using the swept 

wing as a means of attaining supersonic flight.4 In his presentation, he 

2. Quoted in Roy A. Grossnick, et al., United States Naval Aviation 1910–1995 (Washington: 

U.S. Navy, 1997), p. 15; Gordon Swanborough and Peter M. Bowers, United States Navy 

Aircraft Since 1911 (New York: Funk & Wagnalls, 1968), p. 394.

3. Alexander Lippisch, “Recent Tests of Tailless Airplanes,” NACA TM-564 (1930), a NACA translation 

of his article “Les nouveaux essays d’avions sans queue,” l’Aérophile (Feb. 1–15, 1930), pp. 35–39.

4. For Volta, see Theodore von Kármán and Lee Edson, The Wind and Beyond: Theodore von 

Kármán, Pioneer in Aviation and Pathfinder in Space (Boston: Little, Brown and Co., 1967), pp. 

216–217, 221–222; Adolf Busemann, “Compressible Flow in the Thirties,” Annual Review of Fluid 

Mechanics, vol. 3 (1971), pp. 6–11; Carlo Ferrari, “Recalling the Vth Volta Congress: High Speeds in 

Aviation,” Annual Review of Fluid Mechanics, vol. 28 (1996), pp. 1–9; Hans-Ulrich Meier, “Historischer Rückblick zur Entwicklung der Hochgeschwindigkeitsaerodynamik,” in H.-U. Meier, ed., Die 

Pfeilflügelentwicklung in Deutschland bis 1945 (Bonn: Bernard & Graefe Verlag, 2006), pp. 16–36; 

and Michael Eckert, The Dawn of Fluid Dynamics: A Discipline Between 

Science and Technology (Weinheim: Wiley-VCH Verlag, 2006), pp. 228–231.

1
NASA’s Contributions to Aeronautics

6

1 demonstrated the circulation pattern around a swept wing that, essentially, “fooled” it into “believing” it was flying at lower velocities. As well, 

he presented a sketch of an aircraft with such a “Pfielförmiges Tragwerk” 

(“Arrow-Shaped Lifting Surface”), though one that had, by the standards 

of subsequent design, very modest sweep and very high aspect ratio.5

Theodore von Kármán recalled not quite two decades later that afterward, at the conference banquet, “General [Arturo] Crocco, the organizer of the congress and a man of far-reaching vision, went further while 

doodling on the back of the menu card, drawing a plane with sweptback wings and tail, and even swept propeller blades, laughingly calling 

it ‘Busemann’s airplane.’”6 Evidence exists that Crocco took the concept 

beyond mere dinner conversation, for afterward, Guidonia researchers 

evaluated a design blending modestly swept wings with a “push-pull” 

twin-engine fuselage configuration. However, Guidonia soon returned 

to the more conventional, reflecting the Italian air ministry’s increasing emphasis upon building a large and powerful air arm incorporating 

already proven and dependable technology.7

Delegates from other nations present at Busemann’s briefing missed its 

significance altogether, perhaps because his gently swept configuration—in 

the era of the DC-2 and DC-3, which had pronounced leading edge taper—

looked far less radical than the theory and purpose behind it implied. 

NACA Langley Memorial Aeronautical Laboratory researchers had already 

evaluated far more sharply swept planforms at Langley for a seminal 

wing taper study the laboratory issued the next year.8 Thus, at first glance, 

Busemann’s design certainly did not look like a shape that would transform aviation from the firmly subsonic to the transonic, making possible 

the potential of the jet engine, and the jet age (with its jet set) that followed.

5. Adolf Busemann, “Aerodynamische Auftrieb bei Überschallgeschwindigkeit,” Luftfahrtforschung, 

vol. 12, No. 6 (Oct. 3, 1935), pp. 210–220, esp. Abb. 4–5 (Figures 4–5).

6. Theodore von Kármán, Aerodynamics (New York: McGraw-Hill Book Company, Inc., 1963 ed.), p. 133.

7. Ministero dell’Aeronautica, 1° Divisione, Sezione Aerodinamica Resultati di Esperienze (Rome: 

Guidonia, 1936); the swept “double-ender” wind tunnel study (anticipating the layout of Dornier’s 

Do 335 Pfeil [“Arrow”] of the late wartime years) was designated the J-10; its drawing is dated 

March 7, 1936. I thank Professor Claudio Bruno of the Università degli Studi di Roma “La Sapienza”; and Brigadier General Marcello di Lauro and Lieutenant Colonel Massimiliano Barlattani 

of the Stato Maggiore dell’Aeronautica Militare (SMdAM), Rome, for their very great assistance in 

enabling me to examine this study at the Ufficio Storico of the SMdAM in June 2009.

8. Raymond F. Anderson, “Determination of the Characteristics of Tapered Wings,” NACA Report 

No. 572 (1936); see in particular Figs. 15 and 16, p. 11.
Case 1 | Sweep and Swing: Reshaping the Wing for the Jet and Rocket Age

7

Therefore, for the United States and most other nations, over the 

next decade, the normative airplane remained one having straight (if 

tapered) wings and piston propulsion. For Germany, however, the future 

belonged to increasingly sharply swept and delta wings, and jet and 

rocket propulsion as well. Within 5 years of the Volta conference, with 

Europe engulfed in a new war, its engineers had already flown their 

first jet and rocket-powered aircraft, had expanded beyond Busemann’s 

initial conception to derive shapes more closely anticipating subsequent 

high-speed aircraft and missile designs, and were busily testing models 

of swept wing transonic airplanes and supersonic missiles. Lippisch’s 

swept wing sailplanes had presaged a new Messerschmitt rocketpropelled interceptor, the Me 163 Komet (“Comet”), and his broad, high 

aspect ratio deltas had given way to a rounded triangular planform 

that he envisioned as meeting the needs for transonic and supersonic 

flight. While many of these concepts by Lippisch and other German 

designers were impracticable, or unrelated to Germany’s more immediate military needs, others possessed significant military or research 

potential. Only flawed decisions by the Third Reich’s own leadership 

and the Allies’ overrunning of Germany would prevent them from 

being developed and employed before the collapse of the Hitler regime 

in May 1945.9

Birthing the American Delta and Swept Wing

The extent to which the swept wing permeated German aeronautical thought understandably engendered tremendous postwar interest 

9. For an example of such work, see Dr. Richard Lehnert, “Bericht über Dreikomponentenmessungen 

mit den Gleitermodellen A4 V12/a und A4 V12/c,” Archiv Nr 66/34 (Peenemünde: HeeresVersuchsstelle, Nov. 27, 1940), pp. 6-10, Box 674, “C10/V-2/History” file, archives of the National 

Museum of the United States Air Force, Dayton, OH. Re: German research deficiencies, see Adolf 

Baeumker, Ein Beitrag zur Geschichte der Führung der deutschen Luftfahrttechnik im ersten halben Jahrhundert, 1900–1950 (Bad Godesberg: Deutschen Forschungs – und Versuchsanstalt für Luft – und Raumfahrt 

e. V., 1971), pp. 61–74; Col. Leslie E. Simon, German Scientific Establishments (Washington: Office 

of Technical Services, Department of Commerce, 1947), pp. 7–9; Helmuth Trischler, “Self-Mobilization 

or Resistance? Aeronautical Research and National Socialism,” and Ulrich Albrecht, “Military Technology 

and National Socialist Ideology,” in Monika Renneberg and Mark Walker, eds., Science, Technology, 

and National Socialism (Cambridge: Cambridge University Press, 1994), pp. 72–125. For science and 

the Third Reich more generally, see Alan D. Beyerchen, Scientists Under Hitler: Politics and the Physics 

Community in the Third Reich (New Haven: Yale University Press, 1977); and Kristie Macrakis, Surviving 

the Swastika: Scientific Research in Nazi Germany (New York: Oxford University Press, 1993).

1
NASA’s Contributions to Aeronautics

8

1

A sampling of various design concepts for Lippisch swept wing and delta aircraft. These original Lippisch sketches were incorporated in “German Aircraft: New and Projected Types,” a 

1946 Allied technical intelligence summary. USAF.

in the benefits of swept planforms for transonic and supersonic flight 

within the American, European, and Soviet aeronautical communities.10 However, for America, uncovering German swept wing research 

and development furnished the confirmation of its value, not its

discovery, for Robert T. Jones, an aerodynamicist at the Langley Memorial 

Aeronautical Laboratory, had independently discovered its benefits in 

1944, a year before the Allies first entered Germany’s shattered and shuttered research laboratories and design shops.11

The Gluhareff-Griswold Nexus

In 1936, Michael E. Gluhareff, an emigree Russian engineer who was 

chief of design for the Vought-Sikorsky Aircraft Division of United 

10. USAAF, “German Aircraft, New and Projected Types” (1946), Box 568, “A-1A/Germ/1945” 

file, NMUSAF Archives; and J. McMasters and D. Muncy, “The Early Development of Jet Propelled 

Aircraft,” AIAA Paper 2007-0151, Pts. 1–2 (2007).

11. See Richard P. Hallion, “Lippisch Gluhareff, and Jones: The Emergence of the Delta Planform 

and the Origins of the Sweptwing in the United States,” Aerospace Historian, vol. 26, no. 1 (Mar. 

1979), pp. 1–10.
Case 1 | Sweep and Swing: Reshaping the Wing for the Jet and Rocket Age

9

Aircraft Corporation, began examining various tailless aircraft configurations. By July 1941, his study had spawned a proposed interceptor fighter powered by a piston engine driving a contra-rotating pusher 

propeller. It had a rounded delta planform resembling an arrowhead, 

with leading edges swept aft at 56 degrees. It featured a tricycle retractable landing gear, twin ventral vertical fins, an extremely streamlined 

and rounded configuration, provisions for six heavy machine guns, and 

elevons (combined ailerons and elevators) for roll and pitch control. 

Gluhareff informed company founder Igor I. Sikorsky that its sharp 

sweep would delay the onset of transonic compressibility, noting, “The 

general shape and form of the aircraft is, therefore, outstandingly adaptable for extremely high speeds.”12

In retrospect, Gluhareff’s design was a remarkable achievement, 

conceived at just the right time to have been completed with turbojet 

propulsion (for which its configuration and internal layout was eminently suited) though circumstances conspired against its development. 

Sikorsky was then perfecting the first practical helicopter—the VS-300, 

another revolutionary development, of course—and chose understandably to concentrate on rotary wing flight. He did authorize Gluhareff to 

solicit support from inventor-entrepreneur Roger W. Griswold, president of the Ludington-Griswold Company, about building a wind tunnel 

model of the configuration.13 Tests by United Aircraft proved so encouraging that Griswold approached the engineering staff of the Army Air 

Forces (AAF) at its Wright Field Aircraft Laboratory about sponsoring what was now called the “Dart.”14 But having their fill of visitors 

12. Memo, Michael Gluhareff to I.I. Sikorsky, July 1941, copy in the Gluhareff Dart accession 

file, National Air and Space Museum, Smithsonian Institution, Washington, DC. Gluhareff’s Dart 

appeared contemporaneously with a remarkably similar (though with a tractor propeller) Soviet 

design by Alexandr Sergeevich Moskalev. Though unclear, it seems Gluhareff first conceived the 

planform. It is possible that an informal interchange of information between the two occurred, as 

Soviet aeronautics and espionage authorities kept close track of American developments and the 

activities of the emigree Russian community in America.

13. Griswold is best known as coinventor (with Hugh De Haven) of the three-point seat restraint, 

which formed the basis for the modern automotive seat belt; Saab then advanced further, building 

upon their work. See “Three-Point Safety Belt is American, not Swedish, Invention,” Status Report, 

vol. 35, no. 9 (Oct. 21, 2000), p. 7.

14. Vought-Sikorsky, “Aerodynamic Characteristics of the Preliminary Design of a 1/20 Scale Model 

of the Dart Fighter,” Vought-Sikorsky Wind Tunnel Report No. 192 (Nov. 18, 1942), copy in the Gluhareff Dart accession file, National Air and Space Museum, Smithsonian Institution, Washington, DC.

1
NASA’s Contributions to Aeronautics

10

1

The proposed Gluhareff Dart fighter of 1941, showing both its novel layout and, for the time, nearly 

as novel tricycle landing gear layout. National Air and Space Museum, Smithsonian Institution.

bringing a series of the weird and unconventional, and charged with 

ensuring that the AAF acquired large numbers of aircraft, and quickly, 

the AAF’s engineers did not pursue the project.15

So the Gluhareff-Griswold Dart never reached the hardware stage, the 

failure to build it counting as a loss to American midcentury aeronautics. 

As for Gluhareff, though he had made notable contributions to Sikorsky’s 

large flying boats (and would, as well, to his helicopters), he continued 

15. Letter, Roger W. Griswold to Maj. Donald R. Eastman, Oct. 22, 1946, Gluhareff Dart accession file, NASM.
Case 1 | Sweep and Swing: Reshaping the Wing for the Jet and Rocket Age

11

to explore the basic design of his intriguing if abortive configuration, 

proposing a variety of derivatives, including in 1959 a Mach 2+ supersonic transport with a small canard wing and double-deck fuselage.16 If 

the Dart never saw development, its configuration nevertheless proved 

significant. In 1944, Griswold resurrected the Dart shape for a proposed 

2,000-pound guided glide bomb, or “glomb.” The Army Air Forces recommended he obtain the NACA’s opinion of its aerodynamics, and for this, 

Griswold turned to Langley Memorial Aeronautical Laboratory. There, 

on August 19, he met with the NACA’s resident aerodynamic expert on 

“pilotless missiles,” Robert T. Jones. Out of that contact would emerge 

both the American delta and swept wing.

Enter Robert T. Jones . . .

“R.T.” Jones was a brilliant, flight-obsessed, and largely self-taught fluid 

dynamicist, having dropped out of the University of Missouri to join a 

flying circus, then working as a designer for Nicholas-Beazley, a small 

Missouri aircraft company. When the Great Depression collapsed the 

firm, his father used political connections as Chairman of the local 

Democratic Party to secure Jones a job running elevators in the U.S. 

Capitol. In his spare time and evenings, he studied mathematics and 

aerodynamics with Albert Zahm, the aeronautics Chair at the Library of 

Congress, and with Max Munk at Catholic University. Despite his lack 

of a formal engineering degree, through the efforts of Representative 

David Lewis (a homespun Maryland progressive with a strong interest 

in self-improvement who had taken math instruction from the young 

elevator operator), Jones received a temporary appointment as a “scientific aide” to the NACA. There, he quickly proved such a gifted and 

insightful researcher that he soon secured a coveted permanent position at Langley, consorting with the likes of John Stack, Eastman Jacobs, 

and Theodore Theodorsen.17

As he considered Griswold’s “glomb,” Jones recognized that its 

extremely low aspect ratio shape (that is, a shape having a very long 

16. M.E. Gluhareff, “Tailless Airplane,” U.S. patent No. 2,511,502, issued June 13, 1950; 

“Sikorsky Envisions Supersonic Airliner,” Aviation Week (May 4, 1959), pp. 67–68; M.E. Gluhareff, 

“Aircraft with Retractable Auxiliary Airfoil,” U.S. patent No. 2,941,752, issued June 21, 1960.

17. See William Sears’s biographical introduction to the “Collected Works of Robert T. Jones,” 

NASA TM-X-3334 (1976), pp. vii–ix; and Walter G. Vincenti, “Robert Thomas Jones,” in 

Biographical Memoirs, vol. 86 (Washington: National Academy of Sciences, 2005), pp. 3–21.

1
NASA’s Contributions to Aeronautics

12

1 wing root in relation to its total wingspan) could not be adequately analyzed using conventional Prandtl-rooted “lifting line” theory. Instead, 

Jones drew on the work of his mentor Munk, using papers that Munk had 

written on the flow of air around inclined airship hulls and swept wings, 

and one by the Guggenheim Aeronautical Laboratory’s Hsue-shen Tsien, a 

von Kármán associate at the California Institute of Technology (Caltech), 

on airflow around inclined bodies of revolution. He analyzed it using 

linear equations governing two-dimensional incompressible flow, considering his results of little practical value, recalling three decades later, 

“I thought, well, this is so crude, nobody would be interested. So I 

just hid it in my desk.”18

But it sparked his curiosity, and in January 1945, by which time he 

was busy thinking about nonlinear compressible flows, he had a revelation: the equations he had developed months earlier for the glomb 

analysis could be applied to a low aspect triangular wing operating in 

supersonic flow, one whose wing-leading edges were so sharply swept 

as to place them within the shock cone formed around the vehicle and 

hence operating in subsonic flow. In these conditions, the wing was 

essentially “fooled” into behaving as if it were operating at a much lower 

Mach number. As Jones recalled, “It finally dawned on me that the slender wing theory would hold for compressible flow and even at supersonic 

speed if it were near the center of the Mach cone. So, I immediately got 

the paper out and I added the compressible flow parts to it, which was 

really the important part, and then I wondered well, why is it that this 

slender wing doesn’t have an effect on compressibility? Then I realized 

that it was because the obliquity of the edge and that this is the simple sweep theory and would work in spite of the compressibility effect. 

So, I wrote a paper which incorporated the slender wing theory and 

also sweep theory.”19 Jones then moved from considering a slender 

triangular delta [Δ] to the sharply sweptback wing [^], the reverse of 

18. Transcript of interview of R.T. Jones by Walter Bonney, Sept. 24, 1974, p. 5, in Jones biographical file, No. 001147, Archives of the NASA Historical Division, National Aeronautics and 

Space Administration, Washington, DC.

19. Transcript of Jones-Bonney interview, p. 5; Hallion conversation with Dr. Robert T. Jones at 

NASA Ames Research Center, Sunnyvale, CA, July 14, 1977; Max M. Munk, “The Aerodynamic 

Forces on Airship Hills, NACA Report No. 184 (1923); Max M. Munk, “Note on the Relative 

Effect of the Dihedral and the Sweep Back of Airplane Wings,” NACA TN-177 (1924); H.S. Tsien, 

“Supersonic Flow Over an Inclined Body of Revolution,” Journal of the Aeronautical Sciences, vol. 5, 

no. 2 (Oct. 1938), pp. 480–483.
Case 1 | Sweep and Swing: Reshaping the Wing for the Jet and Rocket Age

13

Germany, where the high-speed swept wing had preceded, not followed, 

the delta.20

Jones’s delta and swept wing utilized, for their time, very thin airfoil sections, ones typical of supersonic aircraft to come. In contrast, German swept 

and delta wing developer Alexander Lippisch had employed much thicker 

sections that proved unsuitable for transonic flight. His tailless rocketpropelled swept wing Me 163 Komet (“Comet”) interceptor, for example, 

essentially became uncontrollable at speeds slightly above Mach 0.82 thanks 

to stability changes induced by shock wave formation on its relatively thick 

wing. His design for a rocket-boosted, ramjet-powered delta fighter, the 

P 13, had such thick wing and tail sections—the pilot actually sat within 

the leading edge of the vertical fin—that it could never have achieved its 

desired transonic performance. As discussed subsequently, postwar NACA 

tests of a captured glider configuration of this design, the DFS DM-1, 

confirmed that transonic delta wings should be far thinner, with sharper 

leading edges. As a consequence, NACA researchers rejected the Lippisch 

approach, and, though some of them tried extrapolations of his designs (but 

with lower thickness-chord ratios and sharper leading edges), the NACA 

(and industry as well) adapted instead the thin slender delta, à la Jones.21

Dissemination, Deliberation, and Confirmation

In February 1945, Jones showed his notes on sweep to Jean Roché, the 

Army Air Forces technical liaison at Langley, and informed others as 

well, including Maj. Ezra Kotcher of the AAF’s Air Technical Service 

20. Note that although Lippisch called his tailless aircraft “deltas” as early as 1930, in fact they 

were generally broad high aspect ratio wings with pronounced leading edge taper, akin to the 

wing planform of America’s classic DC-1/2/3 airliners. During the Second World War, Lippisch 

did develop some concepts for sharply swept deltas (though of very thick and impracticable wing 

section). Taken all together, Lippisch’s deltas, whether of high or low aspect ratio planform, were not 

comparable to the thin slender and sharply swept (over 60 degrees) deltas of Jones, and Gluhareff 

before him, or Dietrich Küchemann at the Royal Aircraft Establishment afterwards, which were more 

akin to high-supersonic and hypersonic shapes of the 1950s–1960s.

21. For DM-1 and extrapolative tests, see Herbert A. Wilson, Jr., and J. Calvin Lowell, “Full-Scale 

Investigation of the Maximum Lift and Flow Characteristics of an Airplane Having Approximately 

Triangular Plan Form,” NACA RM-L6K20 (1947); J. Calvin Lovell and Herbert A. Wilson, Jr., 

“Langley Full-Scale-Tunnel Investigation of Maximum Lift and Stability Characteristics of an Airplane 

Having Approximately Triangular Plan Form (DM-1 Glider),” NACA RM-L7F16 (1947); and Edward 

F. Whittle, Jr., and J. Calvin Lovell, “Full-Scale Investigation of an Equilateral Triangular Wing Having 

10-Percent-Thick Biconvex Airfoil Sections,” NACA RM-L8G05 (1948).

1
NASA’s Contributions to Aeronautics

14

1 Command, and NACA colleagues Arthur Kantrowitz and Hartley A. 

Soulé.22 Kotcher passed it along to von Kármán and Tsien—then working as scientific advisers to Gen. Henry H. “Hap” Arnold, the Army Air 

Forces’ Chief of Staff—and Soulé and Kantrowitz urged Jones to inform 

the Agency’s Director of Research, George W. Lewis, of his discovery.23

Accordingly, on March 5, 1945, Jones informed Lewis, “I have recently 

made a theoretical analysis which indicates that a V-shaped wing travelling point foremost would be less affected by compressibility than other 

planforms. In fact, if the angle of the V is kept small relative to the Mach 

angle [the angle of the shockwave], the lift and center of pressure remain 

the same at speeds both above and below the speed of sound.”24 Jones 

subsequently undertook tests in the Langley 9-inch supersonic tunnel of 

a small, 4-inch-long daggerlike sheet-steel triangular wing with rounded 

leading edges and a span of only 1.5 inches, tests that complemented 

other trials at Aberdeen, MD, arranged by von Kármán and Tsien.

The Langley tests, through the transonic region and up to Mach 1.75, 

confirmed his expectations, and Jones published his first test results 

May 11, 1945, noting, “The lift distribution of a pointed airfoil travelling 

point-foremost is relatively unaffected by the compressibility of the air 

below or above the speed of sound.”25 This was almost 2 weeks before 

Lippisch informed von Kármán, then leading an AAF European study 

team, of his high-speed delta concepts (during a technical intelligence 

interrogation at St. Germain, France, on May 23), not quite a month 

before von Kármán assistant Clark Millikan visited the Messerschmitt 

advanced projects group at Oberammergau on June 9–10 and interrogated Waldemar Voigt about his swept wing fighter concepts, and 

well over a month before Millikan journeyed to Völkenrode to inter22. In 1944, Kotcher had conceived a rocket-powered “Mach 0.999” transonic research airplane 

(a humorous reference to the widely accepted notion of an “impenetrable” sonic “barrier”) that 

subsequently inspired the Bell Aircraft Corporation to undertake design of the XS-1, the world’s first 

supersonic manned airplane.

23. Kantrowitz would pioneer high-Mach research facilities design, and Soulé would serve the NACA 

as research airplane projects leader, supervising the Agency’s Research Airplane Projects Panel (RAPP), 

a high-level steering group coordinating the NACA’s X-series experimental aircraft programs.

24. Memo, Jones to Lewis, Mar. 5, 1945; see also ltr., Jones to Ernest O. Pearson, Jr., Feb. 2, 1960, 

and Navy/NACA Record of Invention Sheet, Apr. 10, 1946, Jones biographical file, NASA.

25. Robert T. Jones, “Properties of Low-Aspect-Ratio Pointed Wings at Speeds Below and Above the 

Speed of Sound,” NACA TN-1032 (1946), p. 11 [first issued at NACA LMAL on May 11, 1945].
Case 1 | Sweep and Swing: Reshaping the Wing for the Jet and Rocket Age

15

rogate German swept wing inventor Adolf Busemann, on June 20–21.26

Langley’s peer reviewers and senior Agency official Theodore 

Theodorsen did not immediately accept Jones’s assumption that a unified 

slender wing theory could apply to both compressible and incompressible 

flows and even questioned the evidence of sweep’s benefits. Fortunately, 

Jones was greatly assisted in confounding skeptics by the timely results of 

NACA tunnel tests and falling body experiments, which left little doubt that 

sweep worked. As well, an associate of Jones made a most helpful discovery: locating a 1942 British translation of Busemann’s 1935 paper. Evidence 

of an enemy’s interest coincident with one’s own work always heightens its 

perceived value, and undoubtedly, the Busemann paper, however dated, 

now strongly bolstered Jones’s case. When it became time to assemble a 

bibliography for his swept wing report, Jones added Busemann’s paper 

and other German sources by Albert Betz, H.G. Küssner, Ludwig Prandtl, 

and Hermann Schlichting, though it is unclear whether this reflected a 

collegial respect across the chasm of war or simply a shrewd appreciation of their persuasive value.27

Langley released his report in late June 1945.28 In it, Jones noted: “the 

attachment of plane waves to the airfoil at near-sonic or supersonic speeds 

(Ackeret theory) may be avoided and the pressure drag may be reduced 

by the use of planforms in which the angle of sweepback is greater than 

the Mach angle. The analysis indicates that for aerodynamic efficiency, 

wings designed for flight at supersonic speeds should be swept back at an 

26. For Millikan visit to Germany, see Millikan Diary 6, Box 35, Papers of Clark B. Millikan, 

Archives, California Institute of Technology, Pasadena, CA; Alexander Lippisch, ltr. to editor, Aviation 

Week and Space Technology (Jan. 6, 1975); in 1977, while curator of science and technology 

at the National Air and Space Museum, the author persuaded Jones to donate his historic delta test 

model to the museum; he had been using it for years as a letter opener!

27. Jones noted afterward that at Volta, Busemann “didn’t have the idea of getting the wing inside the 

Mach cone so you got subsonic flow. The real key to [the swept wing] was to get subsonic flow at supersonic speed by getting the wing inside the Mach cone . . . the development of what I would say [was] 

the really correct sweep theory for supersonic speeds occurred in Germany in ’43 or ’44, and with me in 

1945.” (See transcript of Jones-Bonney interview, p. 6). But German researchers had mastered it earlier, 

as evident in a series of papers and presentations in a then-“Geheim” (“Secret”) conference report by 

the Lilienthal-Gesellschaft für Luftfahrtforschung, Allgemeine Strömungsforschung: Bericht über die Sitzung 

Hochgeschwindigkeitsfragen am 29 und 30 Oktober 1942 in Berlin (Berlin: LGF, 1942).

28. For his report, see Robert T. Jones, “Wing Planforms for High-Speed Flight,” NACA TN-1033 

(1946) [first issued at LMAL on June 23, 1945, as Confidential Memorandum Report L5F21]. Jones’s 

tortuous path to publication is related in James R. Hansen’s Engineer in Charge: A History of the Langley Aeronautical Laboratory, 1917–1958, SP-4305 (Washington: NASA, 1987), pp. 284–285.

1
NASA’s Contributions to Aeronautics

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1

Jones showed these notes on the concept of high-speed wing sweep to Langley’s AAF technical 

liaison representative Jean Roché on February 27, 1945. NASA.

angle greater than the Mach angle and the angle of sweepback should be 

such that the component of velocity normal to the leading edge is less than 

the critical speed of the airfoil sections. This principle may also be applied 

to wings designed for subsonic speeds near the speed of sound, for which 

the induced velocities resulting from the thickness might otherwise be sufficiently great to cause shock waves.”29 Such marked the effective birth of 

29. Jones, “Wing Planforms for High-Speed Flight,” NACA TN-1033, p. 1.
Case 1 | Sweep and Swing: Reshaping the Wing for the Jet and Rocket Age

17

the high-speed swept wing airplane in the United States, as his report 

weeks earlier had marked the birth of the American high-speed delta.

By the time Jones’s report appeared, Germany’s aeronautical establishment was already under the microscope of Allied technical intelligence, whose teams swiftly focused on its intensive investment in swept 

wing aerodynamics for its missiles and aircraft. Replicating reaction 

to the earlier “discovery” of “Göttingen aerodynamics” after the First 

World War, the post–Second World War influence of German example 

and practice was even more profound. Indeed, it affected the entire postwar course of European, Soviet, and American high-speed aerodynamic 

research, development, test, evaluation, and acquisition. In the increasingly tense national security environment of the burgeoning Cold War, 

the national intelligence services of the various advanced aeronautical nations understandably maintained very active technical collection 

efforts to learn what they did not already know.30

30. For the United States, this meant that Soviet intelligence collectors increasingly focused on American 

high-speed research. Bell Aircraft Corporation, manufacturer of the first American jet airplane, the first 

supersonic airplane, and advanced swept wing testbeds (the X-2 and X-5), figured prominently as a 

Soviet collection target as did the NACA. NACA engineer William Perl (born Mutterperl), a member 

of the Rosenberg spy ring who passed information on aviation and jet engines to Soviet intelligence, 

worked as a postwar research assistant for Caltech’s Theodore von Kármán, director of the Guggenheim 

Aeronautical Laboratory of the California Institute of Technology (GALCIT), the Nation’s premier academic 

aero research facility. He cultivated a close bond with TvK’s sister Josephine (“Pipa”) and TvK himself. Perl 

had almost unique access to the highest-level NACA and GALCIT reports on high-speed flight, and the 

state of advanced research and facilities planning for them and the U.S. Air Force. He associated as well 

with NACA notables, including Arthur Kantrowitz, Eastman Jacobs, and Robert T. Jones. So closely was 

he associated with von Kármán that he once helpfully reminded him where to find the combination to an 

office safe! He helped screen sensitive NACA data for a presentation TvK was making on high-speed 

stability and control, and TvK recommended Perl for consultation on tunnel development at the proposed 

new Arnold Engineering Development Center (AEDC) in Tennessee. Perl was unmasked by the Venona 

signals intelligence decryption program, interrogated on his associations with known Communists, and 

subsequently arrested and convicted of perjury. (He had falsely denied knowing the Rosenbergs.) More 

serious espionage charges were not brought, lest court proceedings compromise the ongoing Venona

collection effort. The Papers of Theodore von Kármán, Box 31, Folder 31.38, Archives of the California 

Institute of Technology, and the Federal Bureau of Investigations’ extensive Perl documentation contain 

much revealing correspondence on Perl and his associates. I thank Ernest Porter and the FBI historical 

office for arranging access to FBI material. See also Katherine A.S. Sibley, Red Spies in America: Stolen 

Secrets and the Dawn of the Cold War (Lawrence: University Press of Kansas, 2004); and John Earl 

Haynes and Harvey Klehr’s Early Cold War Spies: The Espionage Trials that Shaped American Politics

(Cambridge: Cambridge University Press, 2006) for further details on the Perl case.

1
NASA’s Contributions to Aeronautics

18

1

The North American XP-86, prototype of the F-86 Sabre family, represented an amalgam of 

German and American swept wing and streamlined aerodynamics. USAF.

Swept Wing Challenges

The NACA so rapidly focused its attention on swept planforms that, 

within 2 years of the end of the Second World War, George Gray, author 

of a popular yet surprisingly detailed study of the Agency, could already 

write: “Just how far the sweepback principle can be applied with resulting advantage is a question. . . . At about 90 percent of the speed of 

sound both sweepback and low aspect ratio begin to be of value, and 

wings that combine the two features seem to offer a promising choice. 

At about Mach number 1.50, a sweepback of 60 degrees seems necessary to escape the backward flare of the Mach angle. . . . At Mach number 2.00, the angle is so acute that it is impossible to avoid it and still 

preserve the wings. It may be that designers preparing for flight at this 

speed will return to wings of low angles of sweep, and place their main 

dependence for drag reduction on thinning the profiles, lowering the 

aspect ratio, and sharpening the edges of wings.”31 By 1950, this grow31. George W. Gray, Frontiers of Flight: The Story of NACA Research (New York: Knopf, 1948), 

p. 348.
Case 1 | Sweep and Swing: Reshaping the Wing for the Jet and Rocket Age

19

ing confidence in the old-new swept planform had resulted in transonic 

and supersonic research airplanes, a variety of military prototypes, and 

two operational jet fighters that would shortly clash over North Korea: 

the American F-86 Sabre (first flight in October 1947) and, in the Soviet 

Union, the MiG-15 (first flight in December 1947).32

Swept wing aircraft, for all their high-speed advantages, posed daunting stability, control, and handling qualities challenges. Foremost of 

these was pitch-up at low and high speeds, resulting from deteriorating 

longitudinal stability.33 A swept wing airplane’s lateral-directional stability was compromised as well by so-called “dihedral effects.” Swept 

wing aircraft with excessive dihedral experienced pronounced combined 

rolling and yawing “Dutch roll” motions, which would be unacceptable 

on both production civil and military designs.34 Such motions would 

induce airsickness in passengers on large aircraft and, on bomber, fighter, 

and attack aircraft, prevent accurate tracking of a maneuvering target 

or accurate bomb release. (Indeed, it was largely because of this kind 

of behavior that the U.S. Air Force did not proceed with production of 

Northrop’s YB-49 flying wing jet bomber.) Adverse yaw posed another 

problem. At higher speeds, as a swept wing plane rolled from aileron 

32. Re: German high-speed influence in the U.S., Britain, and Russia, see H.S. Tsien, “Reports on 

the Recent Aeronautical Developments of Several Selected Fields in Germany and Switzerland,” in 

Theodore von Kármán, ed., Where We Stand: First Report to General of the Army H.H. Arnold on 

Long Range Research Problems of the Air Forces with a Review of German Plans and Developments 

(Washington: HQ AAF, Aug. 22, 1945), Microfilm Reel 194, Papers of Gen. Henry H. Arnold, 

Manuscript Division, U.S. Library of Congress, Washington, DC; Ronald Smelt, “A Critical Review 

of German Research on High-Speed Airflow,” Journal of the Royal Aeronautical Society, vol. 50, 

No. 432 (Dec. 1946), pp. 899–934; Andrew Nahum, “I Believe the Americans Have Not Yet 

Taken Them All!” in Helmuth Trischler, Stefan Zeilinger, Robert Bud, and Bernard Finn, eds., Tackling 

Transport (London: Science Museum, 2003), pp. 99–138; Matthew Uttley, “Operation ‘Sturgeon’ 

and Britain’s Post-War Exploitation of Nazi German Aeronautics,” Intelligence and National Security, vol. 17, no. 2 (Sum. 2002), pp. 1–26; M.I. Gurevich, “O Pod’emnoi Sile Strelovidnogo Kryla 

v Sverkhzvukovom Potoke,” Prikladnaya Matematika i Mekhanika, vol. 10 (1946), translated by the 

NACA as “Lift Force of an Arrow-Shaped Wing,” NACA TM-1245 (1949). Gurevich, cofounder 

of the MiG bureau (he is the “G” in “MiG”) was subsequently principal aerodynamicist of the MiG15, the Soviet Union’s swept wing equivalent to the American F-86. For a detailed examination 

of F-86 wing development and the influence of German work (particularly Göthert’s) upon it, see 

Morgan M. Blair, “Evolution of the F-86,” AIAA Paper 80-3039 (1980).

33. Pitch-up was of such significance that it is discussed subsequently in greater detail within this essay.

34. First comprehensively analyzed by Max M. Munk in his “Note on the Relative Effect of the 

Dihedral and the Sweep Back of Airplane Wings,” NACA TN-177 (1924).

1
NASA’s Contributions to Aeronautics

20

1 deflection, it experienced higher drag and loss of lift involving the lowered wing, generating a tendency of the airplane to turn (reverse) into 

the direction of the raised wing, effectively doing the opposite of what 

the pilot intended. Adverse yaw could be caused by aeroelastic effects 

as well. That swept wing aircraft would possess behavior characteristics significantly different than conventional straight wing designs did 

not come as a surprise to the NACA or other aerodynamic researchers 

in America and overseas. But all recognized the need to complement 

theory and ground-test methodologies with flight research.

The peculiarities of swept wing aircraft, at a time when early jet aircraft lacked the power-to-weight advantages of later designs, could—and 

often did—prove fatal. For example, Boeing designed the B-47, America’s 

first large swept wing aircraft, with pod-mounted engines and a broad, 

highly tapered, thin swept wing. During flight-testing at higher speeds, 

test pilots found aileron input to roll the aircraft would twist the wing, 

the aileron effectively acting as a trim-tab does on a control surface. The 

twisted wing would overcome the rolling moment produced by the aileron, rolling the aircraft in the opposite direction. Aeroelastic structural 

divergence caused several accidents of the B-47 during its flight-testing 

and service introduction, forcing the Air Force to limit its permissible 

airspeed to 425 knots, as high as it could be safely flown if roll reversal 

were to be avoided. As a result, Boeing built its successors, the XB-52 

and the Model 367-80 (prototype for the KC-135 family and inspiration 

for the civil 707), with much thicker wing roots and structures that were 

torsion resistant but that could still flex vertically to absorb structural 

loads and gust-induced loads during flight.35

Confronting Pitch-Up

But the most serious swept wing problem in the early jet era was pitchup, a condition affecting both low- and high-speed flight, reflecting 

stall onset either from decreasing speed (low-speed pitch-up) or from 

trim changes during high-speed flight, particularly during accelerated 

35. See John E. Steiner, “Transcontinental Rapid Transit: The 367-80 and a Transport Revolution—

The 1953–1978 Quarter Century,” AIAA Paper 78-3009 (1978), p. 93; John E. Steiner, “Jet Aviation Development: A Company Perspective,” in Walter J. Boyne and Donald H. Lopez, eds., The Jet 

Age: Forty Years of Jet Aviation (Washington: Smithsonian Institution Press, 1979), pp. 145–148; 

and William H. Cook, The Road to the 707: The Inside Story of Designing the 707 (Bellevue, WA: 

TYC Publishing Co., 1991), pp. 145–205.
Case 1 | Sweep and Swing: Reshaping the Wing for the Jet and Rocket Age

21

maneuvers, such as “wind-up” turns that rapidly increased g-loading 

and angle of attack. Pitch-up occurred at the breakpoint in a lift curve, 

immediately beyond the peak point where the airplane’s wing was operating at its highest lift-producing angle of attack, with its lift coefficient 

at maximum value. At the breakpoint, the wing would begin stalling, 

with flow separation from the airfoil, breaking the circulatory flow pattern around the wing. In ideal circumstances with a straight wing aircraft, the change in lift would occur simultaneously spanwise across 

the wing and would typically trigger a nose drop. But in a swept wing 

aircraft, the stall would first begin at the tips and progress inward, the 

center of lift shifting forward. As the plane’s longitudinal (nose-up nosedown) stability decreased, the shifting center of lift would abruptly rotate 

the nose upward (hence the use of the expression “pitch-up”), even at a 

rate of onset beyond the capabilities of its elevator control surfaces to 

correct. As well, of course, since the ailerons that governed lateral control (roll control) were typically located outboard on a wing, a swept 

wing airplane could lose its lateral control authority precisely at a point 

when the pilot needed as much control capability and reserve as possible. Because stall onset is not always triggered uniformly, a swept wing 

airplane nearing the pitch-up point could experience sudden loss of lift 

on one wing, inducing abrupt rolling motions (called “wing dropping”), 

complicating its already dangerous low-speed behavior.

 There was, of course, the possibility of overcoming such problems 

by sweeping a wing forward, not aft. A forward-swept wing (FSW) had 

both desirable high – and low-speed aerodynamic characteristics. Since 

the spanwise flow would run from the tips to the fuselage, the outer 

portions of the wing would stall last, thus preserving lateral control. As 

well, it would have more desirable pitching characteristics. Already, in 

the midst of the Second World War, the Germans had flown an experimental bomber, the Junkers Ju 287, featuring a forward swept wing, 

and a number of aircraft and missile projects were forecast for such 

planforms as well. The forward swept wing, and combined-sweep “M,” 

“W,” and even “X” planforms, received a great deal of postwar attention, both in America and abroad. Researchers at Langley modified wind 

tunnel and configuration models of both the XS-1 and D-558 to employ 

forward-swept wing planforms, and tested conceptual planforms with 

both aft and forward sweep to develop comparison data. But while 

the FSW undoubtedly had better low-speed behavior, it had higher 

profile drag and posed difficult structural problems for designers. In 

1
NASA’s Contributions to Aeronautics

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1 the precomposite structure era, an FSW had to be necessarily heavier 

than an aft-swept wing to avoid aeroelastic flexing that could inhibit 

both good flight performance and even flight safety. Further, the structural and weight limitations also limited the sweep angles that an FSW 

could then have; even as late as the 1960s, when Germany produced 

a business aircraft (the Hamburger Flugzeugbau HFB-320 Hansa 

Jet), it possessed only modest forward sweep and, though flown successfully and built in small numbers, was not a commercial success. 

It would take over three decades before the advent of computerized 

flight control, composite structures, and a more radical vision of forward sweep application would result in experimental planforms like the 

Rockwell Sabrebat FSW design concept, the piloted Air Force–DARPA–

NASA Grumman X-29 (and, in Russia, an X-29-like experimental aircraft, 

the Sukhoi Su-37). Even so, and even though forward sweep would be 

applied to some weapon systems (for example, the AGM-129 stealthy 

cruise missile, where it contributed to its low radar reflectivity), 

forward wing sweeping would remain the exception to “normative” 

aft-swept wing design practice.36

 Pitch-up was profoundly dangerous. At low speeds in proximity to 

the ground, it could—and often did—trigger a disastrous departure and 

crash. The recognition of such problems had caused the U.S. Navy to 

procure two modified Bell P-63 Kingcobra fighters (designated L-39), 

which had their wing panels replaced with 35-degree swept wing sections, and a fuselage extension to accommodate their now-changed center of lift. Not intended for high speeds, these two low-speed swept wing 

research aircraft were extensively flown by various contractor, Navy, and 

NACA research pilots to assess the basic behavior of the swept wing, 

with and without lift-and-control-augmenting devices such as wing 

slats and flaps. They quickly encountered its limitations. On one flight 

with the plane in “clean” (i.e., slat-free) configuration, Bell Company 

test pilot A.M. “Tex” Johnston gradually raised the nose of the plane 

while retarding power. After just “a slight tremor” indicating the onset 

of asymmetric tip stall, it “instantaneously rolled to an almost inverted 

36. See, for example, Richard T. Whitcomb, “An Investigation of the Effects of Sweep on the Characteristics of a High-Aspect-Ratio Wing in the Langley 8-Ft. High Speed Tunnel,” NACA RM-L6J01a 

(1947), conclusion 4, p. 19; Stephen Silverman, “The Next 25 Years of Fighter Aircraft,” AIAA 

Paper No. 78-3013 (1978); Glen Spacht, “X-29 Integrated Technology Demonstrator and ATF,” 

AIAA Paper No. 83-1058 (1983).
Case 1 | Sweep and Swing: Reshaping the Wing for the Jet and Rocket Age

23

position.”37 Grumman test pilot Corwin “Corky” Meyer recalled that while 

the L-39 was “docile” with leading edge slats, without them it “cavorted 

like a cat on catnip.”38 The two L-39 aircraft furnished vital insight into 

the low-speed performance and limitations of swept wing aircraft, but 

they also clearly demonstrated that such aircraft could, in fact, be safely 

flown if their wings incorporated careful design and safety devices such 

as fixed leading edge slots or movable slats.39

 In military aircraft, pitch-up could prevent a pilot from maneuvering 

effectively against a foe, could lead to loss of control of the airplane, and 

could result in such excessive airframe loadings that an airplane would 

break up. It was no respecter of designs, even outstanding ones such as 

North American’s evocative F-86 Sabre, generally considered the finest 

jet fighter of its time by both American and foreign test pilots. First flown 

in October 1947, the Sabre quickly became an internationally recognized 

symbol of aeronautical excellence and advancement. When British test 

pilot Roland Beamont, a distinguished Royal Air Force fighter ace of the 

Second World War, evaluated the Sabre at Muroc Dry Lake in May 1948 

37. A.M. “Tex” Johnston with Charles Barton, Tex Johnston: Jet-Age Test Pilot (Washington: Smithsonian Institution Press, 1991), p. 105. The designation “L-39” could be taken to imply that the swept 

wing testbeds were modifications of Bell’s earlier and smaller P-39 Airacobra. In fact, it was coincidence; the L-39s were P-63 conversions, as is evident from examining photographs of the 

two L-39 aircraft.

38. Corwin H. Meyer, Corky Meyer’s Flight Journal: A Test Pilot’s Tales of Dodging Disasters—Just in 

Time (North Branch, MN: Specialty Press, 2006), p. 193.

39. NACA’s L-39 trials are covered in three reports by S.A. Sjoberg and J.P. Reeder: “Flight Measurements of the Lateral and Directional Stability and Control Characteristics of an Airplane Having 

a 35° Sweptback Wing with 40-Percent-Span slots and a Comparison with Wind-Tunnel Data,” 

NACA TN-1511 (1948); “Flight Measurements of the Longitudinal Stability, Stalling, and Lift Characteristics of an Airplane Having a 35° Sweptback Wing Without Slots and With 40-Percent-Span 

Slots and a Comparison with Wind-Tunnel Data,” NACA TN-1679 (1948); and “Flight Measurements of the Stability, Control, and Stalling Characteristics of an Airplane Having a 35° Sweptback 

Wing Without Slots and With 80-Percent-Span Slots and a Comparison with Wind-Tunnel Data,” 

NACA TN-1743 (1948). The American L-39s were matched by foreign equivalents, most notably 

in Sweden, where the Saab company flew a subscale swept wing variant of its conventional Safir

light aircraft, designated the Saab 201, to support development of its J29 fighter, Western Europe’s 

first production swept wing jet, which first flew in Sept. 1948. Like both the F-86 and MiG-15, it 

owed its design largely to German inspiration. Saab researchers were so impressed with what they 

had learned from the 201 that they subsequently flew another modified Safir, the Saab 202, with 

a more sharply swept wing planform intended for the company’s next jet fighter, the J32 Lansen 

(Lance). See Hans G. Andersson, Saab Aircraft Since 1937 (Washington: Smithsonian Institution 

Press, 1989), pp. 106, 117.

1
NASA’s Contributions to Aeronautics

24

1 (a month after it had dived past the speed of sound, becoming the world’s 

first supersonic turbojet airplane), he likewise dived it through Mach 

1, thus becoming the first supersonic British pilot. Afterward, he noted 

approvingly in his test report, “The P-86 is an outstanding aircraft.”40

The Sabre’s reputation was such that British authorities (frustrated by 

the slow development pace of Albion’s own swept wing aircraft) tellingly referred to it simply as “That Aircraft.” Vickers-Supermarine test 

pilot David Morgan recalled, “No British fighter of the day could match 

the handling of the North American F-86.”41 Indeed, designers from 

his company, frustrated by their slow progress turning the experimental Swift into a decent airplane, even resorted to crude subterfuge in 

an effort to unlock the Sabre’s secrets. When a pair of Canadian pilots 

landed at the Supermarine plant in their Canadair-built Sabres, company officials, with apparent generosity, laid on a fancy lunch, driving 

them off to a local hotel. While the visiting airmen dined and chatted 

with solicitous Supermarine representatives, another team of engineers 

“swarmed over the Sabres to study their construction,” marveling at 

“this splendid aircraft.”42

 Yet however “splendid” “That Aircraft” might otherwise have been, 

the Sabre killed unwary pilots by the dozens in accidents triggered by 

its low-speed pitch-up tendencies. Apollo 11 astronaut Michael Collins 

recalled his introduction to the F-86 at Nellis Air Force Base as “a brutal process. . . . In the eleven weeks I was there, twenty-two people were 

killed. In retrospect it seems preposterous to endure such casualty rates 

without help from the enemy, but at the time the risk appeared perfectly 

acceptable. . . . I’m surprised to have survived. I have never felt quite so 

threatened since.”43 In over a decade of tests with various Sabre variants 

40. XP-86 test report, May 21, 1948, reprinted in Roland Beamont, Testing Early Jets: Compressibility and the Supersonic Era (Shrewsbury: Airlife, 1990), p. 36. Beamont’s achievement remained 

largely secret; the first British pilot to fly through the speed of sound in a British airplane was John 

Derry, who did so in Sept. 1948.

41. Quote from Nigel Walpole, Swift Justice: The Full Story of the Supermarine Swift (Barnsley, UK: 

Pen & Sword Books, 2004), p. 38.

42. Charles Burnet, Three Centuries to Concorde (London: Mechanical Engineering Publications 

Ltd., 1979), pp. 121, 123.

43. Michael Collins, Carrying the Fire: An Astronaut’s Journeys (New York: Farrar, Straus, and 

Giroux, 1974), p. 9. Another Sabre veteran who went through Nellis at the same time recalled to 

the author how he once took off on a training sortie with ominous columns of lingering smoke from 

three earlier Sabre accidents.
Case 1 | Sweep and Swing: Reshaping the Wing for the Jet and Rocket Age

25

to improve their low-speed handling qualities, NACA Ames researchers assessed a variety of technical “fixes.” The most beneficial was the 

combination of artificial feel (to give the pilot more reassuring higher 

maneuvering control forces during the approach-to-landing, combined 

with greater inherent stability than possible with a non–artificial-feel system), coupled with leading-edge suction to draw off the boundary layer 

airflow.44 First evaluated on a test rig installed in the Ames 40-foot by 

80-foot full-scale wind tunnel, the boundary layer control (BLC) experiment on the F-86 proved most valuable. Ames researchers concluded: 

“Leading edge boundary-layer control was most effective in providing a 

large reduction in both stalling speed and approach speed together with 

an increased margin of lift for flare and maneuvering during the [landing] approach,” an important point, particularly for swept wing naval 

aircraft, which had to be controllable down to a landing on the confined deck of an aircraft carrier.45 The trials benefitted not only future 

swept wing studies but, more generally, studies of BLC applications for 

Vertical/Short Take-Off and Landing (V/STOL) aircraft systems as well.46

 Nor were the Sabre’s high-speed pitch-up characteristics innocuous. The NACA flew extensive Sabre evaluations at its High-Speed 

Flight Research Station and at Ames to refine understanding of 

44. For development of control boost, artificial feel, and control limiting, see Robert G. Mungall, 

“Flight Investigation of a Combined Geared Unbalancing-Tab and Servotab Control System as Used 

with an All-Movable Horizontal Tail,” NACA TN-1763 (1948); William H. Phillips, “Theoretical 

Analysis of Some Simple Types of Acceleration Restrictors,” NACA TN-2574 (1951); R. Porter 

Brown, Robert G. Chilton, and James B. Whitten, “Flight Investigation of a Mechanical Feel Device 

in an Irreversible Elevator Control System of a Large Airplane,” NACA Report No. 1101 (1952); 

James J. Adams and James B. Whitten, “Tests of a Centering Spring Used as an Artificial Feel 

Device on the Elevator of a Fighter Airplane,” NACA RM-L52G16; and Marvin Abramovitz, 

Stanley F. Schmidt, and Rudolph D. Van Dyke, Jr., “Investigation of the Use of a Stick Force Proportional to Pitching Acceleration for Normal-Acceleration Warning,” NACA RM-A53E21 (1953).

45. George E. Cooper and Robert C. Innis, “Effect of Area-Suction-Type Boundary-Layer Control on 

the Landing-Approach Characteristics of a 35° Swept-Wing Fighter,” NACA RM-A55K14 (1957), 

p. 11. Other relevant Ames F-86 studies are: George A. Rathert, Jr., L. Stewart Rolls, Lee Winograd, 

and George E. Cooper, “Preliminary Flight Investigation of the Wing-Dropping Tendency and LateralControl Characteristics of a 35° Swept-Wing Airplane at Transonic Mach Numbers,” NACA RMA50H03 (1950); and George A. Rathert, Jr., Howard L. Ziff, and George E. Cooper, “Preliminary 

Flight Investigation of the Maneuvering Accelerations and Buffet Boundary of a 35° Swept-Wing 

Airplane at High Altitude and Transonic Speeds,” NACA RM-A50L04 (1951).

46. Edwin P. Hartman, Adventures in Research: A History of the Ames Research Center, 1940–

1965, SP-4302 (Washington: NASA 1970), p. 252.

1
NASA’s Contributions to Aeronautics

26

1 its transonic pitch-up behavior, which test pilot A. Scott Crossfield 

recalled as “violent and dangerous.”47 It could easily exceed its design 

load factors, sometimes pitching as high as 10 g. At 25,000 feet, in the 

very midst of its combat operating envelope (and at lift coefficients less 

than its maximum attainable lift) the Sabre’s pitch-up onset was so severe 

that g forces once momentarily “blacked out” the test pilot. Overall, after 

extensive Ames tests, the early slatted F-86A with a conventional fixed 

horizontal stabilizer and movable elevator was judged “unsatisfactory” by 

a group of highly experienced fighter test pilots, thanks to its “severe pitchup tendencies.” The same group found the later slat-less “6-3” F-86F (socalled because its wing extended forward 6 inches at the root and 3 inches 

at the tip, a modification made by North American based on Korean 

war experience) had “moderate” pitch-up tendencies. Because of this, 

and because it had an adjustable (not fixed) horizontal stabilizer in 

addition to its elevator, the pilots judged the F-86F’s pitch-up behavior 

“unsatisfactory but acceptable.”48

 Worse swept wing problems plagued the Sabre’s great adversary, the 

Soviet MiG-15. Unlike the Sabre, the MiG-15 had a less aerodynamically 

pleasing configuration, and its fixed horizontal stabilizer and elevator 

combination, located midway up the vertical fin, made it more susceptible to aerodynamic “blanketing” of the tail by the wing and, hence 

severe pitch-up problems, as well as limiting its transonic maneuverability (to the Sabre’s advantage). During the Korean war, Sabre pilots 

frequently saw MiG pilots eject from otherwise perfectly sound aircraft 

that had pitched up during turns, stalled, and entered flat, unrecoverable 

spins. Nearly five decades later, Soviet pilot Stepan Mikoyan (nephew of 

Anushavan “Artem” Mikoyan, cofounder of the MiG design bureau) conceded that high-speed accelerated stalls often triggered unrecoverable 

spins, leading to “a number of ejections and fatal accidents.”49 Postwar 

American testing of a MiG-15 delivered by a defecting North Korean 

47. A. Scott Crossfield with Clay Blair, Always Another Dawn: The Story of a Rocket Test Pilot

(Cleveland: World Publishing Co., 1960), pp. 193–194. See also W.C. Williams and A.S. 

Crossfield, “Handling Qualities of High-Speed Airplanes,” NACA RM-L52A08 (1952), p. 3; 

Melvin Sadoff, John D. Stewart, and George E. Cooper, “Analytical Study of the Comparative PitchUp Behavior of Several Airplanes and Correlation with Pilot Opinion,” NACA RM-A57D04 (1957).

48. Sadoff, Stewart, and Cooper, “Analytical Study of Comparative Pitch-Up Behavior,” p. 12.

49. S.A. Mikoyan, Stepan Anastasovich Mikoyan: An Autobiography (Shrewsbury: Airlife, 1999), 

p. 289.
Case 1 | Sweep and Swing: Reshaping the Wing for the Jet and Rocket Age

27

pilot confirmed the MiG’s marked vulnerability to pitch-up-induced 

stalls and spins; indeed, the defector’s own instructor had been lost in 

one such accident. Not surprisingly, when Mikoyan produced the MiG17—the lineal successor to the MiG-15—it had a very different outer 

wing configuration giving it more benign behavior.50

Western European swept wing aircraft exhibited similar problems 

as their American and Soviet counterparts. For a brief while, influenced 

by the Messerschmitt Me 163 Komet and a variety of other German projects, designers were enthralled with the swept wing tailless configuration, believing it could resolve both the challenges of high-speed flight 

and also furnish inherent stability.51 Then, in September 1946, British 

test pilot Geoffrey de Havilland perished in an experimental tailless 

transonic research aircraft, the de Havilland D.H. 108 Swallow, when 

it began an undamped violently divergent longitudinal pitching oscillation at Mach 0.875, breaking up over the Thames estuary and proving 

that the “sound barrier” could bite.52 Subsequently the NACA evaluated 

the Northrop X-4, a generally similar American configuration. Tested 

at high altitude (and hence, at low dynamic pressure), the X-4 fortunately never “diverged” as violently as the ill-fated D.H. 108. Instead, 

as NACA pilot A. Scott Crossfield remembered, at Mach 0.88 “it broke 

50. Maj. Gen. H.E. “Tom” Collins, USAF (ret.), “Testing the Russian MiG,” in Ken Chilstrom, ed., 

Testing at Old Wright Field (Omaha: Westchester House Publishers, 1991), p. 46.

51. Two not so taken with the swept tailless configuration were Douglas aerodynamicist L. Eugene 

Root and former Focke-Wulf aerodynamicist Hans Multhopp. After an inspection trip to Messerschmitt in August 1945, Root wrote “a tailless design suffers a disadvantage of small allowable 

center of gravity travel. . . . Although equally good flying qualities can be obtained in either [tailless 

or conventional] case, the tailless design is considered more dangerous at very high speeds. For 

example, the Me 262 has been taken to a Mach number of 0.86 without serious difficulty, whereas 

the Me 163 could not exceed M = 0.82. For the Me 163 . . . it was not considered possible 

fundamentally to control the airplane longitudinally past M = 0.82 in view of a sudden diving 

moment and complete loss of elevator effectiveness.” See L.E. Root, “Information of Messerschmitt 

Aircraft Design,” Item Nos. 5, 25, File No. XXXII-37, Copy 079 (Aug. 1945), p. 3, Catalog 

D52.1Messerschmitt/144, in the Wright Field Microfilm Collection, National Air and Space 

Museum Archives, Paul E. Garber Restoration Facility, Silver Hill, MD. Focke-Wulf’s Hans Multhopp, 

designer of the influential T-tail sweptwing Ta 183, was even more dismissive. After the war, while 

working at the Royal Aircraft Establishment, he remarked that it constituted an “awful fashion;” see 

Nahum, “I Believe . . .,” in Trischler, et al., ed., Tackling Transport, p. 118. Multhopp later came to 

America, joining Martin and designing the SV-5 reentry shape that spawned the SV-5D PRIME, the 

X-24A, and the X-38.

52. Burnet, Three Centuries to Concorde, p. 102.

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NASA’s Contributions to Aeronautics

28

1 into a steady porpoising motion, like an automobile cushioning over a 

washboard road.”53 Conventional tailed European swept wing designs 

followed the same steep learning curve as American ones. Britain’s 

Supermarine Swift, a much-touted design from the builder of the 

legendary Spitfire, had a “vicious” transonic pitch-up. By the time it 

entered service, it was years late, obsolescent, and useless for any other 

role save low-level tactical reconnaissance.54

The Skyrocket: The NACA’s Pitch-Up Platform

Pitch-up afflicted a wide range of early transonic and supersonic jet 

fighters, and the NACA was fortunate in having an available research 

airplane that could study swept wing behavior across the transonic 

regime. This aircraft was the Douglas D-558-2 Skyrocket, “Phase II” of 

the larger D-558 research aircraft program, a Douglas Company venture begun in 1945 and sponsored by the U.S. Navy and the NACA. The 

D-558 program had begun as a companion to the XS-1 effort and represented a different design approach. Where the XS-1 was rocket powered, 

the D-558 Skystreak used a turbojet; where the XS-1 employed an ogival projectile shape with a midwing of 8-percent thickness-chord ratio, 

the D-558 used a constant-diameter tube wrapped around an axial-flow 

turbojet engine and a low wing of 10-percent thickness-chord ratio; and 

where the XS-1 was air launched, the D-558 took off from the ground 

as a conventional airplane. Both were straight wing designs, with their 

adjustable stabilizers and movable elevators placed midway up their vertical fins. All together, the Navy ordered six D-558 aircraft from the firm.55

Originally, swept wings had not featured in the D-558 program. 

Then the discovery by Douglas engineers of a plethora of German technical reports (coupled with the work of Jones and others in the United 

States) caused the Navy, the NACA, and Douglas to modify the D-558 

53. Crossfield with Blair, Always Another Dawn, p. 39; Melvin Sadoff and Thomas R. Sisk, 

“Longitudinal-Stability Characteristics of the Northrop X-4 Airplane (USAF No. 46-677),” NACA 

RM-A50D27 (1950); and Williams and Crossfield, “Handling Qualities of High-Speed Airplanes.”

54. Quote from Walpole, Swift Justice, pp. 58, 66. Walpole, a former Swift pilot, writes affectionately but frankly of its strengths and shortcomings. See also Burnet, Three Centuries to Concorde, 

pp. 127–128. Burnet was involved in analyzing Swift performance, and his book is an excellent 

review of Supermarine and other British efforts at this time.

55. For the origins of the D-558 program, see Richard P. Hallion, Supersonic Flight: Breaking the 

Sound Barrier and Beyond—The Story of the Bell X-1 and Douglas D-558 (New York: The Macmillan Co. in association with the Smithsonian Institution, 1972).
Case 1 | Sweep and Swing: Reshaping the Wing for the Jet and Rocket Age

29

program.56 The last three aircraft were completed as a new swept wing 

design. Initially, the planned modification seemed straightforward: 

replace the straight wing and tail surfaces with swept ones. In anticipation, Langley tested models of the D-558 with a variety of swept 

wings. But the possibility of giving the swept wing D-558 supersonic 

performance—something the D-558 straight wing lacked—resulted in a 

more radical redesign. Gone was the simple Pitot intake inlet. Instead, 

designer Edward “Ed” Heinemann and his team chose an ogival body 

shape resembling the XS-1. The new 35-degree slat-equipped swept wing 

was relocated to midfuselage position and given anhedral (droop), with 

the landing gear relocated into the fuselage. In contrast to the original 

single-engine D-558s, the new swept wing design featured both a 6,000-

pound thrust rocket engine and a small turbojet. Thus recast, it received 

the designation D-558-2 and the name Skyrocket, to distinguish it from 

the straight wing Skystreak, itself redesignated D-558-1. The result was 

one of the most elegant and significant aircraft of all time.

The first D-558-2 flew in February 1948, though initial flight tests 

gave little hint of how remarkably versatile and successful it would 

prove. At max takeoff weight, it was so underpowered (and thus so sluggish) that it needed four solid-fuel jettisonable assistance takeoff (JATO) 

rockets to help kick it aloft. Eventually, the Navy and the NACA would 

arrange to take the second and third Skyrockets and modify them for air 

launch from a modified PB2-1S (Navy B-29) Superfortress, dramatically 

improving both their safety and high-speed performance; fuel previously 

56. Particularly Bernard Göthert’s “Hochgeschwindigkeitmessungen an einem Pfeilflügel (Pfeilwinkel φ = 35°),” in the previously cited Lilienthal-Gesellschaft, Allgemeine Strömungsforschung, pp. 

30–40, subsequently translated and issued by the NACA as “High-Speed Measurements on a 

Swept-Back [sic] Wing (Sweepback Angle φ = 35°),” NACA TM-1102 (1947), which directly 

influenced design of the 35-degree swept wings employed on the F-86, the B-47, and the D-558-2. 

Göthert, incidentally, used NACA airfoil sections for his studies, another example of the Agency’s 

pervasive international influence. At war’s end he was in Berlin; when ordered to report to Russian 

authorities, he instead fled the city, making his way back to Göttingen, where he met Douglas 

engineer Apollo M.O. Smith, with the Naval Technical Mission to Europe. Smith arranged for him to 

immigrate to the United States, where he had a long and influential career, rising to Chief Scientist 

of Air Force Systems Command, a position he held from 1964 to 1966. See Tuncer Cebeci, ed., 

Legacy of a Gentle Genius: The Life of A.M.O. Smith (Long Beach: Horizons Publishing, Inc., 

1999), p. 32. I acknowledge with grateful appreciation notes and correspondence received 

from members of the D-558 design team in 1971–1972, including the late Edward Heinemann, 

L. Eugene Root, A.M.O. Smith, Kermit Van Every, and Leo Devlin, illuminating the origins of the 

Skystreak and Skyrocket programs.

1
NASA’s Contributions to Aeronautics

30

1 spent climbing aloft could now be more profitably expended exploring the 

transonic and supersonic regimes. While the third aircraft retained its jet 

and rocket engine, the second had its jet engine removed and additional 

tanks for rocket propellant and oxidizer installed. Thus modified, the 

second aircraft reached Mach 2.01 in November 1953, flown by Scott 

Crossfield, the first piloted Mach 2 flight, having earlier attained an altitude of 83,235 feet, piloted by Lt. Col. Marion Carl, a noted Marine aviator. 

Eventually, the NACA received the first D-558-2 as well (which Douglas had 

employed for contractor testing). The Agency modified it as an all-rocket 

aircraft, though it only completed a single check flight before being retired.

Before all-rocket modification, the second Skyrocket introduced 

Agency pilots to the hazards of pitch-up. On August 8, 1949, during its 

seventh flight, pilot Robert Champine banked into a 4 g turn at Mach 

0.6, and the Skyrocket violently pitched up, reaching 6 g. It responded 

rapidly to full-down elevator, and Champine made an uneventful (if prudently precautionary) landing. Thereafter, until returning the airplane 

to Douglas for all-rocket modification in 1950, the NACA flew extensive 

pitch-up investigations with it. In November, pilot John Griffith replicated the 4 g and Mach 0.6 pitch-up that Champine had experienced 

earlier. This time, however, he attempted to continue flying to more fully 

assess the Skyrocket’s behavior. Thus challenged, it snap-rolled on its 

back. After recovering, Griffith probed its low-speed behavior, gradually 

slowing, with flaps and gear extended and wing slats closed. At 14,000 

feet and 130 mph, the Skyrocket pitched up, rolling into a spin, and losing 7,000 feet of altitude before its pilot could recover.57 Clearly its ugly 

behavior did not match its alluring form.

Focused on extending the Skyrocket’s performance into the supersonic regime by modifying the second aircraft as a pure rocket plane, 

the NACA turned to the third aircraft, which retained its jet engine as 

well as its rocket, for future pitch-up research. Air-launched, the jet-androcket Skyrocket had tremendous research productivity; it could accelerate 

57. Hallion, Supersonic Flight, pp. 151–152, based upon D-558 biweekly progress reports. As 

well, I thank the late Robert Champine for his assistance to my research. See also S.A. Sjoberg and 

R.A. Champine, “Preliminary Flight Measurements of the Static Longitudinal Stability and Stalling 

Characteristics of the Douglas D-558-II Research Airplane (BuAero No. 37974),” NACA RML9H31a (1949); W.H. Stillwell, J.V. Wilmerding, and R.A. Champine, “Flight Measurements with 

the Douglas D-558-II (BuAero No. 37974) Research Airplane Low-Speed Stalling and Lift Characteristics,” NACA RM-L50G10 (1950).
Case 1 | Sweep and Swing: Reshaping the Wing for the Jet and Rocket Age

31

Leading-edge wing chord extensions tested on the third D-558-2 Skyrocket, one of many combinations of flaps, slats, fences, and extensions evaluated in the NACA’s 6-year-long study of 

the Skyrocket’s pitch-up behavior. NASA.

into the supersonic regime, above Mach 1.1, and its jet engine enabled it 

to “loiter” in the transonic region, making repeated data-gathering runs. 

Its comprehensive instrumentation package enabled assessment of loads, 

pressure distributions, and accelerations, evaluated against background 

data on flight conditions, aircraft attitude, and control surface position 

and forces. Between the end of 1950 and the fall of 1956, it completed 66 

research flights on pitch-up and associated transonic phenomena, including the evaluation of the effects external wing stores—tanks and bomb 

shapes—had on aircraft performance. It evaluated a variety of proposed 

aerodynamic solutions and fixes to resolve the pitch-up problem, including various wing fence designs to “channel” airflow and inhibit the characteristic spanwise-flow (flow toward the wingtips) found with swept wing 

planforms, various combinations of slat and flap position, changes to leading edge shape, and “sawtooth” leading edge extensions on its outer wing 

panels. All of this testing reinforced what engineers suspected, namely 

that no one overall technical fix existed that could resolve the pitch-up 

challenge. Rather, swept wing aircraft design was clearly situational, and, 

depending on the mission of the aircraft and its resulting design, combinations of approaches worked best, chief among them being low placement of 

1
NASA’s Contributions to Aeronautics

32

1 the horizontal tail, below the chord-line of the wing, coupled with provision 

of stability augmentation and pitch-damping flight control technology.58

Ensuring Longitudinal Control: Transforming the Horizontal Tail

Though not seemingly connected to the swept wing, the researching and 

documenting of the advantages of low-placed horizontal tail surfaces constituted one of the major NACA postwar contributions to flight, one dramatically improving both the safety and flight performance of swept wing 

designs. As a consequence, the jet fighter and attack aircraft of 1958 looked 

very different than did the initial jet (and rocket) aircraft of the immediate postwar era. Then, high-speed aircraft designers had emphasized 

tailless planforms, or ones in which the horizontal tail was well up the 

vertical fin (for example, both the XS-1 and the D-558 families). A decade 

later, aircraft introduced into test or service—such as the Vought F8U-1 

Crusader, the Republic F-105B Thunderchief, the Grumman F11F-1 Tiger, 

the McDonnell F4H-1 Phantom II, the North American A3J-1 Vigilante, 

and the Northrop N-156 (progenitor of both the T-38 supersonic trainer 

and the F-5 lightweight fighter)—shared common characteristics: irreversible power-operated flight controls, stability augmentation, and damping, large vertical fins for enhanced directional stability, area-ruling, and 

low-placed, all-moving tails. Foreign aircraft exhibited similar features: 

for example, the MiG-21, Folland Gnat, and English Electric Lightning.

Aircraft lacking such features manifested often-perilous behavior. 

The Douglas XF4D-1 Skyray, a graceful rounded delta, had a sudden 

transonic pitch change reflecting its legacy of Messerschmitt-inspired 

tailless aerodynamic design. During one test run to Mach 0.98, it pitched 

58. Jack Fischel and Jack Nugent, “Flight Determination of the Longitudinal Stability in Accelerated 

Maneuvers at Transonic Speeds for the Douglas D-558-II Research Airplane Including the Effects of 

an Outboard Wing Fence,” NACA RM-L53A16 (1953); Jack Fischel, “Effect of Wing Slats and 

Inboard Wing Fences on the Longitudinal Stability Characteristics of the Douglas D-558-II Research 

Airplane in Accelerated Maneuvers at Subsonic and Transonic Speeds,” NACA RM-L53L16 

(1954); Jack Fischel and Cyril D. Brunn, “Longitudinal Stability Characteristics in Accelerated 

Maneuvers at Subsonic and Transonic Speeds of the Douglas D-558-II Research Airplane Equipped 

with a Leading-Edge Wing Chord-Extension,” NACA RM-H54H16 (1954); M.J. Queijo, Byron M. 

Jaquet, and Walter D. Wolmart, “Wind-Tunnel Investigation at Low Speed of the Effects of Chordwise Wing Fences and Horizontal-Tail Position on the Static Longitudinal Stability Characteristics of 

an Airplane Model with a 35° Sweptback Wing,” NACA Report 1203 (1954); Jack Fischel and 

Donald Reisert, “Effect of Several Wing Modifications on the Subsonic and Transonic Longitudinal 

Handling Qualities of the Douglas D-558-II Research Airplane,” NACA RM-H56C30 (1956).
Case 1 | Sweep and Swing: Reshaping the Wing for the Jet and Rocket Age

33

up so violently that test pilot Robert Rahn blacked out, becoming one 

of the first pilots to experience sudden g-induced loss-of-consciousness 

(g-loc). Fortunately, he recovered and returned safely, the battered plane 

now marred by prominent stress-induced wrinkles, giving it a prunelike 

appearance.59 When Grumman entered the transonic swept wing era, it 

did so by converting its conventional straight wing F9F-5 Panther into a 

swept wing design, spawning the F9F-6 Cougar. (The use of an identical 

prefix—“F9F”—indicates just how closely the two aircraft were related.) 

But the Cougar’s swept wing, midplaced horizontal tail, and thick wing 

section (inherited from the firmly subsonic Panther) were ill matched. 

The new Cougar had serious pitch-up and departure characteristics at 

low and high speeds, forcing redesign of its wing before it could be introduced into fleetwide service. Even afterward, however, it retained some 

unpleasant characteristics, particularly a restricted angle-of-attack range 

during carrier landing approaches that gave the pilot only a small maneuver margin before the Cougar would become unstable. Well aware of the 

likely outcome of stalling and pitching up in the last seconds of flight prior 

to “trapping” on a carrier, pilots opted to fly faster, though the safety they 

gained came at the price of less-precise approaches with greater risk of 

“wave-offs” (aborted landings) and “bolters” (touching down beyond the 

cables and having to accelerate back into the air).60

McDonnell’s XF-88, a beefy twin-engine jet fighter prototype from 

the late 1940s, was placed on hold while more powerful engines were 

developed. When finally ordered into development in the early 1950s as 

the F-101 Voodoo, it featured a T-tail, a most unwise choice. Acceptable 

on airliners and transports, the T-tail was anathema for high-performance jet fighters. The Voodoo experienced serious pitch-up problems, 

and the cure was less a “fix” than simply a “patch”: McDonnell installed 

59. Robert O. Rahn, “XF4D Skyray Development: Now It Can Be Told,” 22nd Symposium, Society 

of Experimental Test Pilots, Beverly Hills, CA, Sept. 30, 1978; and Edward H. Heinemann and 

Rosario Rausa, Ed Heinemann: Combat Aircraft Designer (Annapolis: Naval Institute Press, 1980), 

p. 192. Years later, another Skyray pilot at the Naval Air Test Center experienced a similar mishap, 

likewise making a near-miraculous recovery; the plane was so badly stressed that it never flew again.

60. Meyer, Flight Journal, pp. 196–198; he was nearly killed on one low-altitude low-speed pitch-up 

that ended in a near-fatal spin. The Cougar’s approach behavior resulted in a Langley research program flown using a F9F-7 variant, which highlighted the need for more powerful, responsive, and controllable aircraft, such as the later McDonnell F4H-1 Phantom II. See Lindsay J. Lina, Garland J. Morris, 

and Robert A. Champine, “Flight Investigation of Factors Affecting the Choice of Minimum Approach 

Speed for Carrier-Type Landings of a Swept-Wing Jet Fighter Airplane,” NACA RM-L57F13 (1957).

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NASA’s Contributions to Aeronautics

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1 a stick-kicker that would automatically push the stick forward as angle 

of attack increased and the Voodoo approached the pitch-up point. 

Wisely, for their next fighter project (the superlative F4H-1 Phantom II), 

McDonnell designers lowered the horizontal tail location to the base 

of the fin, giving it a characteristically distinctive anhedral (droop).61

Better yet, however, was placing the horizontal tail below the line of 

the wing chord, which, in practical terms, typically meant at the base of 

the rear fuselage, and making it all-moving as well. In 1947, even before the 

first supersonic flights of the XS-1, NACA Langley researchers had evaluated a wind tunnel model of the proposed Bell XS-2 (later X-2) with a lowplaced horizontal tail and a ventral fin, though (unfortunately, given its 

history as related subsequently) Bell completed it with a more conventional 

layout mirroring the XS-1’s midfin location.62 The now-classic jet age low, 

all-moving “stabilator” tail was first incorporated on the North American 

YF-100A Super Sabre, the first of the “Century series” of American fighters.

The low all-moving tail reflected extensive NACA research dating 

to the midst of the Second World War. While the all-moving tail surface 

had been a standard feature of early airplanes such as the German Fokker 

Eindecker (“Monoplane”) and French Morane Bullet fighters of the “Great 

War,” the near constant high workload it made for a pilot caused it to fall 

61. Robert C. Little, “Voodoo! Testing McAir’s Formidable F-101,” Air Power History, vol. 41, no. 

1 (spring 1994), pp. 6–7. In Britain, designer George Edwards likewise added anhedral (though 

more modest than the Phantom’s) to the Supermarine Scimitar, another pitch-up plagued swept wing 

fighter. See Robert Gardner, From Bouncing Bombs to Concorde: the Authorised Biography of Aviation Pioneer Sir George Edwards OM (Stroud, UK: Sutton Publishing, 2006), p. 125. Though not 

per se a swept wing aircraft, the Lockheed F-104 Starfighter, another T-tail design, likewise experienced 

tail-blanketing and consequent pitch-up, necessitating installation of a stick-kicker and imposing of limitations on high angle-of-attack maneuvering. At the time of its design, the benefits of a low-placed tail 

were already recognized, and it is surprising that Clarence “Kelly” Johnson, Lockheed’s legendary 

designer, did not incorporate one. Certainly afterward, he recognized its value, for when, in 1971, 

he proposed a lineal derivative of the F-104, the CL-1200 Lancer (subsequently designated the X-27 

but never built and flown), as a lightweight NATO export fighter, it featured a low, not high, all-moving 

horizontal tail. For X-27 see Jay Miller, The X-Planes: X-1 to X-45 (Hinckley, UK: Midland Publishing, 

2001), pp. 284–289.

62. See Joseph Weil, Paul Comisarow, and Kenneth W. Goodson, “Longitudinal Stability and 

Control Characteristics of an Airplane Model Having a 42.8° Sweptback Circular-Arc Wing with 

Aspect Ratio 4.00, Taper Ratio 0.60, and Sweptback Tail Surfaces,” NACA RM-L7G28 (1947). 

Considerable debate likewise existed on whether the XS-2 should have a shoulder-mounted wing 

with anhedral, a midwing (like the XS-1) without any dihedral or anhedral, or a low wing with or 

without dihedral. Bell opted for a low wing with slight dihedral.
Case 1 | Sweep and Swing: Reshaping the Wing for the Jet and Rocket Age

35

from grace, in favor of a fixed stabilizer and movable elevator surface. But 

by the early 1940s, NACA researchers recognized “its possible advantages 

as a longitudinal control for flight at high Mach numbers.”63 Accordingly, 

researchers at the Langley Memorial Aeronautical Laboratory modified an experimental Curtiss XP-42 fighter on loan from the Army Air 

Forces by removing its conventional horizontal tail surfaces and replacing them with an all-moving tail plane hinged at its aerodynamic center and controlled by a trailing edge servotab. Initial tests during turns 

at 200 mph proved disappointing, with pilots finding the all-moving surface too sensitive and its control forces too light (and thus dangerous, for 

they could easily subject the airplane to excessive maneuvering loads) 

and demanding continuous attention particularly in choppy air. So the 

XP-42 was modified yet again, this time with a geared, not servotab, control mechanism. If not perfect, the results were much better and more 

encouraging, with pilots now having the kind of variation in stick force 

to give them feedback on how effectively they were controlling the airplane.64 Recognizing that the all-moving tail could substantially increase 

longitudinal control authority in the transonic region, NACA researchers continued their study efforts into the postwar years, encouraged by 

initial flight-test results of the Bell XS-1, which began approaching 

higher transonic Mach numbers in the fall of 1946. Though its adjustable 

horizontal stabilizer with a movable elevator constituted an admittedly 

interim step on the path to an all-moving surface, the XS-1’s excursions 

through the speed of sound generated convincing proof that designers 

could dramatically increase transonic longitudinal control authority 

via an all-moving tail.65

63. Harold F. Kleckner, “Preliminary Flight Research on an All-Movable Horizontal Tail as a Longitudinal Control for Flight at High Mach Numbers,” NACA ARR-L5C08 (Mar. 1945), p. 1.

64. Harold F. Kleckner, “Flight Tests of an All-Movable Horizontal Tail with Geared Unbalancing 

Tabs on the Curtiss XP-42 Airplane,” NACA TN-1139 (1946).

65. Hubert M. Drake and John R. Carden, “Elevator-Stabilizer Effectiveness and Trim of the X-1 

Airplane to a Mach Number of 1.06,” NACA RM-L50G20 (1950). Despite the 1950 publication 

date, this report covers the results of XS-1 testing from Oct. 1946 through the first supersonic flight 

to M = 1.06 on Oct. 14, 1947. European designers recognized the value of such a tail layout as 

well. The Miles M.52, a jet-powered supersonic research airplane intemperately canceled by the 

British Labour government, would have incorporated similar surfaces; “This unfortunate decision,” 

Sir Roy Fedden wrote a decade later, “cost us at least ten years in aeronautical progress.” See 

his Britain’s Air Survival: An Appraisement and Strategy for Success (London: Cassell & Co., Ltd., 

1957), p. 20.

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NASA’s Contributions to Aeronautics

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1 Tail location—midfin (as in the XS-1 and D-558), at the base of the 

fin (as with the F-86 and most other jet aircraft), or high (as with the 

T-tail F-101)—was another significant issue. German wartime research 

had favored no tail surfaces or, at the other extreme, high T-tails—for 

example, the DFS 346 supersonic research aircraft under development 

at war’s end or the proposed Focke-Wulf Ta 183 swept wing jet fighter 

(which influenced the design of the MiG-15 and early Lavochkin swept 

wing jet fighters and a proposed British supersonic research aircraft). 

But the pitch-up problems encountered by the Skyrocket and even the 

F-86, as angle of attack increased, argued powerfully against such locations. In 1949, coincident with the Air Force and North American beginning development of the Sabre 45, a 45-degree swept wing successor to 

the F-86, Jack D. Brewer and Jacob H. Lichtenstein, two researchers at 

Langley, undertook a series of studies of tail size, length, and vertical 

location using the Langley stability tunnel and a model having 45-degree 

swept wing and tail surfaces. Their research demonstrated that placing 

a tail well aft of the wing and along the fuselage centerline (as viewed 

from the side) improved longitudinal stability and control.66 Building 

upon their work, Langley researchers William Alford, Jr., and Thomas 

Pasteur, Jr., ran an investigation in the Langley 7-foot by 10-foot highspeed tunnel to determine aspect ratio and location effects on the longitudinal stability of a swept wing model across the transonic regime 

from Mach 0.80 to Mach 0.93. “The results,” they subsequently reported 

in 1953, “indicted that, within the range of variables considered, the 

most favorable pitching-moment characteristics at a Mach number of 

0.90 were obtained by locating the tail below the wing-chord plane.”67

Compared to this, other changes were inconsequential.

Flight tests at Ames in 1952 of a North American YF-86D (an interceptor variant of the F-86) specially modified with a low-placed horizontal tail, confirmed the Langley test results. As researchers noted, “The 

66. Jack D. Brewer and Jacob H. Lichtenstein, “Effect of Horizontal Tail on Low-Speed Static Lateral 

Stability Characteristics of a Model Having 45° Sweptback Wing and Tail Surfaces,” NACA TN2010 (1950); and Jacob H. Lichtenstein, “Experimental Determination of the Effect of Horizontal-Tail 

Size, Tail Length, and Vertical Location on Low-Speed Static Longitudinal Stability and Damping in 

Pitch of a Model Having 45° Sweptback Wing and Tail Surfaces,” NACA Report 1096 (1952).

67. William J. Alford, Jr., and Thomas B. Pasteur, Jr., “The Effects of Changes in Aspect Ratio and 

Tail Height on the Longitudinal Stability Characteristics at High Subsonic Speeds of a Model with a 

Wing Having 32.6° Sweepback,” NACA RM-L53L09 (1953), p. 1.
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37

test airplane, while having essentially the same unstable airplane static 

pitching moments as another version of this airplane [the F-86A] with an 

uncontrollable pitch-up, had only a mild pitch-up which was easily controllable,” and had a nearly 40-percent increase in stabilizer and elevator 

effectiveness at transonic speeds.68 The prototype YF-100 Super Sabre, 

first flown in May 1953, incorporated the fruits of this research. Next came 

the Vought XF8U-1 Crusader and the Republic YF-105 Thunderchief, and 

thereafter a plethora of other types. Aviation had returned full circle to 

the technology with which powered, controlled flight had begun: back to 

pivoted all-moving pitch-control surfaces of a kind the Wrights and other 

pioneers would have immediately recognized and appreciated.

Inertial Coupling: Dangerous Byproduct of High-Speed Design

The progression of aircraft flight speeds from subsonic to transonic and 

on into the supersonic changed the proportional relationship of wing to 

fuselage. As speed rose, the ratio of span to fuselage length decreased. 

At the onset of the subsonic era, the Wright Flyer had a wingspan-tofuselage length ratio of 1.91. The SPAD XIII fighter of World War I was 

1.30. The Second World War’s P-51D decreased to 1.14. Then came the 

supersonic era: the XS-1 was 0.90. In 1953, the F-100A, lowered the ratio 

to 0.80, and the F-104A of 1954 cut this in half, to 0.40. The radical X-3 

had a remarkably slender wingspan-to-fuselage length ratio of just 0.34: 

not without reason was it nicknamed the “Stiletto.” But while the dramatic increase in fuselage length at the expense of span spoke to the 

need to reduce wing-aspect ratio and increase fuselage fineness ratio to 

achieve idealized supersonic shaping, any resulting aerodynamic benefit 

came only at the price of significant performance limitations and risk.

Increasing fuselage length while reducing span dramatically changed 

the mass distribution of these new designs: whereas earlier airplanes 

had most of their mass concentrated along the span of their wings, as 

the wing-fuselage ratio changed from well above 1.0 to well below this 

figure, the distribution of mass shifted to along the fuselage. Since a 

long forward fuselage inherently reduces directional stability, and since 

the small low aspect ratio wings of these airplanes reduced their roll 

stability, a potentially deadly mix of technical circumstances existed to 

68. Norman M. McFadden and Donovan R. Heinle, “Flight Investigation of the Effects of Horizontal-Tail Height, Moment of Inertia, and Control Effectiveness on the Pitch-up Characteristics of a 35° 

Swept-Wing Fighter Airplane at High Subsonic Speeds,” NACA RM-A54F21 (1955).

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1 produce a major crisis: the onset of transonic and supersonic inertial 

coupling, also termed roll-coupling.

William Hewitt Phillips of the NACA’s Langley laboratory had first 

forecast inertial coupling. His pronouncement sprang from a fortuitous 

experience while supervising tests of a large XS-1 “falling body” model in 

the summer of 1947. The model (dropped from a high-flying B-29 over a 

test range near Langley to assess XS-1 elevator control effectiveness as it 

approached Mach 1) incorporated a simple autopilot and was intended 

to rotate slowly as it fell, so as to maintain a “predictable trajectory.”69

But after the drop, things went rapidly awry. The model experienced 

violent pitching and rapid rolling “well below” the speed of sound and 

fell so far from its planned impact point that it literally disappeared 

from history. But optical observations, coupled with telemetric data, 

led Phillips to conclude that “some kind of gyroscopic effect” had taken 

place. Intrigued, he drew upon coursework from Professors Manfred 

Rauscher and Charles Stark Draper of the Massachusetts Institute of 

Technology, using the analogy of the coupling dynamics of a rotating 

rod. He substituted the values obtained from the falling XS-1 model, 

discovering that “the results clearly showed the possibility of a divergent motion. . . . The instability was likely to occur when the values of 

longitudinal stability and directional stability were markedly different and when a large amount of the weight was distributed along the 

fuselage.”70 Hewitt subsequently published a seminal NACA Technical 

Note in 1948, which presciently concluded: “Design trends of very highspeed aircraft, which include short wing spans, fuselages of high density, 

and flight at high altitude, all tend to increase the inertia forces due to 

rolling in comparison with the aerodynamic restoring forces provided 

by the longitudinal and directional stabilities. It is therefore desirable 

to investigate the effects of rolling on the longitudinal and directional 

stabilities of these aircraft. . . . The rolling motion introduces coupling 

between the longitudinal and lateral motion of the aircraft.”71 Out of 

69. W. Hewitt Phillips, Journey in Aeronautical Research: A Career at NASA Langley Research 

Center, No. 12 in the Monographs in Aerospace History Series (Washington: NASA, 1998),

p. 70; for an excellent survey, see Richard E. Day, Coupling Dynamics in Aircraft: A Historical 

Perspective, SP-532 (Washington: NASA, 1997).

70. Phillips, Journey in Aeronautical Research, p. 72.

71. William H. Phillips, “Effect of Steady Rolling on Longitudinal and Directional Stability,” NACA 

TN-1627 (1948), pp. 1–2.
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39

this came the expression “inertial coupling” and its more descriptive 

equivalent, “roll-coupling.” Phillips continued his research on roll-coupling and rolling maneuvers in accelerated flight, noting in 1949 that 

high-speed rolls could generate “exceptionally large” sideslip loads 

on a vertical fin that might risk structural failure. He concluded: “The 

provision of adequate directional stability, especially at small angles 

of sideslip, in order to prevent excessive sideslipping in rolls at high 

speed is therefore important from structural considerations as well as 

from the standpoint of providing desirable flying qualities.”72

In the summer of 1952, as part of an investigation effort studying 

coupled lateral and longitudinal oscillations, researchers at the NACA’s 

Pilotless Aircraft Research Division at Wallops Island, VA, fired a series 

of large rocket-boosted swept wing model airplanes. Spanning over 

3 feet, but with a length of nearly 6 feet, they had the general aerodynamic shape of the D-558-2 as originally conceived: with a slightly 

shorter vertical fin. These models accelerated to supersonic speed 

and then, after rocket burnout and separation, glided onward while 

onboard telemetry instrumentation relayed a continuous stream of key 

performance and behavior parameters as they decelerated through the 

speed of sound before diving into the sea. On August 6, 1952, technicians launched one equipped with a small pulse rocket to deliberately 

destabilize it with a timely burst of rocket thrust. After booster burnout, as the model decelerated below Mach 1, the small nose thruster 

fired, inducing combined yawing, sideslip, and rolling motions. But 

instead of damping out, the model swiftly went out of control, as if a 

replay of the XS-1 falling body test 5 years previously. It rolled, pitched, 

and yawed until it plunged into the Atlantic, its death throes caught 

by onboard instrumentation and radioed to a NACA ground station. 

If dry, the summary words of the resulting test report held ominous 

import for future flight-testing of full-size piloted aircraft: “From the 

flight time history of a rocket-propelled model of a representative 35° 

sweptback wing airplane, it is indicated that coupled longitudinal 

motions were excited and sustained by pure lateral oscillations. The 

resulting longitudinal motions had twice the frequency of the lateral 

oscillations and rapidly developed lift loads of appreciable magnitude. 

The longitudinal moments are attributed to two sources, aerodynamic 

72. William H. Phillips, “Appreciation and Prediction of Flying Qualities,” NACA Report No. 927 

(1949), p. 32.

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NASA’s Contributions to Aeronautics

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1 moments due to sideslip and inertial cross-coupling. The roll characteristics are indicated to be the predominating influence in the inertial 

cross-coupling terms.”73

Two model tests, 5 years apart, had shown that roll coupling was 

clearly more than a theoretical possibility. Shortly thereafter it turned 

into an alarming reality when the Bell X-1A, North American YF-100 

Super Sabre, and Douglas X-3 entered flight-testing. Each of these encountered it with varying degrees of severity. The first to do so was the Bell 

X-1A, a longer, more streamlined, and more powerful derivative of the 

original XS-1.74 The X-1A arrived at Edwards in early 1953, flew a brief 

contractor program, and then entered Air Force evaluation in November. 

On December 12, 1953, test pilot Charles E. “Chuck” Yeager nearly died 

when it went out of control at Mach 2.44 at nearly 80,000 feet. In the low 

dynamic pressure (“low q” in engineering parlance) of the upper atmosphere, a slight engine thrust misalignment likely caused it to begin a 

slow left roll. As Yeager attempted to control it, the X-1A rolled rapidly 

to the right, then violently back to the left, tumbling completely out of 

control and falling over 50,000 feet before the badly battered Yeager managed to regain control. Gliding back to Edwards, he succinctly radioed: 

“You know, if I’d had a seat, you wouldn’t still see me in this thing.”75

Afterward, NACA engineers concluded that “lateral stability difficulties 

were encountered which resulted in uncontrolled rolling motions of the 

airplane at Mach numbers near 2.0. Analysis indicates that this behavior apparently results from a combination of low directional stability 

73. James H. Parks, “Experimental Evidence of Sustained Coupled Longitudinal and Lateral Oscillations from a Rocket-Propelled Model of a 35° Swept Wing Airplane Configuration,” NACA 

RM-L54D15 (1954). For more on Wallops testing, see Joseph A. Shortal, A New Dimension: 

Wallops Island Flight Test Range: The First Fifteen Years, RP-1028 (Washington: NASA, 1978), pp. 

256–257. For the record, the wingspan-to-fuselage ratio of the model was 0.59, significantly lower 

than the XS-1.

74. It is worth noting that the advanced X-1A (and X-1B and X-1D) had a wingspan-to-fuselage 

length ratio of 0.79, compared to the 0.90 XS-1, the drop model of which first encountered inertial 

coupling. Their longer fuselage forebody likewise contributed even further to their tendency toward 

lateral-directional instability.

75. Yeager pilot report and attached transcript, Dec. 23, 1953; J.L. Powell, Jr., “X-1A Airplane 

Contract W33-038-ac-20062, Flight Test Progress Report No. 15, Period From 9 December 

through 20 December 1953,” Bell Aircraft Corporation Report No. 58-980-019 (Feb. 3, 1954), 

both from AFFTC History Office archives. I thank the staff of the AFFTC History Office and the 

NASA DFRC Library and Archives for locating these and other documents.
Case 1 | Sweep and Swing: Reshaping the Wing for the Jet and Rocket Age

41

and damping in roll.”76 The predictions made in Phillips’s 1948 NACA 

Technical Note had come to life, and even worse would soon follow.

By the time of Yeager’s harrowing X-1A flight, the prototype YF-100, 

having first flown in May 1953, was well into its flight-test program. North 

American and the Air Force were moving quickly to fulfill ambitious production plans for this new fighter. Yet all was not well. The prototype 

Super Sabre had sharply swept wings, a long fuselage, and a small vertical fin. While fighter pilots, entranced by its speed, were enthusiastic 

about the new plane, Air Force test pilots were far less sanguine, noting 

its lateral-directional stability was “unsatisfactory throughout the entire 

combat speed range,” with lateral-directional oscillations showing “no 

tendency to damp at all.”77 Even so, in the interest of reducing weight 

and drag, North American actually shrank the size of the vertical fin for 

the production F-100A, lowering its height, reducing its area and aspect 

ratio, and increasing its taper ratio. The changes further cut the directional stability of the F-100A, by some estimates as much as half, over the 

YF-100.78 The first production F-100As entered service in the late summer of 1954. Inertial coupling now struck with a vengeance. In October 

and November, two accidents claimed North American’s chief test pilot, 

George “Wheaties” Welch, and Royal Air Force Air Commodore Geoffrey 

Stephenson, commander of Britain’s Central Fighter Establishment. 

Others followed. The accidents resulted in an immediate grounding while 

the Air Force, North American, and the NACA crafted complementary 

research programs to analyze and fix the troubled program.79

Then, in the midst of the F-100’s travail, inertial coupling struck the 

Douglas X-3. First flown in October 1952, the X-3 had vestigial straight 

wings and tail surfaces joined to a missile-like fuselage. Though it was 

76. Hubert M. Drake and Wendell H. Stillwell, “Behavior of the Bell X-1A Research Airplane During 

Exploratory Flights at Mach Numbers Near 2.0 and at Extreme Altitude,” NACA RM-H55G25 

(1955), p. 10.

77. Alfred D. Phillips and Lt. Col. Frank K. Everest, USAF, “Phase II Flight Test of the North American 

YF-100 Airplane USAF No. 52-5754,” AFFTC TR-53-33 (1953), Appendix I, p. 8. For the difference between fighter pilots and test pilots in regarding the F-100, see Brig. Gen. Frank. K. Everest, 

Jr., with John Guenther, The Fastest Man Alive (New York: Bantam, 1990 ed.), pp. 6, 11–13.

78. James R. Peele, “Memorandum for Research Airplane Projects Leader [Hartley A. Soulé, hereafter RAPL]: Results of Flights 2, 3, and 4 of the F-100A (52-5778) airplane” (Nov. 19, 1954), 

DFRC Archives.

79. Ronald-Bel Stiffler, The History of the Air Force Flight Test Center: 1 January 1954–30 June 1954

(Edwards AFB: AFFTC, July 13, 1955), vol. 1, pp. 66–67, copy in AFFTC History Office archives.

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NASA’s Contributions to Aeronautics

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1 the most highly streamlined airplane of its time, mediocre engines confounded hopes it might achieve Mach 2 speeds, and it never flew faster 

than Mach 1.21, and that only in a dive. The NACA acquired it for research 

in December 1953, following contractor flights and a brief Air Force evaluation. On October 27, 1954, during its 10th NACA flight, test pilot Joseph 

A. Walker initiated an abrupt left aileron roll at Mach 0.92 at 30,000 feet. 

The X-3 pitched up as it rolled, sideslipping as well. After it returned to 

stable flight, Walker initiated another left roll at Mach 1.05. This time, it 

responded even more violently. Sideslip angle exceeded 21 degrees, and 

it reached –6.7 g during a pitch-down, immediately pitching up to over 

7 g. Fortunately, the wild motions subsided, and Walker, like Yeager 

before him, returned safely to Earth.80 With the example of the X-1A, the 

F-100A, and the X-3, researchers had conclusive proof of a newly emergent crisis imperiling the practical exploitation of the high-speed frontier.

The F-100A raised the most concern, for it was the first of an entire 

new class of supersonic fighter aircraft, the “Century series,” with which 

the United States Air Force and at least some of its allies hoped to 

reequip. Welch’s F-100A had sideslipped and promptly disintegrated 

during a diving left roll initiated at Mach 1.5 at 25,000 feet. As Phillips 

had predicted in 1949, the loads had proven too great for the fin to withstand (afterward, North American engineers “admitted they had been 

naive in estimating the effects of reducing the aspect ratio and area of 

the YF-100 prototype tail”).81 Curing the F-100’s inertial coupling problems took months of extensive NACA and Air Force flight-testing, much 

of it very high-risk, coupled with analytical studies by Langley personnel 

using a Reeves Electronic Analogue Computer (REAC), an early form 

of a digital analyzer. During one roll at Mach 0.7 (and only using twothirds of available aileron travel), NACA test pilot A. Scott Crossfield 

experienced “a large yaw divergence accompanied by a violent pitchdown . . . which subjected the airplane to approximately –4.4g vertical 

acceleration.”82 Clearly the F-100A needed significant redesign: the Super 

80. NACA High-Speed Flight Station, “Flight Experience with Two High-Speed Airplanes Having 

Violent Lateral-Longitudinal Coupling in Aileron Rolls,” NACA RM-H55A13 (1955), p. 4.

81. Joseph Weil, “Memo to RAPL: Visit of HSFS personnel to North American Aviation, Inc. on Nov. 

8, 1954” (Nov. 19, 1954), DFRC Archives.

82. Peele memo to RAPL, Nov. 19, 1954; for Langley REAC studies, see Charles J. Donlan [NACA 

LRC], “Memo for Associate Director [Floyd Thompson]: Industry-Service-NACA Conference on 

F-100, Dec. 16, 1954” (Dec. 28. 1954), DFRC Archives.
Case 1 | Sweep and Swing: Reshaping the Wing for the Jet and Rocket Age

43

Sabre’s accidents and behavior (and that of the X-3 as well) highlighted 

that streamlined supersonic aircraft needed greatly increased tail area, 

coupled with artificial stability and motion damping, to keep sideslip 

from developing to dangerous values. North American subsequently 

dramatically increased the size of the F-100’s vertical fin, increased its 

wingspan by 2 feet (to shift the plane’s center of gravity forward), and 

incorporated a yaw damper to control sideslip. Though the F-100 subsequently became a reliable fighter-bomber (it flew in American service 

for almost a quarter century and longer in foreign air arms), it remained 

one that demanded the constant attention and respect of its pilots.83

 Inertial coupling was not, of course, a byproduct of conceptualizing 

the swept and delta wings, nor was it limited (as the experience of the 

XS-1 falling model, X-1A, and X-3 indicated) just to aircraft possessing 

swept or delta planforms. Rather, it was a byproduct of the revolution 

in high-speed flight, reflecting the overall change in the parametric relationship between span and length that characterized aircraft design in 

the jet age. Low aspect ratio straight wing aircraft like the X-3 and the 

later Lockheed F-104 were severely constrained by the threat of inertial coupling, even more than many swept wing aircraft were.84 But for 

swept wing and delta designers, inertial coupling became a particular 

challenge they had to resolve, along with pitch-up. As the low-placed 

horizontal tail reflected the problem of pitch-up, the increasing size of 

vertical fins (and the addition of ventral fins and strakes as well) incorporated on new aircraft such as the Navy’s F8U-1 and the Air Force’s 

83. Joseph Weil and Walter C. Williams, “Memo for RAPL: Meeting of NACA and Air Force 

personnel at North American Aviation, Inc on Monday, Nov. 22, 1954, to discuss means of expediting solution of stability and control problems on the F-100A airplane” (Nov. 26, 1954), DFRC 

Archives; Thomas W. Finch, “Memo for RAPL: Progress report for the F-100A (52-5778) airplane 

for the period Nov. 1 to Nov. 30, 1954” (Dec. 20, 1954), DFRC Archives; Hubert M. Drake, 

Thomas W. Finch, and James R. Peele, “Flight Measurements of Directional Stability to a Mach 

Number of 1.48 for an Airplane Tested with Three Different Vertical Tail Configurations,” NACA 

RM-H55G26 (1955); Marion H. Yancey, Jr., and Maj. Stuart R. Childs, USAF, “Phase IV Stability 

Tests of the F-100A Aircraft, USAF S/N 52-5767,” AFFTC TR-55-9 (1955); 1st Lt. David C. Leisy, 

USAF, and Capt. Hugh P. Hunerwadel, USAF, “ARDC F-100D Category II Performance Stability 

and Control Tests,” AFFTC TR-58-27 (1958).

84. See, for example, Robert G. Hoey and Capt. Iven C. Kincheloe, USAF, “ARDC F-104A Stability and Control,” AFFTC TR-58-14 (1958); and Capt. Slayton L. Johns, USAF, and Capt. James 

W. Wood, USAF, “ARDC F-104A Stability and Control with External Stores,” AFFTC TR-58-14, 

Addendum 1 (July 1959).

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1 F-105B (and the twin-fins that followed in the 1970s on aircraft such as 

the F-14A, F-15A, and F/A-18A) spoke to the serious challenge the inertial coupling phenomenon posed to aircraft design. Not visible were such 

“under the skin” systems as yaw dampers and the strict limitations on 

abrupt transonic and supersonic rolling taught to pilots transitioning 

into these and many other first-generation supersonic designs.85

 The story of the first encounters with inertial coupling is a salutary, cautionary tale. A key model test had resulted in analysis leading 

to the issuance of a seminal report but one recognized as such only in 

retrospect. A half decade after the report’s release, pilots died because 

the significance of the report for future aircraft design and behavior 

had been missed. Even within the NACA, recognition of seriousness of 

reduced transonic and supersonic lateral-directional stability had been 

slow. When, in August 1953, NACA engineers submitted thoughts for 

a tentative research plan for an F-100A that the Agency would receive, 

attention focused on longitudinal pitch-up, assessing its handling qualities (particularly its suitability as a gun platform, something seemingly 

more appropriately done by the Air Force Flight Test Center or the Air 

Proving Ground at Eglin), and the correlation of flight and wind tunnel measurements.86 Even after the experience of the X-1A, F-100A, and 

X-3, even after all the fixes and training, it is disturbing how inertial 

coupling stilled claimed the unwary.87 Over time, the combination of 

refined design, advances in stability augmentation (and eventually the 

advent of computer-controlled fly-by-wire flight) would largely render 

85. For example, Thomas R. Sisk and William H. Andrews, “Flight Experience with a Delta-Wing 

Airplane Having Violent Lateral-Longitudinal Coupling in Aileron Rolls,” NACA RM-H55H03 (1955).

86. William H. Phillips [NACA LRC], “Memo for Associate Director: Flight program for F-100A 

airplane” (Aug. 10, 1953), DFRC Archives. Even odder, it was Phillips who had identified inertial 

coupling in TN-1627 in 1948!

87. The best known was Capt. Milburn “Mel” Apt, who died in late 1956. His Bell X-2 went out of 

control as he turned back to Edwards after having attained Mach 3.2, possibly because of lagging 

instrumentation readings leading him to conclude he was flying at a slower speed. Undoubtedly the 

nearly decade-old design of the X-2 contributed to its violent coupling tendencies. It is sobering that 

in 1947 NACA had evaluated some design options (tail location, vertical fin design) that, had Bell 

incorporated them on the X-2, might have turned Apt’s accident into an incident. See Ronald Bel 

Stiffler, The Bell X-2 Rocket Research Aircraft: The Flight Test Program (Edwards AFB: Air Force Flight 

Test Center, 1957); and Richard E. Day and Donald Reisert, “Flight Behavior of the X-2 Research 

Airplane to a Mach Number of 3.20 and a Geometric Altitude of 126,200 Feet,” NACA TM-X137 (1959).
Case 1 | Sweep and Swing: Reshaping the Wing for the Jet and Rocket Age

45

inertial coupling a curiosity. But for pilots of a certain age—those who 

remember aircraft such as the X-3, F-100, F-101, F-102, and F-104—

the expression “inertial coupling,” like “pitch-up,” will always serve to 

remind that what is an analytical curiosity in the engineer’s laboratory 

is a harsh reality in the pilot’s cockpit.

Implementing the Delta Planform

While swept wing adaptation in Europe, Russia, and America followed 

a generally similar pattern, the delta wing underwent markedly different international development. Generally, European designers initially 

emulated the Lippisch approach, resulting in designs with relatively 

thick wing sections (exemplified by the Avro Vulcan bomber and the 

“tailed” Gloster Javelin interceptor) that inhibited their ability to operate 

beyond the transonic. Only after the practical demonstration of Convair’s 

emerging family of thin-wing delta designs—the XF-92A research aircraft, the F-102 interceptor, the XF2Y-1 experimental naval fighter, the 

B-58 supersonic bomber, and the F-106 interceptor—did they conceptualize more “supersonic friendly” designs, typified by the Swedish Saab 

J35 Draken (“Dragon”), the British Fairey F.D.2 research airplane, the 

French Dassault Mirage I (progenitor of the Mirage fighter and bomber 

family). By the late 1950s, British and French aerodynamicists had so 

completely “closed” any “delta gap” that might have existed between 

Europe and America that they were already conceptualizing development 

of a Mach 2 supersonic transatlantic transport using a shapely “ogee” 

reflexive delta planform, a study effort that would, a decade later, spawn 

the Anglo-French Concorde.88 Not so taken with the pure delta, Soviet 

designers joined American-like thin delta wings to the low-placed horizontal tail, generating advanced MiG and Sukhoi fighters and interceptors. These “tailed deltas” (particularly the MiG-21) possessed far better 

transonic and supersonic turning performance than could be attained 

by a conventional delta with its high induced drag onset at the increasing angles of attack characteristic of hard-maneuvering. (An American 

equivalent was the Douglas Company’s superlative A4D-1 Skyhawk, 

88. Keneth Owen, Concorde: Story of a Supersonic Pioneer (London: Science Museum, 2001), 

pp. 21–60; Andrew Nahum, “The Royal Aircraft Establishment from 1945 to Concorde,” in Robert 

Bud and Philip Gummett, eds., Cold War, Hot Science: Applied Research in Britain’s Defence 

Laboratories, 1945–1990 (London: Science Museum, 1999), pp. 29–58; and Andersson, Saab 

Aircraft, pp. 124–129.

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1 a light attack bomber with maneuvering performance better than 

most fighters.)

Although it is commonly accepted that American delta aircraft 

owe their inspiration to the work of Lippisch—Convair’s delta aircraft 

repeatedly being cited as the products of his influence—in fact, they do 

not.89 Unlike, say, the swept wing F-86 and B-47, which directly reflected 

German aerodynamic thought and example, America’s delta wing aircraft reflected indigenous, not foreign, research and inspiration. By the 

time that Lippisch first met with Allied technical intelligence experts, 

American aerodynamicists were already advancing along a very different path than the one he had followed. Jones had already enunciated his 

thin, sharply swept delta theory and undertaken his first tunnel tests of 

it. In June 1946—a full year after the German collapse—Convair engineers developing the experimental delta XP-92 interceptor had their 

chance to meet with Lippisch at Wright Field. By then, however, they had 

already independently decided upon a thin delta planform. “We had heard 

about Dr. Lippisch’s work and this gave us some moral support,” Convair 

designer Adolph Burstein recalled, adding: “but not much else. . . . We 

did not go along with many of his ideas, such as a very thick airfoil.”90

Burstein and his colleagues arrived at their delta shape by beginning with 

a 45-degree swept wing, gradually increasing its sweepback angle, and 

then “filling in” the ever-closing trailing edges, until they arrived at the 

classic 60-degree triangular delta planform the company incorporated 

on all its subsequent delta aircraft. With a 6.5 thickness-chord ratio—

less than half that of Lippisch’s DM-1—it was an altogether differentlooking airplane.91 Nor was Convair alone in going its own way; Douglas 

naval aircraft designer Edward Heinemann acknowledged that “At the 

close of World War II the work with delta planforms accomplished by 

89. Even the official Air Force history of the service’s postwar fighter development repeats the 

canard, though it does acknowledge that “low-aspect-ratio wing forms were also studied by the 

U.S. National Advisory Committee for Aeronautics.” See Marcelle Size Knaack, Post-World War II 

Fighters 1945–1973, vol. 1 of Encyclopedia of U.S. Air Force Aircraft and Missile Systems (Washington: Office of Air Force History, 1978), p. 159, no. 1.

90. Letter, Adolph Burstein to Richard P. Hallion, Jan. 25, 1972. Despite his “Germanic” name, 

Burstein, one of the XF-92A’s designers, was not a German scientist or engineer who came to 

America after 1945. Rather, he was a Russian emigree from St. Petersburg who had come to the 

United States in 1925.

91. See Hallion, “Lippisch Gluhareff, and Jones,” and R.P. Hallion, “Convair’s Delta Alpha,” Air 

Enthusiast Quarterly, No. 2 (1976).
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Lippisch DM-1 glider in the Langley Full Scale Tunnel, 1946. The thick-wing section is readily apparent, 

as is the oversize vertical fin, both of which rendered the concept unsuitable for transonic flight. NASA.

Dr. Lippisch in Germany became generally known and appreciated,” 

but that “Extensive wind tunnel tests showed there was no special merit 

to an equilateral triangle planform—especially those designed with 

thicker airfoils.”92

The chronology of American delta development, and the technical 

choices and paths followed by American engineers, supports both statements. At war’s end, advancing ground forces at Prien, Austria, had discovered a thick-wing wooden delta glider, the DM-1, which Lippisch 

had intended as a low-speed testbed for a proposed supersonic fighter, 

the P 13. At Army Air Forces’ request, it was shipped back to America 

in January 1946 for comprehensive testing in the Full-Scale Tunnel at 

the NACA’s Langley Aeronautical Laboratory. Had the tests gone well, 

92. Edward H. Heinemann, “Design of High-Speed Aircraft,” a paper presented at the Fifth International Aeronautical Conference, Royal Aeronautical Society-Institute of the Aeronautical Sciences, Los 

Angeles, CA, June 20–24, 1955, p. 3. Copy from the Boeing-McDonnell Douglas Archives.

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1 the possibility existed as that, as the Germans had intended, it might 

be flown as a glider. But the tunnel tests quickly disabused delta enthusiasts of these hopes. As the AAF’s Langley liaison officer subsequently 

reported, the “Initial test results were very disappointing; the lift coefficient was low, the drag was high, the directional stability was unsatisfactory, and the craft was considered unsafe for flight tests.”93

Afterward, Langley engineers undertook a comprehensive study of 

the DM-1 configuration, not in the spirit of emulation but rather attempting to find a way to fix it. After giving its wings sharp leading edges, sealing all slots and gaps around control surfaces, and removing the thick 

vertical fin and replacing it with a thin one (relocating the pilot under 

a streamlined bubble canopy), they had markedly improved its performance, doubling its lift coefficient, from 0.6 to over 1.2. But it remained 

an unsatisfactory design, proof enough that the Lippisch concept of deltas was hardly one that could serve—or did serve—as a veritable 

template (as has been so often alleged) for the supersonic American, 

Swedish, and French delta fighters and bombers that flew over the next 

decade.94 Subsequently, NACA engineers looked to far thinner and more 

streamlined configurations that, if not yet as extreme as Robert T. Jones’s 

original daggerlike concept, were even more amenable to the rigors of 

transonic and supersonic flight than the generously rounded contours 

of Lippisch’s thick wings and awkward pilot-enclosing vertical 

fins. By the beginning of 1947, they were already examining the 

technical requirements of slender, low aspect ratio delta configurations 

93. Ltr., Maj. Howard C. Goodell, USAF, to Paul E. Garber, “DM-1 Glider Disposal,” Nov. 28, 

1949, in Gluhareff Dart accession file, National Air and Space Museum.

94. For Langley’s progressive evaluation and modification of the DM-1, see two reports by Herbert 

A. Wilson, Jr., and J. Calvin Lovell, “Full Scale Investigation of the Maximum Lift and Flow Characteristics of an Airplane Having Approximately Triangular Plan Form,” NACA RM-L6K20 (1947); and 

“Langley Full-Scale Tunnel Investigation of Maximum Lift and Stability Characteristics of an Airplane 

Having Approximately Triangular Plan Form (DM-1 Glider), NACA RM-L7F16 (1947). Changes 

are detailed in RM L7F16, Fig. 4. The closest expression of Germanic delta philosophy in America 

was not a Convair delta, but a Douglas one: the Navy-Marine F4D-1 Skyray fighter. Its design was 

greatly influenced by German tailless and swept wing reports Douglas engineers L. Eugene Root 

and Apollo M.O. Smith had discovered while assigned to an Allied technical intelligence team 

examining the Messerschmitt advanced projects office at Oberammergau and interviewing its senior 

personnel, particularly chief designer Woldemar Voigt; I wish to acknowledge with gratitude notes 

on their experiences received in 1972 from both the late L. Eugene Root and A.M.O. Smith. See 

also Cebeci, ed., Legacy of a Gentle Genius, pp. 30–36.
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The Convair XF-92A, the world’s first delta jet airplane, at the NACA High-Speed Flight Research 

Station, now the Dryden Flight Research Center, in 1953. NASA.

to meet emerging military specifications for a Mach 1.5, 60,000-foot 

bomber interceptor.95

First Flight Experiences

Out of this mutually reinforcing climate of thought emerged the world’s 

first delta jet airplane, the Convair XF-92A, first flown in September 

1948. This technology explorer (for despite its “fighter” designation, it 

was always intended for research purposes) demonstrated the potential of the delta wing and encouraged Convair and Air Force authorities 

to pursue a delta planform for a future interceptor design. Originally, 

that design had been the “XP-92,” an impractical barrel-shaped rocketboosted ramjet with the pilot sitting in a conical nose within the ramjet’s 

95. R.M. Cross, “Characteristics of a Triangular-Winged Aircraft: 2: Stability and Control,” 

in NACA, Conference on Aerodynamic Problems of Transonic Airplane Design (1947), pp. 

163–186, and Figs. 6 and 12. See also Edward F. Whittle, Jr., and J. Calvin Lovell, “Full-Scale 

Investigation of an Equilateral Triangular Wing Having 10-percent-Thick Biconvex Airfoil Sections,” 

NACA RM-L8G05 (1948), Fig. 2.

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1 circular inlet, similar to René Leduc’s straight wing air-launched French 

ramjet designs of the same period. Following its cancellation, work on 

the XF-92A continued, supporting the Air Force’s “1954 Interceptor” initiative, which Convair hoped to win with, essentially, a bigger and more 

powerful version of the XF-92A. Aside from greater power, the interceptor would have to have a nose radar and thus “cheek” inlets rather than 

the simple Pitot nose inlet of the smaller testbed. The “1954 Interceptor” 

eventually became two: the “interim” Mach 1+ F-102 Delta Dagger and 

the “ultimate” Mach 2+ F-106 Delta Dart.

The XF-92A contributed markedly to delta understanding but was 

far from a trouble-free design. Deltas evinced a variety of quirks and performance deficiencies, some of which they shared with their swept wing 

brethren. Deltas manifested the same tendency to persistent combined 

lateral-directional Dutch roll motions, as well as pitch-up, from Mach number effects as they entered further into the transonic regime. The extreme 

sweep of their wings accentuated spanwise flow tendencies, making wing 

fences almost mandatory in all cases. Their high angle-of-attack (“hi AoA”) 

landing approaches highlighted potentially serious control deficiencies, for, 

unlike a conventional fighter, the delta lacked separate elevators and ailerons. It relied instead on elevons—combined elevator-ailerons—for pitch 

and roll control. Thus, with the stick pulled back on final approach, the 

nose would rise, and if the plane encountered a sudden gust that induced 

a rolling motion, the pilot might lack sufficient remaining reserve “travel” 

from the deflected elevon to correct for the rolling motion. Further complicating landing approaches and turn performance was the delta’s inherently high-induced drag as it turned or was at higher angles of attack. 

Deltas needed lots of power. The high-induced drag of the delta led to a 

rapid bleeding off of airspeed during turns and thus inhibited its holding 

altitude during turning maneuvers. Tests with the little XF-92A in 1953 

by NACA research pilot Scott Crossfield indicated that as much as 3,000 

feet of altitude could be lost trying to maintain constant speed in a turning maneuver—and this was after it had been modified to incorporate an 

afterburner for greater power. “Every time I took off in that plane I held my 

brief until I reached sufficient altitude to use the ejection seat,” Crossfield 

recollected later. “The pilot never really flew that airplane, he corralled it.”96

96. Crossfield with Blair, Always Another Dawn, p. 167; Thomas R. Sisk and Duane O. Muhleman, 

“Longitudinal Stability Characteristics in Maneuvering Flight of the Convair XF-92A Delta-Wing 

Airplane Including the Effects of Wing Fences,” NACA RM-H4J27 (1955).
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All together, the NACA completed 25 flights in the XF-92A before a landing gear collapse brought its research career to an end.

Tests of the XF-92A foreshadowed similar challenges with the next 

Convair delta, the prototype YF-102 interceptor. The YF-102 is infamous 

for having suffered from such high transonic drag rise that it could not 

accelerate through the speed of sound, a discovery that led, as Air Force 

test pilot Lt. Col. Frank K. “Pete” Everest recalled, to “surprise and concern. . . . The National Advisory Committee for Aeronautics had claimed 

all along that the airplane would not go supersonic, and now their predictions came true.”97 (How the YF-102 was transformed from embarrassing failure to operational success, thanks to Richard Whitcomb’s “area 

rule” theory and its practical application to the F-102 design, is covered 

elsewhere in this volume in a case study on Whitcomb’s contributions to 

aeronautics, by historian Jeremy Kinney.) But more than reshaping of 

its fuselage was required before the F-102 became a success. Instead, its 

wing underwent fundamental aerodynamic redesign reflecting the second 

stage in American delta development and its third stage overall.

Reshaping the Delta: Deriving Conical Camber

Having preceded the explication of the swept wing in Jones’s original 

research, the roots of the delta’s redesign now lay, somewhat ironically, 

in his expanding upon the slender swept wing research he had first 

begun at Langley. After the war, Jones had left Virginia’s Tidewater region 

for the equally pleasant Bay area environment of Sunnyvale, CA, and 

there had continued his swept wing studies. By 1947, he had evolved a 

sharply swept symmetrical airfoil planform he considered suitable for 

a supersonic jet transport. Such a planform, with the leading edges of 

the wings within the shock cone formed around the vehicle and thus in 

a region of subsonic flow, could perhaps have a lift-to-drag (L/D) ratio 

as high as 10, though at the price of much higher landing speeds.98

Tests of a small model in the Ames 1-foot by 3-foot supersonic tunnel and a larger one in the Ames 40-foot by 80-foot tunnel encouraged 

Jones and inspired fellow Ames researchers Charles F. Hall and John 

C. Heitmeyer to build upon his work. Hall and Heitmeyer considered 

the behavior of the combined wing-body, with the wing twisted and 

97. Everest with Guenther, Fastest Man Alive, p. 109.

98. R.T. Jones, “Characteristics of a Configuration with a Large Angle of Sweepback,” in NACA, 

Conference on Aerodynamic Problems of Transonic Airplane Design (1947), pp. 165–168, Figs. 1–6.

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1 given camber (curvature) to evenly distribute the flight loads, deriving a sharply swept and tapered wing configuration that demonstrated 

an L/D of 8.9 during tunnel tests to Mach 1.53.99 In the refinement of 

its planform, it called to mind the shape (though, of course, not the 

airfoil section) of Whitcomb’s later supercritical transonic transport 

wing conceptualizations.100

Hall and Heitmeyer next broadened their research to examine slender deltas likewise featuring aerodynamic twist and camber. In 2 years, 

1951–1952, they coauthored a dozen reports, culminating in the issuance of a seminal study by Hall in the spring of 1953 that summarized 

the lift, drag, pitching moment, and load distribution data on a variety 

of thin delta wings of varying aspect ratios operating from Mach 0.25 

(touchdown velocity) to Mach 1.9. Out of this came the concept of leading edge “conical camber”: twisting and rounding the leading edge of 

a delta wing to minimize performance-robbing drag generated by the 

wing’s own lifting force. The modified delta had minimal camber at the 

wing root and maximum camber at the tip, the lineal development of 

the camber along the leading edge effectively representing the surface of 

a steadily expanding cone nestled under the leading edge of the wing.101

Hall’s conical camber, like Whitcomb’s area rule, came just in time 

to save the F-102 program. Both were necessary to make it a success: 

Whitcomb’s area rule to get it through the sound barrier, and Hall’s 

to give it acceptable transonic and supersonic flying qualities. If overshadowed by Whitcomb’s achievement—which resulted in the young 

Langley aerodynamicist receiving the Robert J. Collier Trophy, American 

aviation’s most prestigious award, in 1954—Hall’s conical camber concept was nevertheless a critical one. Comparative flight-testing of the 

YF-102 at the NACA High-Speed Flight Station at Edwards from late 

99. Charles F. Hall and John C. Heitmeyer, “Aerodynamic Study of a Wing-Fuselage Combination 

Employing a Wing Swept Back 63°—Characteristics at Supersonic Speeds of a Model with the 

Wing Twisted and Cambered for Uniform Load,” NACA RM-A9J24 (1950).

100. Though no transport or military aircraft ever flew with such a slender swept wing, just such a 

configuration was subsequently employed on the largest swept wing tailless vehicle ever flown, the 

Northrop Snark intercontinental cruise missile. Though the Snark did not enter operational service for 

a variety of other reasons, it did demonstrate that, aerodynamically, such a wing configuration was 

eminently suitable for long-range transonic cruising flight.

101. Charles F. Hall, “Lift, Drag, and Pitching Moment of Low-Aspect Ratio Wings at Subsonic and 

Supersonic Speeds,” NACA RM-A53A30 (1953). For the views of an Ames onlooker, see Hartman, Adventures in Research, pp. 202–207.
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1954 to mid-1955 with and without conical camber indicated that conical camber gave it lower drag and increased its maximum lift-to-drag 

ratio by approximately 20 percent over a test Mach number range of 0.6 

to 1.17, at altitudes of 25,000, 40,000, and 50,000 feet. Transonic stability of the cambered versus symmetrical YF-102 more than doubled, 

and “no severe pitch-up tendencies were exhibited, except when accelerating or decelerating through the trim-change region.”102

With the advent of conical camber, the age of the practical transonicsupersonic delta wing had arrived. By mid-decade, the F-102’s aerodynamic deficiencies had been cured, and it was well on its way to 

service use.103 Convair designers were refining the delta planform to 

generate the F-102’s successor, the superlative F-106, and a four-engine 

Mach 2+ bomber, the delta wing B-58 Hustler. Overseas, Britain’s Fairey 

Company had under test a delta of its own, the F.D.2, which would 

shortly establish an international speed record, while, in France, Dassault 

engineers were conceptualizing a design that would spawn the Mirage 

family and be responsible, in 1967, for one of the most remarkable aerial 

victories of all time. Jones’s supersonic delta vision from over a decade 

previously had become reality, thanks in part to Whitcomb’s interference studies (which Jones himself would expand at Ames) and Hall’s 

conceptualization of conical camber. 

Extending the Delta into the Hypersonic and Orbital Frontier

The next stage in delta development took it from the realm of the transonic and supersonic into the hypersonic, again thanks to a healthy 

rivalry and differing technical perspective between those two great 

research centers, Ames and Langley. The area was hypersonics: flight 

at speeds higher than Mach 5, an area of intense inquiry in the mid1950s following upon the success of the supersonic Round One research 

102. Quoted in William E. Andrews, Thomas R. Sisk, and Robert W. Darville, “Longitudinal Stability Characteristics of the Convair YF-102 Airplane Determined from Flight Tests,” NACA RM-H56I17 

(1956), p. 1; see also Edwin J. Saltzman, Donald R. Bellman, and Norman T. Musialowski, 

“Flight-Determined Transonic Lift and Drag Characteristics of the YF-102 Airplane With Two Wing 

Configurations,” NACA RM-H56E08 (1956).

103. Although it still experienced some troubled sailing: like most of the Century series fighters, the 

F-102 had other, more tortuous acquisition and program management problems unrelated to its 

aerodynamics that contributed to its delayed service entry. See Thomas A. Marschak, The Role of 

Project Histories in the Study of R&D, Rand report P-2850 (Santa Monica: The Rand Corporation, 

1965), pp. 66–81; and Knaack, Fighters, pp. 163–167.

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1 aircraft. Already a Round Two hypersonic test vehicle, the soon-to-emerge 

North American X-15 was underway. But what of high-hypersonics, the 

hypersonics of flight at Mach 10 to orbital velocity?

Hypersonics constituted a natural application for the low aspect ratio 

delta planform. Before the Second World War, Austrian engineer Eugen 

Sänger and his mathematician wife, Irene Sänger-Bredt, had conceptualized the Silbervogel (“Silver Bird”), a flat-bottom, half ogive body 

shape as a potential Earth-girdling hypersonic boost-glider. It had, for 

its time, a remarkable advanced aerodynamic profile, introducing the 

flat bottom and ogival configuration that did, in fact, come to characterize hypersonic aerothermodynamic design. But in one respect it did 

not: Sänger-Bredt’s “antipodal aircraft” had a conventional wing (though 

of low aspect planform and with supersonic wedge airfoils). Although 

it proved very influential on the course of postwar hypersonics, by the 

mid-1950s, as high-speed aerodynamic thinking advanced beyond the 

supersonic and into the hypersonic realm, attention increasingly turned 

toward the sharply swept delta planform.

In 1951, Ames researchers H. Julian Allen and Alfred Eggers, Jr., 

had postulated the blunt-body reentry theory that led to the advent of 

the practical reentry shape used subsequently both for missile warheads 

and the first human presence in space.104 (Their work, and the emergence of the hypersonics field generally, are discussed in greater detail in 

T.A. Heppenheimer’s accompanying essay on transatmospherics.) While 

blunt-body theory enabled safely transiting the atmosphere, it did not 

furnish the flexibility of a large landing “footprint”; indeed, in practice, 

blunt-body reentry was limited to “throwaway” reentry shapes and programs such as Mercury, Gemini, and Apollo that necessitated a large 

and cumbersome investment in oceanic recovery of returning spacecraft. Some sort of lifting vehicle that could fly at hypersonic velocities 

would have far greater flexibility.

Related to the problem of hypersonic flight was the challenge of 

increasing lift-to-drag ratios at high supersonic speeds. Eggers, working with Ames researcher Clarence A. Syvertson, now turned away 

from blunt-body theory to examine thin, slender deltas. The two recognized that “the components of the aircraft should be individually 

104. H. Julian Allen and A.J. Eggers, Jr., “A Study of the Motion and Aerodynamic Heating of Ballistic Missiles Entering the Earth’s Atmosphere at High Supersonic Speeds,” NACA TR-1381 (1953); 

Hartman, Adventures in Research, pp. 215–218.
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and collectively arranged to impart the maximum downward and the 

minimum forward momentum to the surrounding air.”105 Out of this 

emerged the hypersonic “flattop” delta, a high-wing concept having the 

wing perched above the body (in this case, surmounting the classic 

half-ogive hypersonic shape), incongruously much like a general 

aviation light airplane such as a Cessna 152. At mid-span, its tips 

would angle sharply downward, capturing the momentum of flow 

imparted laterally outward from the body and deflecting it into 

downward momentum, thus greatly increasing lift. The tips as well 

furnished directional stability. This flattop concept, which Eggers 

and Syvertson enunciated in 1956, spawned an Ames concept for a 

hypersonic “beyond X-15” Round Three research vehicle that could 

be air-launched from a modified Convair B-36 bomber for initial trials to Mach 6 and, once proven, could then be launched 

vertically as the second stage of a two-stage system capable of reaching Mach 10 and transiting the United States. The Ames vehicle, with 

an overall length of 70 feet and a span of just 25 feet, represented a 

bold concept that seemed likely to spawn the anticipated Round Three 

hypersonic boost-glider.106

But the flattop delta was swiftly undone by a rival Round Three 

Langley concept that echoed more the earlier work of Sänger-Bredt. 

A 1957 study by Peter Korycinski and John Becker demonstrated that 

a flat-bottom (that is, low-wing) delta boost-glider would have better cooling characteristics (a vital concern at hypersonic velocities) 

and thus require less weight for thermal protection systems. Any 

lift-to-drag advantages of the Ames flattop high-wing concept were 

thus nullified. Round Three went forward, evolving into the abortive 

Air Force–NASA X-20 Dyna-Soar program, which employed the Langley 

105. A.J. Eggers, Jr., and Clarence A. Syvertson, “Aircraft Configurations Developing High Lift-Drag 

Ratios at High Supersonic Speeds,” NACA RM-A55L05 (1956), p. 1.

106. Ames staff, “Preliminary Investigation of a New Research Airplane for Exploring the Problems 

of Efficient Hypersonic Flight,” (Jan. 18, 1957), copy in the archives of the Historical Office, NASA 

Johnson Space Center, Houston, TX. Drawings and more data on this concept can be found in 

Richard P. Hallion, ed., From Max Valier to Project PRIME (1924–1967), vol. 1 of The Hypersonic 

Revolution: Case Studies in the History of Hypersonic Technology (Washington: USAF, 1998), pp.IIvi–II-x. Round One, in NACA parlance, was the original X-1 and D-558 programs. Round Two was 

the X-15. Round Three was what eventually emerged as the X-20 Dyna-Soar development effort.

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1 flat-bottom approach, not the high-wing flattop delta of Ames.107 Ames 

and Langley contested a decade later, this time in rival lifting bodies, 

with the Ames half-cone flattop M-2 (the product of Allen, Eggers, 

Syvertson, George Edwards, and George Kenyon) competing against 

Langley’s HL-10 fattened flat-bottom delta (by Eugene S. Love). Again, 

it was the flat-bottom delta that proved superior, confirmed by tests in 

the mid-1970s with an even more refined flat-bottom Air Force-derived 

slender delta body shape, the Martin X-24B.108

When orbital cross range proved even of greater significance, Shuttle 

proponents from the National Aeronautics and Space Administration 

(NASA) and the Air Force in the 1970s looked away from flattop and lifting body approaches and more toward blended bodies, modified delta 

planforms, and exotic delta “wave riders.” Though NASA’s Spacecraft 

Design Division briefly considered a conventionally tailed, straight 

and swept wing Shuttle concepts, reflecting an influential study by 

Johnson’s Maxime Faget, it moved rapidly toward deltas after analysis 

indicated such designs had a tendency of hypersonic spins, suspect aerothermal survivability, and too small a cross range during return from 

orbit. Between mid-1971 and the late summer of 1972, the Spacecraft 

Design Division evaluated no less than 37 separate delta configurations, 

ranging from simple triangular shapes echoing the early days of Jones 

to much more complex ogee shape reflecting the refinement of the delta 

as exemplified by the Anglo-French Concorde. Aside from continuous 

review by the Manned Spacecraft Center (MSC; subsequently the NASA 

Lyndon B. Johnson Space Center), these evaluations benefitted greatly 

from aerodynamic analysis by NASA’s Ames and Langley hypersonic 

107. See John V. Becker, “The Development of Winged Reentry Vehicles, 1952–1963,” in Hallion, 

ed., Hypersonic Revolution, vol. 1, pp. 379–448. It is worth noting that one significant aircraft project did use the Eggers-Syvertson wing but in a modified form: the massive North American XB-70A 

Valkyrie Mach 3+ experimental bomber. The XB-70 had its six engines, landing gear, and weapons 

bays located under the wing in a large wedge-shaped centerbody. The long, cobralike nose ran 

forward from the wing and featured canard control surfaces. Its sharply swept delta wing had 

outer wing panels that could entrap the lateral momentum off the ventral centerbody and transfer it 

downward to furnish compression lift.

108. For further detail, see R. Dale Reed with Darlene Lister, Wingless Flight: The Lifting Body Story, 

SP-4220 (Washington: NASA 1997); Milton O. Thompson and Curtis Peebles, Flying Without 

Wings: NASA Lifting Bodies and The Birth of the Space Shuttle (Washington: Smithsonian Institution Press, 1999); and Johnny G. Armstrong, “Flight Planning and Conduct of the X-24B Research 

Aircraft Flight Test Program,” Air Force Flight Test Center TR-76-11 (1977).
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57

communities, the practical low lift-to-drag-ratio flight-test experience 

of researchers at the NASA Flight Research Center, and the rocketry 

and space flight expertise of the Marshall Space Flight Center, whose 

experts assessed each proposal from the standpoint of technical feasibility and launch vehicle practicality. This multi-Center review strongly 

endorsed development of a modified delta planform, in part because 

the delta had inherently better stability characteristics during the high 

angle-of-attack reentry profile that any returning Shuttle would have to 

experience. Two families emerged as finalists: The 036 series, with small 

payload bays and three engines, and the 040 family, of similar planform 

but with larger payload bays and four engines. Then, in late January 

1972, MSC engineers evolved the 040C configuration: a three-engine 

design using new high-pressure engines. The 040C design became the 

baseline for subsequent Orbiter studies. While many questions remained 

over the final form that Shuttle’s launch system would take, with the 

040C study, the shape of the orbiter, and its all-important wing, was 

essentially fixed. Again, the flat-bottom delta had carried the day.109

Swing Wing: The Path to Variable Geometry

The notion of variable wing-sweeping dates to the earliest days of aviation and, in many respects, represents an expression of the “bird imitative” philosophy of flight that gave the ornithopter and other flexible 

wing concepts to aviation. Varying the sweep of a wing was first conceptualized as a means of adjusting longitudinal trim. Subsequently, 

109. Spacecraft Design Division, Summary of MSC Shuttle Configurations (External HO Tanks) (Houston: Manned Spacecraft Center, June 30, 1972, rev. ed.), passim. I thank the late Dr. Edward C. Ezell 

for making a copy of this document available for my research. The range of configurations and wind 

tunnel testing done in support of Shuttle development is in A. Miles Whitnah and Ernest R. Hillje, “Space 

Shuttle Wind Tunnel Testing Summary,” NASA Reference Publication 1125 (1984), esp. pp. 5–7. See 

also Alfred C. Draper, Melvin L. Buck, and William H. Goesch, “A Delta Shuttle Orbiter.” Astronautics & 

Aeronautics, vol. 9, No. 1 (Jan. 1971), pp. 26–35 (I acknowledge with gratitude the assistance and 

advice of the late Al Draper, while we both worked at Aeronautical Systems Division, Wright-Patterson 

AFB, in 1986–1987); Joseph Weil and Bruce G. Powers, “Correlation of Predicted and Flight Derived 

Stability and Control Derivatives with Particular Application to Tailless Delta Wing Configurations,” 

NASA TM-81361 (July 1981); and J.P. Loftus, Jr., et al. “The Evolution of the Space Shuttle Design,” 

a reference paper prepared for the Rogers Commission, 1986 (copy in NASA JSC History Office 

archives). The evolution of Shuttle configuration evolution is examined more broadly in Richard P. Hallion 

and James O. Young, “Space Shuttle: Fulfillment of a Dream,” in Hallion, ed., From Scramjet to the 

National Aero-Space Plane (1964–1986), vol. 2 of The Hypersonic Revolution: Case Studies in the 

History of Hypersonic Technology (Washington: USAF, 1998) pp. 947–1173.

1
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1

A time-lapse photograph of the Bell X-5, showing the range of its wing sweep. Note how the wing roots 

translated fore and aft to accommodate changes in center of lift with varying sweep angles. NASA.

variable-geometry advocates postulated possible use of asymmetric 

sweeping as a means of roll control. Lippisch, pioneer of tailless and delta 

design, likewise filed a patent in 1942 for a scheme of wing sweeping, 

but it was another German, Waldemar Voigt (the chief of advanced design 

for the Messerschmitt firm) who triggered the path to modern variable 

wing-sweeping. Ironically, at the time he did so, he had no plan to make use 

of such a scheme himself. Rather, he designed a graceful midwing turbojet 

swept wing fighter, the P 1101. The German air ministry rejected its development based upon assessments of its likely utility. Voigt decided to continue its development, planning to use the airplane as an in-house swept 

wing research aircraft, fitted with wings of varying sweep and ballasted 

to accommodate changes in center of lift.110

110. The best survey of v-g origins remains Robert L. Perry’s Innovation and Military Requirements: A 

Comparative Study, Rand Report RM-5182PR (Santa Monica: The Rand Corporation, 1967), upon 

which this account is based.
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By war’s end, when the Oberammergau plant was overrun by 

American forces, the P 1101 was over 80-percent complete. A technical team led by Robert J. Woods, a member of the NACA Aerodynamics 

Committee, moved in to assess the plant and its projects. Woods immediately recognized the value of the P 1101 program, but with a twist: he 

proposed to Voigt that the plane be finished with a wing that could be 

variably swept in flight, rather than with multiple wings that could be 

installed and removed on the ground. Woods’s advocacy, and the results 

of NACA variable-sweep tests by Charles Donlan of a modified XS-1 

model in the Langley 7-foot by 10-foot wind tunnel, convinced the NACA 

to support development of such an aircraft. In May 1949, the Air Force 

Air Materiel Command issued a contract covering development of two 

Bell variable sweep airplanes, to be designated X-5. They were effectively 

American-built versions of the P 1101, but with American, not German, 

propulsion, larger cockpit canopies for greater pilot visibility, and, of 

course, variable sweep wings that could range from 20 to 60 degrees.111

The first X-5 flew in June 1951 and within 5 weeks had demonstrated 

variable in-flight wing sweep to its maximum 60-degree aft position. 

Slightly over a year later, Grumman flew a prototype variable wing-sweep 

naval fighter, the XF10F-1 Jaguar. Neither aircraft represented a mature 

application of variable sweep design. The mechanism in each was heavy 

and complex and shifted the wing roots back and forth down the centerline of the aircraft to accommodate center of lift changes as the wing 

was swept and unswept. Each of the two had poor flying qualities unrelated to the variable-sweep concept, reflecting badly on their design. 

The XF10F-1 was merely unpleasant (its test pilot, the colorful Corwin 

“Corky” Meyer, tellingly recollected later “I had never attended a test 

pilots’ school, but, for me, the F10F provided the complete curriculum”), 

but the X-5 was lethal.112 It had a vicious pitch-up at higher-sweep angles, 

and its aerodynamic design ensured that it would have very great difficulty 

when it departed into a spin. The combination of the two led to the death 

of Air Force test pilot Raymond Popson in the crash of the second X-5 

111. The history of the X-5 is examined minutely in Warren E. Green’s The Bell X-5 Research Airplane (Wright-Patterson AFB: Wright Air Development Center, March 1954). For NACA work, see 

LRC staff, “Summary of NACA/NASA Variable-Sweep Research and Development Leading to the 

F-111 (TFX),” Langley Working Paper LWP-285 (Dec. 22, 1966).

112. Corwin H. Meyer, “Wild, Wild Cat: The XF10F,” 20th Symposium, The Society of Experimental Test Pilots, Beverly Hills, CA, Sept. 15, 1976.

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1 in 1953. More fortunate, NACA pilots completed 133 research flights in 

the first X-5 before retiring it in 1955.

The X-5 experience demonstrated that variable geometry worked, and 

the potential of combining good low-speed performance with high-speed 

supersonic dash intrigued military authorities looking at future interceptor and long-range strike aircraft concepts. Coincidentally, in the late 

1950s, Langley developed increasingly close ties with the British aeronautical community, largely a result of the personal influence of John Stack 

of Langley Research Center, who, in characteristic fashion, used his forceful personality to secure a strong transatlantic partnership. This partnership, best known for its influence upon Anglo-American V/STOL research 

leading to the Harrier strike fighter, influenced as well the course of 

variable-geometry research. Barnes Wallis of Vickers had conceptualized a 

sharply swept variable-geometry tailless design, the Swallow, but was not 

satisfied with the degree of support he was receiving for the idea within 

British aeronautical and governmental circles. Accordingly, he turned to 

the United States. Over November 13–18, 1958, Stack sponsored an AngloAmerican meeting at Langley to craft a joint research program, in which 

Wallis and his senior staff briefed the Swallow design.113 As revealed by 

subsequent Langley tunnel tests over the next 6 months, Wallis’s Swallow 

had many stability and control deficiencies but one significant attribute: 

its outboard wing-pivot design. Unlike the X-5 and Jaguar and other 

early symmetrical-sweep v-g concepts, the wing did not adjust for changing center of lift position by translating fore and aft along the fuselage 

centerline using a track-type approach and a single pivot point. Rather, 

slightly outboard of the fuselage centerline, each wing panel had its own 

independent pivot point. This permitted elimination of the complex track 

and allowed use of a sharply swept forebody to address at least some of 

the changes in center-of-lift location as the wings moved aft and forward. 

The remainder could be accommodated by control surface deflection and 

shifting fuel. Studies in Langley’s 7-foot by 10-foot tunnel led to refinement 

of the outboard pivot concept and, eventually, a patent to William J. Alford 

and E.C. Polhamus for its concept, awarded in September 1962. Wallis’s 

inspiration, joined with insightful research by Alford and Polhamus and 

113. For meeting, see LRC staff, “Summary of NACA/NASA Variable-Sweep Research and 

Development,” p. 8; and J.E. Morpurgo, Barnes Wallis: a Biography (Harmondsworth, UK: Penguin 

Books, 1973), p. 423. NASA Langley Photograph L58-771a, dated Nov. 13, 1958, documents 

the Stack-Wallis meeting; it is also catalogued as NASA LaRC image EL-2008-00001.
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61

followed by adaptation of a conventional “tailed” configuration (a critical necessity in the pre-fly-by-wire computer-controlled era), made 

variable wing sweep a practical reality.114 (Understandably, after returning to Britain, Wallis had mixed feelings about the NASA involvement. 

On one hand, he had sought it after what he perceived as a “go slow” 

approach to his idea in Britain. On the other, following enunciation 

of outboard wing sweep, he believed—as his biographer subsequently 

wrote—“The Americans stole his ideas.”)115

Thus, by the early 1960s, multiple developments—swept wings, 

high-performance afterburning turbofans, area ruling, the outboard 

wing pivot, low horizontal tail, advanced stability augmentation systems, to select just a few—made possible the design of variablegeometry combat aircraft. The first of these was the General Dynamics 

Tactical Fighter Experimental (TFX), which became the F-111. It was a 

troubled program, though, like most of the Century series that had preceded it (the F-102 in particular), this had essentially nothing to do with 

the adaptation of a variably swept wing. Instead, a poorly written specification emphasizing joint service over practical, attainable military utility resulted in development of a compromised design. The result was a 

decade of lost fighter time for the U.S. Navy, which never did receive the 

aircraft it sought, and a constrained Air Force program that resulted in 

the eventual development of a satisfactory strike aircraft—the F-111F—

but years late and at tremendous cost. Throughout the evolution of the 

F-111, NASA research proved of crucial importance to saving the program. NASA Langley, Ames, and Lewis researchers invested over 30,000 

114. LRC, “Summary of NACA/NASA Variable-Sweep Research;” see also William J. Alford, Jr., and 

William P. Henderson, “An Exploratory Investigation of Variable-Wing-Sweep Airplane Configurations,” 

NASA TM-X-142 (1959); William J. Alford, Jr., Arvo A. Luoma, and William P. Henderson, “WindTunnel Studies at Subsonic and Transonic Speeds of a Multiple-Mission Variable-Wing-Sweep Airplane 

Configuration,” NASA TM-X-206 (1959); and Gerald V. Foster and Odell A. Morris, “Aerodynamic 

Characteristics in Pitch at a Mach Number of 1.97 of Two Variable-Wing-Sweep V/STOL Configurations with Outboard Wing Panels Swept Back 75°,” NASA TM-X-322 (1960). 

115. Morpurgo, Wallis, p. 422, and Derek Wood, Project Cancelled: British Aircraft that Never Flew

(Indianapolis: The Bobbs-Merrill Company, Inc., 1975), pp. 182–195. After the Nov. 1958 meeting, 

NASA tunnel tests revealed very great deficiencies attending his tailless concept that Stack and others 

reported back to Vickers in June 1959. In short, the outboard pivot was but one element necessary 

for making a successful v-g aircraft. Others were provision for a conventional tail and design of a 

practicable airframe. In short, Wallis had an idea, but it took Alford and Polhamus and other NASA 

researchers to refine it and render it achievable.

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NASA’s Contributions to Aeronautics

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1 hours of wind tunnel test time in the F-111 (over 22,000 at Langley alone), 

addressing various shortcomings in its design, including excessive drag, 

lack of transonic and supersonic maneuverability, deficient directional 

stability, and inlet distortion that plagued its engine performance. As a 

result, the Air Force F-111 became a reliable weapon system, evidenced 

by its performance in Desert Storm, where it flew long-range strike missions, performed electronic jamming, and proved the war’s single most 

successful “tank plinker,” on occasion destroying upward of 150 tanks 

per night and 1,500 over the length of the 43-day conflict.116

From the experience gained with the F-111 program sprang the 

Grumman F-14 Tomcat naval fighter and the Rockwell B-1 bomber, 

both of which experienced fewer development problems, benefitting 

greatly from NASA tunnel and other analytical research.117 Emulating 

American variable-geometry development, Britain, France, and the Soviet 

Union undertook their own development efforts, spawning the experimental Dassault Mirage G (test-flown, though never placed in service), 

the multipartner NATO Tornado interceptor and strike fighter program, 

and a range of Soviet fighter and bomber aircraft, including the MiG23/27 Flogger, the Sukhoi Su-17/22 Fitter, the Su-24 Fencer, the Tupolev 

Tu-22M Backfire, and the Tu-160 Blackjack.118

Variable geometry has had a mixed history since; in the heyday of the 

space program, many proposals existed for tailored lifting body shapes 

deploying “switchblade” wings, and the variable-sweep wing was a prominent feature of the Boeing SST concept before its subsequent rejection. 

The tailored aerodynamics and power available with modern aircraft have 

rendered variable-geometry approaches less attractive than they once 

were, particularly because, no matter how well thought out, they invari116. NASA F-111 tunnel research, analysis, and support is detailed in Testimony of Edward C. Polhamus, in U.S. Senate, TFX Contract Investigation (Second Series): Hearings Before the Permanent 

Subcommittee on Investigations of the Committee on Government Operations, United States Senate, 

91st Congress, 2nd Session, Part 2 (Washington: GPO, 1970), pp. 339–363; for the F-111 in 

Desert Storm, see Tom Clancy with Gen. Chuck Horner (New York: G.P. Putnam’s Sons, 1999), pp. 

318, 417, 424, and 450.

117. See Joseph R. Chambers, Partners in Freedom: Contributions of the Langley Research Center 

to U.S. Military Aircraft of the 1990s, SP-2000-4519 (Washington; NASA, 2000), which treats 

these and other programs in great and authoritative detail.

118. Robert W. Kress, “Variable Sweep Wing Design,” AIAA Paper No. 83-1051 (1983) is an 

excellent survey. The Su-24 was clearly F-111 inspired, and the Tu-160 was embarrassingly similar 

in configuration to the American B-1.
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The Grumman F-14A Tomcat naval fighter marked the maturation of the variable wing-sweep concept. This is one was assigned to Dryden for high angle of attack and departure flight-testing. NASA.

ably involve greater cost, weight, and structural complexity. In 1945–1946, 

John Campbell and Hubert Drake undertook tests in the Langley Free 

Flight Tunnel of a simple model with a single pivot, so that its wing could 

be skewed over a range of sweep angles. This concept, which German 

aerodynamicists had earlier proposed in the Second World War, demonstrated “that an airplane wing can be skewed as a unit to angles as great as 

40° without encountering serious stability and control difficulties.”119 This 

concept, the simplest of all variable-geometry schemes, returned to the 

fore in the late 1970s, thanks to the work of Robert T. Jones, who adopted 

and expanded upon it to generate the so-called “oblique wing” design concept. Jones conceptualized the oblique wing as a means of producing a 

transonic transport that would have minimal drag and a minimal sonic 

boom; he even foresaw possible twin fuselage transports with a skewed 

wing shifting their relative position back and forth. Tests with a subscale 

turbojet demonstrator, the AD-1 (for Ames-Dryden), at the Dryden Flight 

Research Center confirmed what Campbell and Drake had discovered 

119. John P. Campbell and Hubert M. Drake, “Investigation of Stability and Control Characteristics 

of an Airplane Model with Skewed Wing in the Langley Free-Flight Tunnel,” NACA TN-1208 (May 

1947), p. 10.

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1 nearly four decades previously, namely that at moderate sweep angles 

the oblique wing possessed few vices. But at higher sweep angles near 

60 degrees, its deficits became more pronounced, calling into question 

whether its promise could ever actually be achieved.120 On the whole, 

the variable-geometry wing has not enjoyed the kind of widespread success that its adherents hoped. While it may be expected that, from time 

to time, variable sweep aircraft will be designed and flown for particular purposes, overall the fixed conventional planform, outfitted with 

all manner of flaps and slats and blowing, sucking, and perhaps even 

warping technology, continues to prevail.

The Quest for Refinement

By the end of the 1960s, the “classic” era of aircraft design was arguably at an end. As exemplars of the highest state of aviation technology, 

the piston engine had given way to the gas turbine, the wood-and-fabric 

aircraft to the all-metal, the straight wing had given way to the swept and 

delta. Aircraft flight speeds had risen from a mere 40 mph at the time 

of the Wright brothers to over 100 times as fast, as the X-15A-2 demonstrated when it streaked to Mach 6.70 (4,520 mph) in October 1967, 

piloted by Maj. William J. Knight. Fighters, by that time, had been flying 

on a Mach 2 plateau for a decade and transports on a Mach 0.82 plateau 

for roughly the same amount of time. In space, Americans were basking 

in the glow of the recent Apollo triumph, where a team of astronauts, led 

by former NACA–NASA research pilot Neil Armstrong—a Round One 

and Round Two veteran whose experience included both the X-1 and 

the X-15—journeyed to the Moon, landed two of their number upon it, 

and then returned to Earth.

Such accomplishments hardly meant that the frontiers of the sky were 

closing, or that NASA had little to do. Indeed, in some respects, it was 

facing even greater challenges: conducting comprehensive aeronautical 

research at a time when, increasingly, more people identified it with space 

than aeronautics and when, in the aftermath of the Apollo success, monies were increasingly tight. Added to this was a dramatically transforming 

world situation: increasing tension in the Middle East, a growing Soviet 

threat, rising oil prices, open concern over environmental stewardship, 

120. Richard P. Hallion and Michael H. Gorn, On the Frontier: Experimental Flight at NASA 

Dryden (Washington: Smithsonian Books, 2002), pp. 256–260, and personal recollections of the 

program from the time.
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and a national turning away from the reflexive perception that limitless 

technological progress was both a given and a good thing.

Within this framework, NASA work increasingly turned to achieving 

efficiencies: more fuel-efficient and energy-efficient civilian flight, and 

more efficient military systems. It was not NASA’s business to, per se, 

design new aircraft, but, as NACA–NASA history amply demonstrated, 

the Agency’s mark could be found on many aircraft and their innovations. Little things counted for much. When, for example, NACA HighSpeed Flight Research Station pilots flew a Douglas D-558-1 Skystreak 

modified with a row of small vortex generators (little rectangular fins 

of 0.5-inch chord standing vertically like a row of razor blades) on its 

upper wing surface, they hardly expected that such a small energyimparting modification would so dramatically improve its transonic 

handling qualities that rows of vortex generators would become a commonly recognized feature on many aircraft, including such “classics” as 

the B-52, the 707, and the A-4.121 In the post-1970 period, NASA assiduously pursued three concepts related to swept wing and delta flight, in 

hopes that each would pay great dividends: the supercritical wing, the 

winglet, and the arrow wing.122 All had roots embedded and nourished 

in the earliest days of the supersonic and swept/delta revolution. Each 

reflected Whitcomb’s passion—indeed obsession, in its most positive 

sense—with minimizing interference effects and achieving the greatest 

possible aerodynamic efficiency without incurring performance-robbing 

complexity. Many had researched configurations approaching the purity 

of the arrow wing, but it was Whitcomb who first actually achieved such 

a configuration, as part of Langley’s Supersonic Transport study effort.

Long a subject of individual research and thought, Langley’s institutional SST studies had begun in 1958, when the ever-enthusiastic John 

Stack formed a Supersonic Transport Research Committee (STRC). It 

evaluated the maturity of various disciplines—particularly the “classics” 

of aerodynamics, structures, propulsion, and controls—and then forecast the overall feasibility of a Supersonic Transport. The Stack team 

presented the results of their studies to the head of the Federal Aviation 

121. De E. Beeler, Donald R. Bellman, and John H. Griffith, “Flight Determination of the Effects of 

Wing Vortex Generators on the Aerodynamic Characteristics of the Douglas D-558-I Airplane,” 

NACA RM-L51A23 (1951).

122. All three bore the imprint of Richard Whitcomb and thus, in this survey, are not examined in 

detail, since his work is more thoroughly treated in a companion essay by Jeremy Kinney.

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1 Administration (FAA), Elwood Quesada, a retired Air Force general, 

in December 1959. Their report, issued the following year, concluded: 

“the state of the art appears sufficiently advanced to permit the design 

of an airplane at least marginally capable of performing the supersonic transport mission.”123 NASA swiftly ramped up to match growing 

interest in the FAA in such aircraft; within a decade, SST-focused 

research would constitute over a quarter of all NASA aeronautics research 

undertaken at the Langley, Ames, and Lewis Centers.124

Given that the British and French subsequently designed the Mach 

2+ Concorde, and the Soviets the Tupolev Tu-144, NASA Langley’s technological optimism in 1959–1960 was, within limits, technically well 

justified, and such optimism infused Washington’s political community 

as well. In March 1966, President Lyndon Johnson announced that the 

first American SST, designed to cruise at Mach 2.7, would fly at decade’s 

end and enter commercial service in 1974.125 But such expectations would 

prove overly optimistic. As Mach number rose, so too did a number of 

daunting technical challenges encountered by the more ambitious aircraft American SST proponents favored. Assessing the technology alone 

did not address the serious questions—research and development investment, production costs, operating economics, and environmental concerns, for example—such aircraft would pose and would limit the airline 

acceptance (and, hence, market success) of even the “modest” Concorde 

and Tu-144. Air transport constitutes a system of systems, and excellence 

in some does not guarantee or imply excellence overall. Political support, 

strongly bipartisan over the Kennedy-Johnson era, withered in the Nixon 

123. LRC staff, “The Supersonic Transport—A Technical Summary,” NASA TN-D-423 (1960), p. 

93; this was the summary report of the briefings presented the previous fall to Quesada. NASA 

research on supersonic cruise is the subject of a companion essay in this study, by William Flanagan, and Whitcomb’s work is detailed in the previously cited Kinney study in this volume.

124. In FY 1968, NASA expended $10.8 million in then-year dollars on SST research at Langley, 

Ames, and Lewis, against a total aeronautics research expenditure of $42.9 million at those 

Centers. See Testimony of James E. Webb in U.S. Senate, Aeronautical Research and Development 

Policy: Hearings Before the Committee on Aeronautical and Space Sciences, United States Senate, 

90th Congress, 1st Session (Washington: GPO, 1967), p. 39.

125. Lyndon B. Johnson, “President’s Message on Transportation,” Mar. 2, 1966, reprinted in Legislative Reference Service of the Library of Congress, Policy Planning for Aeronautical Research and 

Development: Staff Report Prepared for Use of the Committee on Aeronautical and Space Sciences 

United States Senate by the Legislative Reference Service Library of Congress, Document No. 90, 

U.S. Senate, 89th Congress, 2nd Session (Washington: GPO, 1966), pp. 50–51.
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years as technical and other challenges arose, and a re-action against 

the SST set in, fueled by questions over the value of high technology and 

reaction to the long and costly war in Southeast Asia.126

From the standpoint of aircraft design, from Langley’s interest emerged 

a series of Supersonic Commercial Air Transport (SCAT) design studies, 

most of which incorporated variable-geometry planforms reflecting a 

growing popular wisdom that future military or civilian supersonic cruise 

designs would necessarily incorporate such wings. Whitcomb, focused 

on simplicity and efficiency, demurred, preferring instead a sharply swept 

arrow configuration, the SCAT-4, which he had derived. It drew upon a 

two-decade tradition of Langley swept and delta studies running through 

those of Clinton E. Brown and F. Edward McLean in the 1950s, back to 

the thin swept and delta research manifested in Robert T. Jones’s original concepts in 1944–1945. Though he was not successful at the time at 

selling his vision of what such an aircraft should be (and, in fact, left the 

Stack SST study effort as a result), in time the fixed wing predominated. 

In 1964, a Langley team comprised of Harry Carlson, Roy Harris, Ed 

McLean, Wilbur Middleton, and A. Warner Robins derived a fixed wing 

variant of the variable-sweep SCAT-15, generating an elegant slender 

arrow wing called the SCAT-15F. SCAT-15F had an incredible lift-to-drag 

ratio of 9.3 at Mach 2.6, well beyond what previous analysis and thought 

had deemed possible, though it also had serious low-speed pitch-up and 

deep-stall tendencies that triggered intensive investigations by researchers using the Langley Full-Scale Tunnel.127 Out of this came a revised 

SCAT-15F configuration, with leading edge flaps, wing notches, area-andcamber-increasing Fowler flaps, and a small, horizontal tail, all of which 

worked to make it a much more acceptable planform. The development 

126. For various perspectives on Anglo-French-Soviet-American SST development, see Kenneth Owen, 

Concorde: Story of a Supersonic Pioneer (London: Science Museum, 2001); Howard Moon, Soviet 

SST: The Technopolitics of the Tupolev Tu-144 (New York: Orion Books, 1989); R.E.G. Davies, Supersonic (Airliner) Non-Sense: A Case Study in Applied Market Research (McLean, VA: Paladwr Press, 

1998); Mel Horwitch, Clipped Wings: The American SST Conflict (Cambridge: The MIT Press, 1982); 

and Eric M. Conway, High-Speed Dreams: NASA and the Technopolitics of Supersonic Transportation, 

1945–1999 (Baltimore: The Johns Hopkins Press, 2005).

127. Deep stall is a dangerous condition wherein an airplane pitches to a high angle of attack, stalls, 

and then descends in a stabilized stalled attitude, impervious to corrective control inputs. It is more 

typically encountered by swept wing T-tail aircraft, and one infamous British accident, to a BAC 1-11 

airliner, claimed the life of a crack flight-test crew captained by the legendary Mike Lithgow, an early 

supersonic and sweptwing pioneer.

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1 of the high supersonic L/D fixed wing eventually led Boeing (winner of 

the Government’s SST design competition) to abandon variable-sweep in 

favor of a highly refined small-tailed delta, for its final SST proposal, 

though congressional refusal to furnish needed developmental monies 

brought the American SST development effort to a sorry end.128 It did not, 

however, end interest in similar configurations for a range of other missions. Today, in an era of vastly different technology, with much higherperforming engines, better structures, and better means of modeling and 

simulating the aerodynamic and propulsive performance of such designs, 

tailored fixed arrow wing configurations are commonplace for future 

advanced high-speed civil and military aircraft applications.

As the American SST program, plagued by controversy and 

numerous wounds (many self-inflicted), died amid performance and 

environmental concerns, Whitcomb increasingly turned his attention to 

the transonic, thereby giving to aviation one of its most compelling images, 

that of the graceful supercritical wing and, of less aesthetic appeal but 

no less significance, the wingtip winglet. Both, in various forms, became 

standard design elements of future civil and military transport design 

and are examined elsewhere (by historian Jeremy Kinney) in this work.

128. Langley’s SCAT studies are summarized in David A. Anderton, Sixty Years of Aeronautical 

Research, 1917–1977, EP-145 (Washington: NASA, 1978), pp. 54–58. Relevant reports on 

specific configurations and predecessors include: Donald D. Baals, Thomas A. Toll, and Owen G. 

Morris, “Airplane Configurations for Cruise at a Mach Number of 3,” NACA RM-L58E14a (1958); 

Odell A. Morris and A. Warner Robins, “Aerodynamic Characteristics at Mach Number 2.01 of an 

Airplane Configuration Having a Cambered and Twisted Arrow Wing Designed for a Mach Number 

of 3.0,” NASA TM-X-115 (1959); Cornelius Driver, M. Leroy Spearman, and William A. Corlett, 

“Aerodynamic Characteristics at Mach Numbers From 1.61 to 2.86 of a Supersonic Transport Model 

With a Blended Wing-Body, Variable-Sweep Auxiliary Wing Panels, Outboard Tail Surfaces, and a 

Design Mach Number of 2.2,” NASA TM-X-817 (1963); Odell A. Morris and James C. Patterson, Jr., 

“Transonic Aerodynamic Characteristics of Supersonic Transport Model With a Fixed, Warped Wing 

Having 74° Sweep,” NASA TM-X-1167 (1965); Odell A. Morris, and Roger H. Fournier, “Aerodynamic Characteristics at Mach Numbers 2.30, 2.60, and 2.96 of a Supersonic Transport Model 

Having Fixed, Warped Wing,” NASA TM-X-1115 (1965); A. Warner Robins, Odell A. Morris, and 

Roy V. Harris, Jr., “Recent Research Results in the Aerodynamics of Supersonic Vehicles,” AIAA Paper 

65-717 (1965); Donald D. Baals, A. Warner Robins, and Roy V. Harris, Jr., “Aerodynamic Design 

Integration of Supersonic Aircraft,” AIAA Paper 68-1018 (1968); Odell A. Morris, Dennis E. Fuller, and 

Carolyn B. Watson, “Aerodynamic Characteristics of a Fixed Arrow-Wing Supersonic Cruise Aircraft 

at Mach Numbers of 2.30, 2.70, and 2.95,” NASA TM-78706 (1978); and John P. Decker and 

Peter F. Jacobs, “Stability and Performance Characteristics of a Fixed Arrow Wing Supersonic Transport 

Configuration (SCAT 15F-9898) at Mach Numbers from 0.60 to 1.20,” NASA TM-78726 (1978).
Case 1 | Sweep and Swing: Reshaping the Wing for the Jet and Rocket Age

69

As for the arrow wing, military exigency and the Cold War combined to ensure that studies of this most promising configuration 

spawned the “cranked arrow wing” of the late 1970s. Following cancellation of the national SST effort, NASA researchers continued studying supersonic cruise for both military and civil applications, under the 

guise of a new study effort, the Advanced Supersonic Technology (AST) 

effort. AST was succeeded by another Langley-run cruise-focused effort, 

the Supersonic Cruise Aircraft Research (SCAR, later shortened to SCR) 

program. SCR lasted until 1982, when NASA terminated it to focus more 

attention and resources on the already troubled Shuttle program. But 

meantime, it had spawned the Supersonic Cruise and Maneuver Prototype 

(SCAMP), a derivative of the F-16 designed to cruise at supersonic speeds. 

Its “cranked arrow” wing, blending a 70-degree swept inboard leading 

edge and a 50-degree swept outboard leading edge, looked deceptively 

simple but embodied sophisticated shaping and camber (reflecting the 

long legacy of SCAT studies, particularly the refinement of the SCAT15F), with leading edge vortex flaps to improve both transonic and lowspeed performance. General Dynamics’ F-16 designer Harry J. Hillaker 

adopted the planform for a proposed strike fighter version of the F-16 

because it reduced supersonic wave drag, increasing the F-16’s potential 

combat mission radius by as much as 65 percent and more than doubling 

its permissible angle-of-attack range as well. In the early 1980s, SCAMP, 

now designated the F-16XL, competed with the prototype F-15E Strike 

Eagle at Edwards Air Force Base for an Air Force deep-strike fighter contract. But the F-16XL was too small an airplane to win the completion; 

with greater internal fuel and volume, the larger Strike Eagle offered 

more growth potential and versatility. The two F-16XL aircraft, among 

the most beautiful ever flown, remained at Edwards, where they flew a 

variety of research missions at NASA Dryden, refining understanding of 

the complex flows around cranked arrow profiles and addressing such 

technical issues as the possibility of supersonic laminar flow control by 

using active suction. Interest in the cranked arrow has persisted, as it 

remains a most attractive design option for future supersonic cruise aircraft, whether piloted or not, both civil and military.129

129. Harry J. Hillaker, “The F-16: A Technology Demonstrator, a Prototype, and a Flight Demonstrator,” AIAA Paper No. 83-1063 (1983). The “XL” designation for the cranked-arrow F-16 reflected 

Harry Hillaker’s passionate interest in golf, for it echoed the name of a particularly popular longdistance golf ball, the Top Flite XL. See also Chambers, Innovation in Flight, pp. 42, 48, 58–59.

1
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1 By the end of the 1980s, for military aircraft, concern over aerodynamic shaping of aircraft was beginning to take second place behind 

concern over their electromagnetic signature. Where something such 

as the blended wing-body delta SR-71 possessed an innate purity and 

beauty of form, inherent when aerodynamics is given the position of 

primacy in aircraft design, something such as the swept wing, V-tail 

F-117 stealth fighter did not: all angles and panels, it hardly looked aerodynamic, and, indeed, it had numerous deficits cured only by its being 

birthed in the electronic fly-by-wire and composites era. But in other 

aspects it performed with equal brilliance: not the brilliance of Mach 

3+, but the quiet brilliance of penetrating a high-threat integrated air 

defense network, attacking a key target, and escaping without detection.

For the future of the swept surface, one had to look elsewhere, 

back to the transonic, where it could be glimpsed in the boldly imaginative lines of the Blended Wing-Body (BWB) transport. Conceived by 

Robert H. Liebeck, a gifted Boeing designer who had begun his career at 

Douglas, where he worked with the legendary A.M.O. Smith, the BWB 

represented a conception of pure aerodynamic efficiency predating 

NASA, the NACA that had preceded it, and even, indeed, Jack Northrop 

and the Horten brothers. It hearkened back to the earliest concepts for 

Nurflügeln (flying wings) by Hugo Junkers before the First World War, 

the first designer to appreciate how one could insightfully incorporate 

the cantilever all-metal structure to achieve a pure lifting surface.130

Conceived while Liebeck worked for McDonnell-Douglas in the latter 

years before its own merger with Boeing, the graceful BWB was not 

strictly a flying wing but, rather, a hybrid wing-body combination whose 

elegant high aspect ratio wing blended smoothly into a wide, flat-bottom 

fuselage, the wings sprouting tall winglets at their tips for lateral control, thus differing significantly from earlier concepts such as the Boeing 

“Spanloader” and the Horten, Armstrong-Whitworth, and Northrop flying wings. Early design conceptions envisioned upward of 800 passen130. For Junkers, see Hugo Junkers, Gleitflieger mit zur Aufnahme von nicht Auftrieg erzeugen Teilen 

dienenden Hohlkörpen, Patentschrift Nr. 253788, Klasse 77h, Gruppe 5 (Berlin: Reichspatentamt, 

Nov. 14, 1912). For Liebeck, see Robert H. Liebeck, Mark A. Page, Blaine K. Rawdon, Paul W. 

Scott, and Robert A. Wright, “Concepts for Advanced Subsonic Transports,” NASA CR-4624 

(1994); Robert H. Liebeck, “Design of the Blended Wing Body Subsonic Transport,” Journal of 

Aircraft, vol. 41, no. 1 (Jan.–Feb. 2004). pp. 10–25; and Chambers, Innovation in Flight, pp. 

86–92.
Case 1 | Sweep and Swing: Reshaping the Wing for the Jet and Rocket Age

71

gers flying in a three-engine, double-deck, 823,000-pound, manta-shaped 

BWB (spanning 289 feet with a length of 161 feet), cruising across the 

globe at Mach 0.85. Subsequent analysis resulted in a smaller design 

sized for 450 passengers, the BWB-450, which served as the baseline 

for later research and evaluation, which concluded that the most suitable role for the BWB might be for a range of global heavy-lift multipurpose military missions rather than passenger-carrying.131 Extensive 

studies by NASA Langley and Lewis researchers; McDonnell-Douglas 

(now Boeing) BWB team members; and academic researchers from 

Stanford University, the University of Southern California, Clark Atlanta 

University, and the University of Florida confirmed the aerodynamic 

and propulsive promise inherent in the BWB, particularly its potential to carry great loads at transonic speeds over global distances with 

unprecedented aerodynamic and energy efficiency, resulting in potentially 30-percent better fuel economy than that achievable by traditional 

“tube and wing” airlifters.132

 These and many other studies, including tests by Boeing and the 

United States Air Force, encouraged the next logical step: developing a 

subscale unmanned aerial vehicle (UAV) to assess the low-speed flightcontrol characteristics of the BWB in actual flight. This became the 

X-48B, a 21-foot span, 8.5-percent scale UAV testbed of the BWB-450 

configuration, powered by three 240-pound thrust Williams turbojets. 

Boeing had Cranfield Aerospace, Ltd., in Great Britain build two X-48Bs 

for the company’s Phantom Works. After completion, the first X-48B completed 250 hours of tunnel tests in the Langley Full-Scale Tunnel (run 

by Old Dominion University) in May 2006. Readying the BWB for flight 

131. These included heavy-lift cargo, air-refueling, and other military missions rather than use as 

a civil airliner. See NASA LRC, “The Blended-Wing-Body: Super Jumbo Jet Concept Would Carry 

800 Passengers,” NASA Facts, FS-1997-07-24-LaRC (July 1997); and NASA LRC, “The Blended 

Wing Body: A Revolutionary Concept in Aircraft Design,” NASA Facts, FS-2001-04-24-LaRC (Apr. 

2001). For an early appreciation of the military value of BWB designs, see Gene H. McCall, et 

al., Aircraft & Propulsion, a volume in the New World Vistas: Air and Space Power for the 21st 

Century series (Washington: HQ USAF Scientific Advisory Board, 1995), p. 6.

132. Robert H. Liebeck, Mark A. Page and Blaine K. Rawdon, “Blended-Wing-Body Subsonic 

Commercial Transport,” AIAA Paper 98-0438 (1998); Sean Wakayama, “Multidisciplinary Design 

Optimization of the Blended-Wing-Body,” AIAA Paper 98-4938 (1998); Dino Roman, J.B. Allen, 

and Robert H. Liebeck, “Aerodynamic Design Challenges of the Blended-Wing-Body Subsonic 

Transport,” AIAA Paper 2000-4335 (2000). Fuel economy figure from Dryden Flight Research 

Center, “X-24B Blended Wing-Body” (Apr. 2, 2009).

1
NASA’s Contributions to Aeronautics

72

1

The NASA F-16XL cranked-arrow research aircraft aloft over the Dryden Flight Research Center 

on December 16, 1997. NASA.

consumed another year until, on July 20, 2007, the second example took 

to the air at Dryden, becoming the first of the X-48B testbeds to fly. By 

the end of the year, it had completed five research flights. Subsequent 

testing explored its stability and control at increasing angles of attack 

(to as great as 16-degree AoA), pointing to possible ways of furnishing 

improved controllability at even higher angles of attack.133 Time will 

tell if the world’s skies will fill with blended wing-body shapes. But to 

those who follow the technology of the sky, if seemingly fantastic, it is 

well within the realm of the possible, given the history of the swept and 

delta wings—and NACA–NASA’s role in furthering them.

In conclusion, the invention of the swept and delta wing blended 

creative and imaginative analysis and insight, great risk, and steadfast 

research. If in introspect their story has a clarity and a cohesiveness 

that was not necessarily visible to those at the time, it is because time 

has stripped the story to its essence. It is unfortunate that the perception that America was “given” (or “took”) the swept and delta wing in 

full-blown maturity from the laboratories of the Third Reich possesses 

133. DFRC, “X-24B Blended Wing-Body” (Apr. 2, 2009).
Case 1 | Sweep and Swing: Reshaping the Wing for the Jet and Rocket Age

73

such persistency, for it obscures the complex roots of the swept and delta 

wing in both Europe and America, the role of the NACA and NASA in 

maturing them, and, at heart, the accomplishments of successive generations of Americans within the NACA–NASA and elsewhere who worked 

to take what were, in most cases, very immature concepts and turn them 

into practical reality. Doing so required achieving many other things, 

among which were securing a practical means of effective longitudinal control at transonic speeds (the low, all-moving, and powered tail), 

reducing transonic drag rise, developing stability augmentation systems, and refining aircraft handling qualities. Defeating the transonic 

drag “hump”; reducing pitch-up to nuance, not nuisance; and overcoming the danger of inertial coupling were all crucial to ensuring that the 

swept and delta wing could fulfill their transforming promise. Once 

achieved, that gave to the world the means to fulfill the promise of the 

jet engine. As a result, international security and global transportation 

patterns were dramatically altered and a new transnational global consciousness born. It is something that workers of the NACA past, and 

NASA past, present, and future, can look back upon with a sense of both 

pride and accomplishment.

1
NASA’s Contributions to Aeronautics

74

1 Recommended Additional Readings

Reports, Papers, Articles, and Presentations:

Marvin Abramovitz, Stanley F. Schmidt, and Rudolph D. Van Dyke, Jr., 

“Investigation of the Use of a Stick Force Proportional to Pitching 

Acceleration for Normal-Acceleration Warning,” NACA RM-A53E21 

(1953).

James J. Adams and James B. Whitten, “Tests of a Centering Spring Used 

as an Artificial Feel Device on the Elevator of a Fighter Airplane,” 

NACA RM-L52G16 (1952).

William J. Alford, Jr., and Thomas B. Pasteur, Jr., “The Effects of Changes 

in Aspect Ratio and Tail Height on the Longitudinal Stability 

Characteristics at High Subsonic Speeds of a Model with a Wing 

Having 32.6° Sweepback,” NACA RM-L53L09 (1953).

William J. Alford, Jr., and William P. Henderson, “An Exploratory 

Investigation of Variable-Wing-Sweep Airplane Configurations,” 

NASA TM-X-142 (1959).

William J. Alford, Jr., Arvo A. Luoma, and William P. Henderson, “WindTunnel Studies at Subsonic and Transonic Speeds of a MultipleMission Variable-Wing-Sweep Airplane Configuration,” NASA 

TM-X-206 (1959).

H. Julian Allen and A.J. Eggers, Jr., “A Study of the Motion and 

Aerodynamic Heating of Ballistic Missiles Entering the Earth’s 

Atmosphere at High Supersonic Speeds,” NACA TR-1381 (1953).

Ames Research Center staff, “Collected Works of Robert T. Jones,” NASA 

TM-X-3334 (1976).

Raymond F. Anderson, “Determination of the Characteristics of Tapered 

Wings,” NACA Report No. 572 (1936).

William E. Andrews, Thomas R. Sisk, and Robert W. Darville, 

“Longitudinal Stability Characteristics of the Convair YF-102 

Airplane Determined from Flight Tests,” NACA RM-H56I17 (1956).
Case 1 | Sweep and Swing: Reshaping the Wing for the Jet and Rocket Age

75

Johnny G. Armstrong, “Flight Planning and Conduct of the X-24B 

Research Aircraft Flight Test Program,” Air Force Flight Test Center 

TR-76-11 (1977).

Theodore G. Ayers and James B. Hallissy, “Historical Background and 

Design Evolution of the Transonic Aircraft Technology Supercritical 

Wing,” NASA TM-81356 (1981).

Donald D. Baals, Thomas A. Toll, and Owen G. Morris, “Airplane 

Configurations for Cruise at a Mach Number of 3,” NACA 

RM-L58E14a (1958).

Donald D. Baals, A. Warner Robins, and Roy V. Harris, Jr., “Aerodynamic 

Design Integration of Supersonic Aircraft,” AIAA Paper 68-1018 (1968).

De E. Beeler, Donald R. Bellman, and John H. Griffith, “Flight 

Determination of the Effects of Wing Vortex Generators on the 

Aerodynamic Characteristics of the Douglas D-558-I Airplane,” NACA 

RM-L51A23 (1951).

Morgan M. Blair, “Evolution of the F-86,” AIAA Paper 80-3039 (1980).

Jack D. Brewer and Jacob H. Lichtenstein, “Effect of Horizontal Tail on 

Low-Speed Static Lateral Stability Characteristics of a Model Having 

45° Sweptback Wing and Tail Surfaces,” NACA TN-2010 (1950).

R. Porter Brown, Robert G. Chilton, and James B. Whitten, “Flight 

Investigation of a Mechanical Feel Device in an Irreversible Elevator 

Control System of a Large Airplane,” NACA Report No. 1101 (1952).

Adolf Busemann, “Aerodynamische Auftrieb bei Überschallgeschwindigkeit,” 

Luftfahrtforschung, vol. 12, no. 6 (Oct. 3, 1935).

Adolf Busemann, “Compressible Flow in the Thirties,” Annual Review 

of Fluid Mechanics, vol. 3 (1971).

John P. Campbell and Hubert M. Drake, “Investigation of Stability and 

Control Characteristics of an Airplane Model with Skewed Wing in 

the Langley Free-Flight Tunnel,” NACA TN-1208 (May 1947).

1
NASA’s Contributions to Aeronautics

76

1 George E. Cooper and Robert C. Innis, “Effect of Area-Suction-Type 

Boundary-Layer Control on the Landing-Approach Characteristics 

of a 35° Swept-Wing Fighter,” NACA RM-A55K14 (1957).

R.M. Cross, “Characteristics of a Triangular-Winged Aircraft: 2: Stability 

and Control,” in NACA, Conference on Aerodynamic Problems of 

Transonic Airplane Design (1947).

Richard E. Day and Donald Reisert, “Flight Behavior of the X-2 Research 

Airplane to a Mach Number of 3.20 and a Geometric Altitude of 

126,200 Feet,” NACA TM-X-137 (1959).

John P. Decker and Peter F. Jacobs, “Stability and Performance 

Characteristics of a Fixed Arrow Wing Supersonic Transport 

Configuration (SCAT 15F-9898) at Mach Numbers from 0.60 to 

1.20,” NASA TM-78726 (1978).

Hubert M. Drake and John R. Carden, “Elevator-Stabilizer Effectiveness 

and Trim of the X-1 Airplane to a Mach Number of 1.06,” NACA 

RM-L50G20 (1950).

Hubert M. Drake and Wendell H. Stillwell, “Behavior of the Bell X-1A 

Research Airplane During Exploratory Flights at Mach Numbers 

Near 2.0 and at Extreme Altitude,” NACA RM-H55G25 (1955).

Hubert M. Drake, Thomas W. Finch, and James R. Peele, “Flight 

Measurements of Directional Stability to a Mach Number of 1.48 for 

an Airplane Tested with Three Different Vertical Tail Configurations,” 

NACA RM-H55G26 (1955).

Alfred C. Draper, Melvin L. Buck, and William H. Goesch, “A Delta 

Shuttle Orbiter.” Astronautics & Aeronautics, vol. 9, no. 1 (Jan. 1971).

Cornelius Driver, M. Leroy Spearman, and William A. Corlett, 

“Aerodynamic Characteristics at Mach Numbers From 1.61 to 2.86 of 

a Supersonic Transport Model With a Blended Wing-Body, VariableSweep Auxiliary Wing Panels, Outboard Tail Surfaces, and a Design 

Mach Number of 2.2,” NASA TM-X-817 (1963).
Case 1 | Sweep and Swing: Reshaping the Wing for the Jet and Rocket Age

77

Alfred J. Eggers, Jr., and Clarence A. Syvertson, “Aircraft Configurations 

Developing High Lift-Drag Ratios at High Supersonic Speeds,” NACA 

RM-A55L05 (1956).

Carlo Ferrari, “Recalling the Vth Volta Congress: High Speeds in 

Aviation,” Annual Review of Fluid Mechanics, vol. 28 (1996).

Jack Fischel and Jack Nugent, “Flight Determination of the Longitudinal 

Stability in Accelerated Maneuvers at Transonic Speeds for the 

Douglas D-558-II Research Airplane Including the Effects of an 

Outboard Wing Fence,” NACA RM-L53A16 (1953).

Jack Fischel, “Effect of Wing Slats and Inboard Wing Fences on the 

Longitudinal Stability Characteristics of the Douglas D-558-II 

Research Airplane in Accelerated Maneuvers at Subsonic and 

Transonic Speeds,” NACA RM-L53L16 (1954).

Jack Fischel and Cyril D. Brunn, “Longitudinal Stability Characteristics 

in Accelerated Maneuvers at Subsonic and Transonic Speeds of the 

Douglas D-558-II Research Airplane Equipped with a Leading-Edge 

Wing Chord-Extension,” NACA RM-H54H16 (1954).

Jack Fischel and Donald Reisert, “Effect of Several Wing Modifications 

on the Subsonic and Transonic Longitudinal Handling Qualities of 

the Douglas D-558-II Research Airplane,” NACA RM-H56C30 (1956).

Gerald V. Foster and Odell A. Morris, “Aerodynamic Characteristics in 

Pitch at a Mach Number of 1.97 of Two Variable-Wing-Sweep V/

STOL Configurations with Outboard Wing Panels Swept Back 75°,” 

NASA TM-X-322 (1960).

Bernard Göthert, “High-Speed Measurements on a Swept-Back Wing 

(Sweepback Angle φ = 35°),” NACA TM-1102 (1947).

Charles F. Hall and John C. Heitmeyer, “Aerodynamic Study of a 

Wing-Fuselage Combination Employing a Wing Swept Back 63°—

Characteristics at Supersonic Speeds of a Model with the Wing 

Twisted and Cambered for Uniform Load,” NACA RM-A9J24 (1950).

1
NASA’s Contributions to Aeronautics

78

1 Charles F. Hall, “Lift, Drag, and Pitching Moment of Low-Aspect Ratio 

Wings at Subsonic and Supersonic Speeds,” NACA RM-A53A30 (1953).

Richard P. Hallion, “Convair’s Delta Alpha,” Air Enthusiast Quarterly, 

no. 2 (1976).

Richard P. Hallion, “Lippisch Gluhareff, and Jones: The Emergence of 

the Delta Planform and the Origins of the Sweptwing in the United 

States,” Aerospace Historian, vol. 26, no. 1 (Mar. 1979).

Robert G. Hoey and Capt. Iven C. Kincheloe, USAF, “ARDC F-104A 

Stability and Control,” AFFTC TR-58-14 (1958).

Capt. Slayton L. Johns, USAF, and Capt. James W. Wood, USAF, “ARDC 

F-104A Stability and Control with External Stores,” AFFTC TR-58-

14 Addendum 1 (July 1959).

Robert T. Jones, “Properties of Low-Aspect-Ratio Pointed Wings at Speeds 

Below and Above the Speed of Sound,” NACA TN-1032 (1946).

Robert T. Jones, “Wing Planforms for High-Speed Flight,” NACA TN-1033 

(1946).

Robert T. Jones, “Characteristics of a Configuration with a Large Angle 

of Sweepback,” in NACA, Conference on Aerodynamic Problems of 

Transonic Airplane Design (1947).

Harold F. Kleckner, “Preliminary Flight Research on an All-Movable 

Horizontal Tail as a Longitudinal Control for Flight at High Mach 

Numbers,” NACA ARR-L5C08 (March 1945).

Harold F. Kleckner, “Flight Tests of an All-Movable Horizontal Tail with 

Geared Unbalancing Tabs on the Curtiss XP-42 Airplane,” NACA 

TN-1139 (1946).

Robert W. Kress, “Variable Sweep Wing Design,” AIAA Paper No. 83-1051 

(1983).
Case 1 | Sweep and Swing: Reshaping the Wing for the Jet and Rocket Age

79

1st Lt. David C. Leisy and Capt. Hugh P. Hunerwadel, “ARDC F-100D 

Category II Performance Stability and Control Tests,” AFFTC TR-58-

27 (1958).

Jacob H. Lichtenstein, “Experimental Determination of the Effect of 

Horizontal-Tail Size, Tail Length, and Vertical Location on LowSpeed Static Longitudinal Stability and Damping in Pitch of a Model 

Having 45° Sweptback Wing and Tail Surfaces,” NACA Report 1096 

(1952).

Robert H. Liebeck, “Design of the Blended Wing Body Subsonic 

Transport,” Journal of Aircraft, vol. 41, no. 1 (Jan.–Feb. 2004). pp. 

10–25.

Robert H. Liebeck, Mark A. Page, and Blaine K. Rawdon, “Blended-WingBody Subsonic Commercial Transport,” AIAA Paper 98-0438 (1998).

Robert H. Liebeck, Mark A. Page, Blaine K. Rawdon, Paul W. Scott, and 

Robert A. Wright, “Concepts for Advanced Subsonic Transports,” 

NASA CR-4624 (1994).

Lindsay J. Lina, Garland J. Morris, and Robert A. Champine, “Flight 

Investigation of Factors Affecting the Choice of Minimum Approach 

Speed for Carrier-Type Landings of a Swept-Wing Jet Fighter 

Airplane,” NACA RM-L57F13 (1957).

Alexander Lippisch, “Recent Tests of Tailless Airplanes,” NACA TM-564 

(1930), a NACA translation of his article “Les nouveaux essays 

d’avions sans queue,” l’Aérophile (Feb. 1–15, 1930).

J. Calvin Lovell and Herbert A. Wilson, Jr., “Langley Full-Scale-Tunnel 

Investigation of Maximum Lift and Stability Characteristics of 

an Airplane Having Approximately Triangular Plan Form (DM-1 

Glider),” NACA RM-L7F16 (1947).

Norman M. McFadden and Donovan R. Heinle, “Flight Investigation of 

the Effects of Horizontal-Tail Height, Moment of Inertia, and Control 

Effectiveness on the Pitch-up Characteristics of a 35° Swept-Wing 

Fighter Airplane at High Subsonic Speeds,” NACA RM-A54F21 (1955).

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NASA’s Contributions to Aeronautics

80

1 John McMasters and D. Muncy, “The Early Development of Jet Propelled 

Aircraft,” AIAA Paper 2007-0151, Pts. 1–2 (2007).

Odell A. Morris and A. Warner Robins, “Aerodynamic Characteristics at 

Mach Number 2.01 of an Airplane Configuration Having a Cambered 

and Twisted Arrow Wing Designed for a Mach Number of 3.0,” NASA 

TM-X-115 (1959).

Odell A. Morris and James C. Patterson, Jr., “Transonic Aerodynamic 

Characteristics of Supersonic Transport Model With a Fixed, Warped 

Wing Having 74° Sweep,” NASA TM-X-1167 (1965).

 Odell A. Morris and Roger H. Fournier, “Aerodynamic Characteristics 

at Mach Numbers 2.30, 2.60, and 2.96 of a Supersonic Transport 

Model Having Fixed, Warped Wing,” NASA TM-X-1115 (1965).

Odell A. Morris, Dennis E. Fuller, and Carolyn B. Watson, “Aerodynamic 

Characteristics of a Fixed Arrow-Wing Supersonic Cruise Aircraft 

at Mach Numbers of 2.30, 2.70, and 2.95,” NASA TM-78706 (1978).

Robert G. Mungall, “Flight Investigation of a Combined Geared 

Unbalancing-Tab and Servotab Control System as Used with an 

All-Movable Horizontal Tail,” NACA TN-1763 (1948).

Max M. Munk, “The Aerodynamic Forces on Airship Hulls,” NACA Report 

No. 184 (1923).

Max M. Munk, “Note on the Relative Effect of the Dihedral and the 

Sweep Back of Airplane Wings,” NACA TN-177 (1924).

Andrew Nahum, “The Royal Aircraft Establishment from 1945 to 

Concorde,” in Robert Bud and Philip Gummett, eds., Cold War, Hot 

Science: Applied Research in Britain’s Defence Laboratories, 1945–

1990 (London: Science Museum, 1999), pp. 29–58.

Andrew Nahum, “I Believe the Americans Have Not Yet Taken Them All!” 

in Helmuth Trischler, Stefan Zeilinger, Robert Bud, and Bernard Finn, 

eds., Tackling Transport (London: Science Museum, 2003), pp. 99–138.
Case 1 | Sweep and Swing: Reshaping the Wing for the Jet and Rocket Age

81

NASA Langley Research Center staff, “The Supersonic Transport—A 

Technical Summary,” NASA TN-D-423 (1960).

NASA Langley Research Center staff, “Summary of NACA/NASA VariableSweep Research and Development Leading to the F-111 (TFX),” 

Langley Working Paper LWP-285 (Dec. 22, 1966).

Roy J. Niewald and Jack D. Brewer, “Bibliography of NACA Reports 

Related to Aircraft Control and Guidance Systems, January 1949–

April 1954,” NACA RM-54F01 (1954).

James H. Parks, “Experimental Evidence of Sustained Coupled Longitudinal 

and Lateral Oscillations from a Rocket-Propelled Model of a 35° Swept 

Wing Airplane Configuration,” NACA RM-L54D15 (1954).

Paul Pellicano, Joseph Krumenacker, and David Van Hoy, “X-29 High 

Angle-of-Attack Flight Test Procedures, Results, and Lessons 

Learned,” Society of Flight Test Engineers 21st Annual Symposium 

(1990).

Alfred D. Phillips and Lt. Col. Frank K. Everest, USAF, “Phase II Flight 

Test of the North American YF-100 Airplane USAF No. 52-5754,” 

AFFTC TR-53-33 (1953).

William H. Phillips, “Effect of Steady Rolling on Longitudinal and 

Directional Stability,” NACA TN-1627 (1948).

William H. Phillips, “Appreciation and Prediction of Flying Qualities,” 

NACA Report No. 927 (1949).

William H. Phillips, “Theoretical Analysis of Some Simple Types of 

Acceleration Restrictors,” NACA TN-2574 (1951).

M.J. Queijo, Byron M. Jaquet, and Walter D. Wolmart, “Wind-Tunnel 

Investigation at Low Speed of the Effects of Chordwise Wing Fences 

and Horizontal-Tail Position on the Static Longitudinal Stability 

Characteristics of an Airplane Model with a 35° Sweptback Wing,” 

NACA Report 1203 (1954).

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NASA’s Contributions to Aeronautics

82

1 George A. Rathert, Jr., L. Stewart Rolls, Lee Winograd, and George E. 

Cooper, “Preliminary Flight Investigation of the Wing-Dropping 

Tendency and Lateral-Control Characteristics of a 35° Swept-Wing 

Airplane at Transonic Mach Numbers,” NACA RM-A50H03 (1950).

George A. Rathert, Jr., Howard L. Ziff, and George E. Cooper, 

“Preliminary Flight Investigation of the Maneuvering Accelerations 

and Buffet Boundary of a 35° Swept-Wing Airplane at High Altitude 

and Transonic Speeds,” NACA RM-A50L04 (1951).

A. Warner Robins, Odell A. Morris, and Roy V. Harris, Jr., “Recent 

Research Results in the Aerodynamics of Supersonic Vehicles,” 

AIAA Paper 65-717 (1965).

Dino Roman, J.B. Allen, and Robert H. Liebeck, “Aerodynamic Design 

Challenges of the Blended-Wing-Body Subsonic Transport,” AIAA 

Paper 2000-4335 (2000).

Melvin Sadoff and Thomas R. Sisk, “Longitudinal-Stability 

Characteristics of the Northrop X-4 Airplane (USAF No. 46-677),” 

NACA RM-A50D27 (1950).

Melvin Sadoff, John D. Stewart, and George E. Cooper, “Analytical Study 

of the Comparative Pitch-Up Behavior of Several Airplanes and 

Correlation with Pilot Opinion,” NACA RM-A57D04 (1957).

Edwin J. Saltzman, Donald R. Bellman, and Norman T. Musialowski, 

“Flight-Determined Transonic Lift and Drag Characteristics of the 

YF-102 Airplane With Two Wing Configurations,” NACA RM-H56E08 

(1956).

Thomas R. Sisk and Duane O. Muhleman, “Longitudinal Stability 

Characteristics in Maneuvering Flight of the Convair XF-92A 

Delta-Wing Airplane Including the Effects of Wing Fences,” NACA 

RM-H54J27 (1955).

Thomas R. Sisk and William H. Andrews, “Flight Experience with a 

Delta-Wing Airplane Having Violent Lateral-Longitudinal Coupling 

in Aileron Rolls,” NACA RM-H55H03 (1955).
Case 1 | Sweep and Swing: Reshaping the Wing for the Jet and Rocket Age

83

S.A. Sjoberg and J.P. Reeder: “Flight Measurements of the Lateral and 

Directional Stability and Control Characteristics of an Airplane 

Having a 35° Sweptback Wing with 40-Percent-Span slots and a 

Comparison with Wind-Tunnel Data,” NACA TN-1511 (1948).

S.A. Sjoberg and J.P. Reeder, “Flight Measurements of the Longitudinal 

Stability, Stalling, and Lift Characteristics of an Airplane Having a 

35° Sweptback Wing Without Slots and With 40-Percent-Span Slots 

and a Comparison with Wind-Tunnel Data,” NACA TN-1679 (1948).

S.A. Sjoberg and J.P. Reeder, “Flight Measurements of the Stability, 

Control, and Stalling Characteristics of an Airplane Having a 35° 

Sweptback Wing Without Slots and With 80-Percent-Span Slots 

and a Comparison with Wind-Tunnel Data,” NACA TN-1743 (1948).

S.A. Sjoberg and R.A. Champine, “Preliminary Flight Measurements of 

the Static Longitudinal Stability and Stalling Characteristics of the 

Douglas D-558-II Research Airplane (BuAero No. 37974),” NACA 

RM-L9H31a (1949).

Ronald Smelt, “A Critical Review of German Research on High-Speed 

Airflow,” Journal of the Royal Aeronautical Society, vol. 50, no. 432 

(Dec. 1946).

John E. Steiner, “Transcontinental Rapid Transit: The 367-80 and a 

Transport Revolution—The 1953–1978 Quarter Century,” AIAA Paper 

78-3009 (1978).

Ronald Bel Stiffler, The Bell X-2 Rocket Research Aircraft: The Flight Test 

Program (Edwards AFB: Air Force Flight Test Center, 1957).

W.H. Stillwell, J.V. Wilmerding, and R.A. Champine, “Flight 

Measurements with the Douglas D-558-II (BuAero No. 37974) 

Research Airplane Low-Speed Stalling and Lift Characteristics,” 

NACA RM-L50G10 (1950).

H.S. Tsien, “Supersonic Flow Over an Inclined Body of Revolution,” 

Journal of the Aeronautical Sciences, vol. 5, no. 2 (Oct. 1938).

1
NASA’s Contributions to Aeronautics

84

1 United States Senate, TFX Contract Investigation (Second Series): 

Hearings Before the Permanent Subcommittee on Investigations 

of the Committee on Government Operations, United States Senate, 

91st Congress, 2nd Session, Pts. 1–3 (Washington: GPO, 1970).

Walter G. Vincenti, “Robert Thomas Jones,” in Biographical Memoirs, 

vol. 86 (Washington: National Academy of Sciences, 2005).

Marion H. Yancey, Jr., and Maj. Stuart R. Childs, USAF, “Phase IV 

Stability Tests of the F-100A Aircraft, USAF S/N 52-5767,” AFFTC 

TR-55-9 (1955).

Sean Wakayama, “Multidisciplinary Design Optimization of the BlendedWing-Body,” AIAA Paper 98-4938 (1998).

Joseph Weil, Paul Comisarow, and Kenneth W. Goodson, “Longitudinal 

Stability and Control Characteristics of an Airplane Model Having 

a 42.8° Sweptback Circular-Arc Wing with Aspect Ratio 4.00, Taper 

Ratio 0.60, and Sweptback Tail Surfaces,” NACA RM-L7G28 (1947).

A. Miles Whitnah and Ernest R. Hillje, “Space Shuttle Wind Tunnel 

Testing Summary,” NASA Reference Publication 1125 (1984).

Richard T. Whitcomb, “An Investigation of the Effects of Sweep on the 

Characteristics of a High-Aspect-Ratio Wing in the Langley 8-Ft. 

High Speed Tunnel,” NACA RM-L6J01a (1947).

Edward F. Whittle, Jr., and J. Calvin Lovell, “Full-Scale Investigation of 

an Equilateral Triangular Wing Having 10-Percent-Thick Biconvex 

Airfoil Sections,” NACA RM-L8G05 (1948).

W.C. Williams and A.S. Crossfield, “Handling Qualities of High-Speed 

Airplanes,” NACA RM-L52A08 (1952).

Herbert A. Wilson, Jr., and J. Calvin Lovell, “Full Scale Investigation of 

the Maximum Lift and Flow Characteristics of an Airplane Having 

Approximately Triangular Plan Form,” NACA RM-L6K20 (1947). 
Case 1 | Sweep and Swing: Reshaping the Wing for the Jet and Rocket Age

85

Herbert A. Wilson, Jr., and J. Calvin Lovell, “Langley Full-Scale Tunnel 

Investigation of Maximum Lift and Stability Characteristics of an 

Airplane Having Approximately Triangular Plan Form (DM-1 Glider), 

NACA RM-L7F16 (1947).

Books and Monographs:

David A. Anderton, Sixty Years of Aeronautical Research, 1917–1977, 

EP-145 (Washington: NASA, 1978).

Roland Beamont, Testing Early Jets: Compressibility and the Supersonic 

Era (Shrewsbury: Airlife, 1990).

Walter J. Boyne and Donald H. Lopez, eds., The Jet Age: Forty Years of 

Jet Aviation (Washington: Smithsonian Institution Press, 1979).

Peter W. Brooks, The Modern Airliner: Its Origins and Development 

(London: Putnam & Co., Ltd., 1961).

Charles Burnet, Three Centuries to Concorde (London: Mechanical 

Engineering Publications Ltd., 1979).

Joseph R. Chambers, Partners in Freedom: Contributions of the Langley 

Research Center to U.S. Military Aircraft of the 1990s, SP-2000-4519 

(Washington: NASA, 2000).

Joseph R. Chambers, Innovation in Flight: Research of the NASA Langley 

Research Center on Revolutionary Advanced Concepts for Aeronautics, 

SP-2005-4539 (Washington: NASA, 2005).

Eric M. Conway, High-Speed Dreams: NASA and the Technopolitics of 

Supersonic Transportation, 1945–1999 (Baltimore: The Johns Hopkins 

Press, 2005).

William H. Cook, The Road to the 707: The Inside Story of Designing the 

707 (Bellevue, WA: TYC Publishing Co., 1991).

A. Scott Crossfield with Clay Blair, Always Another Dawn: The Story of a 

Rocket Test Pilot (Cleveland: World Publishing Co., 1960).

1
NASA’s Contributions to Aeronautics

86

1 Richard E. Day, Coupling Dynamics in Aircraft: A Historical Perspective,

SP-532 (Washington: NASA, 1997).

Michael Eckert, The Dawn of Fluid Dynamics: A Discipline Between 

Science and Technology (Weinheim: Wiley-VCH Verlag, 2006).

Brig. Gen. Frank K. Everest, Jr., with John Guenther, The Fastest Man 

Alive (New York: Bantam, 1990 ed.).

George W. Gray, Frontiers of Flight: The Story of NACA Research (New 

York: Knopf, 1948).

Roy A. Grossnick, et al., United States Naval Aviation 1910–1995 

(Washington: U.S. Navy, 1997).

Richard P. Hallion, Supersonic Flight: Breaking the Sound Barrier and 

Beyond—The Story of the Bell X-1 and Douglas D-558 (New York: The 

Macmillan Co. in association with the Smithsonian Institution, 1972).

Richard P. Hallion, ed., The Hypersonic Revolution: Case Studies in the 

History of Hypersonic Technology, vols. 1–2 (Washington: USAF, 1998).

Richard P. Hallion and Michael H. Gorn, On the Frontier: Experimental 

Flight at NASA Dryden (Washington: Smithsonian Books, 2002).

James R. Hansen, Engineer in Charge: A History of the Langley Aeronautical 

Laboratory, 1917–1958, SP-4305 (Washington: NASA, 1987).

Edwin P. Hartman, Adventures in Research: A History of the Ames Research 

Center, 1940–1965, SP-4302 (Washington: NASA 1970).

Theodore von Kármán, Aerodynamics (New York: McGraw-Hill Book 

Company, Inc., 1963 ed.).

Theodore von Kármán and Lee Edson, The Wind and Beyond: Theodore 

von Kármán, Pioneer in Aviation and Pathfinder in Space (Boston: 

Little, Brown and Co., 1967).
Case 1 | Sweep and Swing: Reshaping the Wing for the Jet and Rocket Age

87

Thomas A. Marschak, The Role of Project Histories in the Study of R&D, 

Rand Report P-2850 (Santa Monica: The Rand Corporation, 1965).

Hans-Ulrich Meier, ed., Die Pfeilflügelentwicklung in Deutschland bis 

1945 (Bonn: Bernard & Graefe Verlag, 2006).

Jay Miller, The X-Planes: X-1 to X-45 (Hinckley, UK: Midland Publishing, 

2001).

Kenneth Owen, Concorde: Story of a Supersonic Pioneer (London: Science 

Museum, 2001).

Robert L. Perry, Innovation and Military Requirements: A Comparative 

Study, Rand Report RM-5182PR (Santa Monica: The Rand 

Corporation, 1967).

W. Hewitt Phillips, Journey in Aeronautical Research: A Career at NASA 

Langley Research Center, No. 12 in Monographs in Aerospace History

(Washington: NASA, 1998).

R. Dale Reed with Darlene Lister, Wingless Flight: The Lifting Body Story, 

SP-4220 (Washington: NASA 1997).

Joseph A. Shortal, A New Dimension: Wallops Island Flight Test Range: 

The First Fifteen Years, RP-1028 (Washington: NASA, 1978).

Milton O. Thompson and Curtis Peebles, Flying Without Wings: NASA 

Lifting Bodies and The Birth of the Space Shuttle (Washington: 

Smithsonian Institution Press, 1999).

Milton O. Thompson with J.D. Hunley, Flight Research: Problems 

Encountered and What They Should Teach Us, SP-2000-4522 

(Washington: NASA, 2000).

Helmuth Trischler, Stefan Zeilinger, Robert Bud, and Bernard Finn, eds., 

Tackling Transport (London: Science Museum, 2003).

1
88

Whitcomb evaluates the shape of one of his area rule models in the 8-foot High Speed Tunnel. 

NASA.
89

2 Richard Whitcomb 

and the Quest for 

Aerodynamic Efficiency

Jeremy Kinney

CASE

2

Much of the history of aircraft design in the postwar era is encapsulated 

by the remarkable work of NACA–NASA engineer Richard T. Whitcomb. 

Whitcomb, a transonic and supersonic pioneer, gave to aeronautics the 

wasp-waisted area ruled transonic airplane, the graceful and highly 

efficient supercritical wing, and the distinctive wingtip winglet. But he 

also contributed greatly to the development of advanced wind tunnel 

design and testing. His life offers insights into the process of aeronautical 

creativity and the role of the genius figure in advancing flight.

O N DECEMBER 21, 1954, Convair test pilot Richard L. “Dick” 

Johnson flew the YF-102A Delta Dagger prototype to Mach 1, 

an achievement that marked the meeting of a challenge that had 

been facing the American aeronautical community. The Delta Dagger’s 

contoured fuselage, shaped by a new design concept, the area rule, 

enabled an efficient transition from subsonic to supersonic via the transonic regime. Seventeen years later, test pilot Thomas C. “Tom” McMurtry 

made the first flight in the F-8 Supercritical Wing flight research vehicle on March 9, 1971. The flying testbed featured a new wing designed 

to cruise at near-supersonic speeds for improved fuel economy. Another 

17 years later, the Boeing Company announced the successful maiden 

flight of what would be the manufacturer’s best-selling airliner, the 747-

400, on April 29, 1988. Incorporated into the design of the jumbo jet 

were winglets: small vertical surfaces that reduced drag by smoothing 

turbulent airflow at the wingtips to increase fuel efficiency.1 All three of 

these revolutionary innovations originated with one person, Richard T. 

1. David A. Anderton, “NACA Formula Eases Supersonic Flight,” Aviation Week vol. 63 (Sept. 12, 

1955): p. 15; Marvin Miles, “New Fighter Jet Gets Test,” Los Angeles Times, Mar. 10, 1971, p. 

26; “Boeing’s 747-400 Jet Makes Maiden Flight,” Wall Street Journal, May 2, 1988, p. 8.
NASA’s Contributions to Aeronautics

90

2 “Dick” Whitcomb, an aeronautical engineer working for the National 

Advisory Committee for Aeronautics (NACA) and its successor, the 

National Aeronautics and Space Administration (NASA).

A major aeronautical revolution was shaping the direction and use 

of the airplane during the latter half of the 20th century. The invention 

of the turbojet engine in Europe and its incorporation into the airplane 

transformed aviation. The aeronautical community followed a basic 

premise—to make the airplane fly higher, faster, farther, and cheaper 

than ever before—as national, military, industrial, and economic factors shaped requirements. As a researcher at the Langley Memorial 

Aeronautical Laboratory in Hampton, VA, Dick Whitcomb was part 

of this movement, which was central to the missions of both the NACA 

and NASA.2 His three fundamental contributions, the area rule fuselage, 

the supercritical wing, and the winglet, each in their own aerodynamic 

ways offered an increase in speed and performance without an increase 

in power. Whitcomb was highly individualistic, visionary, creative, and 

practical, and his personality, engineering style, and the working environment nurtured at Langley facilitated his quest for aerodynamic efficiency.

The Making of an Engineer

Richard Travis Whitcomb was born on February 21, 1921, in Evanston, 

IL, and grew up in Worcester, MA. He was the eldest of four children in a 

family led by mathematician-engineer Kenneth F. Whitcomb.3 Whitcomb 

was one of the many air-minded American children building and testing 

aircraft models throughout the 1920s and 1930s.4 At the age of 12, he 

created an aeronautical laboratory in his family’s basement. Whitcomb 

spent the majority of his time there building, flying, and innovating rubberband-powered model airplanes, with the exception of reluctantly 

eating, sleeping, and going to school. He never had a desire to fly himself, but, in his words, he pursued aeronautics for the “fascination of 

2. Whitcomb’s story has been interpreted from the viewpoint of the NACA and NASA’s overall 

contributions to aeronautics by several historians and engineers. This chapter depends heavily on the 

work of James Hansen, Richard Hallion, Michael Gorn, Lane Wallace, John Becker, Donald Baals, 

and William Corliss.

3. Clay Blair, Jr., “The Man Who Put the Squeeze on Aircraft Design,” Air Force Magazine, vol. 39 

(Jan. 1956): p. 50.

4. “Richard Travis Whitcomb: Distinguished Research Associate,” NASA Langley Research Center, 

Apr. 1983, File CW-463000-01, National Air and Space Museum Archives.
Case 2 | Richard Whitcomb and the Quest for Aerodynamic Efficiency

91

2 making a model that would fly.” One innovation Whitcomb developed 

was a propeller that folded back when it stopped spinning to reduce 

aerodynamic drag. He won several model airplane contests and was a 

prizewinner in the Fisher Body Company automobile model competition; both were formative events for young American men who would 

become the aeronautical engineers of the 1940s. Even as a young man, 

Whitcomb exhibited an enthusiastic drive that could not be diverted 

until the challenge was overcome.5

A major influence on Whitcomb during his early years was his paternal grandfather, who had left farming in Illinois to become a manufacturer of mechanical vending machines. Independent and driven, the 

grandfather was also an acquaintance of Thomas A. Edison. Whitcomb 

listened attentively to his grandfather’s stories about Edison and soon 

came to idolize the inventor for his ideas as well as for his freethinking 

individuality.6 The admiration for his grandfather and for Edison shaped 

Whitcomb’s approach to aeronautical engineering.

Whitcomb received a scholarship to nearby Worcester Polytechnic 

Institute and entered the prestigious school’s engineering program in 

1939. He lived at home to save money and spent the majority of his 

time in the institute’s wind tunnel. Interested in helping with the war 

effort, Whitcomb’s senior project was the design of a guided bomb. 

He graduated with distinction with a bachelor’s of science degree in 

mechanical engineering. A 1943 Fortune magazine article on the NACA 

convinced Whitcomb to join the Government-civilian research facility 

at Hampton, VA.7

Airplanes ventured into a new aerodynamic regime, the so-called 

“transonic barrier,” as Whitcomb entered into his second year at 

Worcester. At speeds approaching Mach 1, aircraft experienced sudden 

changes in stability and control, extreme buffeting, and, most importantly, a dramatic increase in drag, which exposed three challenges to 

5. Richard Witkin, “Air Scientist Got His Start When 12,” New York Times, Oct. 3, 1955, p. 20 

(quote); Ray Bert, “Winged Victory: Meet Richard Whitcomb,” Transformations (fall 2002), 

http://www.wpi.edu/News/Transformations/2002Fall/whitcomb.html (Accessed Feb. 14, 2009); 

“Jet Pioneers—Richard T. Whitcomb,” n.d., File CW-463000-01, National Air and Space Museum 

Archives.

6. Barbara Rowes, “When You Ride Tomorrow’s Airplanes, You’ll Thank Dick Whitcomb,” Washington Post-Times Herald, Aug. 31, 1969, p. 165.

7. Bert, “Winged Victory”; Witkin, “Air Scientist Got His Start When 12”; Brian Welch, “Whitcomb: Aeronautical Research and the Better Shape,” Langley Researcher (Mar. 21, 1980): p. 4.
NASA’s Contributions to Aeronautics

92

2 the aeronautical community, involving propulsion, research facilities, and aerodynamics. The first challenge involved the propeller and 

piston-engine propulsion system. The highly developed and reliable system was at a plateau and incapable of powering the airplane in the transonic regime. The turbojet revolution brought forth by the introduction 

of jet engines in Great Britain and Germany in the early 1940s provided 

the power needed for transonic flight. The latter two challenges directly 

involved the NACA and, to an extent, Dick Whitcomb, during the course 

of the 1940s. Bridging the gap between subsonic and supersonic speeds 

was a major aerodynamic challenge.8

Little was known about the transonic regime, which falls between 

Mach 0.8 and 1.2. Aeronautical engineers faced a daunting challenge 

rooted in developing new tools and concepts. The aerodynamicist’s primary tool, the wind tunnel, was unable to operate and generate data at 

transonic speeds. Four approaches were used in lieu of an available wind 

tunnel in the 1940s for transonic research. One way to generate data for 

speeds beyond 350 mph was through aircraft diving at terminal velocity, 

which was dangerous for test pilots and of limited value for aeronautical engineers. Moreover, a representative drag-weight ratio for a 1940-

era airplane ensured that it was unable to exceed Mach 0.8. Another 

way was the use of a falling body, an instrumented missile dropped from 

the bomb bay of a Boeing B-29 Superfortress. A third method was the 

wing-flow model. NACA personnel mounted a small, instrumented airfoil on top of the wing of a North American P-51 Mustang fighter. The 

Mustang traveled at high subsonic speeds and provided a recoverable 

method in real-time conditions. Finally, the NACA launched small models mounted atop rockets from the Wallops Island facility on Virginia’s 

Eastern Shore.9 The disadvantages for these three methods were that 

they only generated data for short periods of time and that there were 

many variables regarding conditions that could affect the tests.

Even if a wind tunnel existed that was capable of evaluating aircraft 

at transonic speeds, there was no concept that guaranteed a successful 

8. John Becker, The High Speed Frontier: Case Histories of Four NACA Programs 1920–1950, 

NASA SP-445 (Washington, DC: U.S. Government Printing Office, 1980), p. 61.

9. Becker, High Speed Frontier, p. 61; Lane E. Wallace, “The Whitcomb Area Rule: NACA Aerodynamics Research and Innovation,” in Pam E. Mack, ed., From Engineering Science to Big Science: The NACA and NASA Collier Trophy Research Project Winners, (Washington, DC: National 

Aeronautics and Space Administration, 1998), p. 137.
Case 2 | Richard Whitcomb and the Quest for Aerodynamic Efficiency

93

2 transonic aircraft design. A growing base of knowledge in supersonic 

aircraft design emerged in Europe beginning in the 1930s. Jakob Ackeret 

operated the first wind tunnel capable of generating Mach 2 in Zurich, 

Switzerland, and designed tunnels for other countries. The international 

high-speed aerodynamics community met at the Volta Conference held 

in Rome in 1935. A paper presented by German aerodynamicist Adolf 

Busemann argued that if aircraft designers swept the wing back from the 

fuselage, it would offset the increase in drag beyond speeds of Mach 1. 

Busemann offered a revolutionary answer to the problem of high-speed 

aerodynamics and the sound barrier. In retrospect, the Volta Conference 

proved to be a turning point in high-speed aerodynamics research, especially for Nazi Germany. In 1944, Dietrich Küchemann discovered that a 

contoured fuselage resembling the now-iconic Coca-Cola soft drink bottle was ideal when combined with Busemann’s swept wings. American 

researcher Robert T. Jones independently discovered the swept wing at 

NACA Langley almost a decade after the Volta Conference. Jones was 

a respected Langley aerodynamicist, and his five-page 1945 report provided a standard definition of the aerodynamics of a swept wing. The 

report appeared at the same time that high-speed aerodynamic information from Nazi Germany was reaching the United States.10

As the German and American high-speed traditions merged after 

World War II, the American aeronautical community realized that there 

were still many questions to be answered regarding high-speed flight. 

Three NACA programs in the late 1940s and early 1950s overcame the 

remaining aerodynamic and facility “barriers” in what John Becker characterized as “one of the most effective team efforts in the annals of aeronautics.” The National Aeronautics Association recognized these NACA 

achievements three times through aviation’s highest award, the Collier 

Trophy, for 1947, 1951, and 1954. The first award, for the achievement 

of supersonic flight by the X-1, was presented jointly to John Stack of 

the NACA, manufacturer Lawrence D. Bell, and Air Force test pilot Capt. 

Charles E. “Chuck” Yeager. The second award in 1952 recognized the 

slotted transonic tunnel development pioneered by John Stack and his 

associates at NACA Langley.11 The third award recognized the direct 

byproduct of the development of a wind tunnel in which the visionary 

10. John D. Anderson, Jr., A History of Aerodynamics and its Impact on Flying Machines (New 

York: Cambridge University Press, 1997), pp. 419, 424–425.

11. Becker, High Speed Frontier, p. 61.
NASA’s Contributions to Aeronautics

94

2 mind of Dick Whitcomb developed the design concept that would enable 

aircraft to efficiently transition from subsonic to supersonic speeds 

through the transonic regime.

Dick Whitcomb and the Transonic-Supersonic Breakthrough

Whitcomb joined the research community at Langley in 1943 as a member of Stack’s Transonic Aerodynamics Branch working in the 8-foot 

High-Speed Tunnel (HST). Initially, NACA managers placed him in the 

Flight Instrument Research Division, but Whitcomb’s force of personality ensured that he would be working directly on problems related to 

aircraft design. As many of his colleagues and historians would attest, 

Whitcomb quickly became known for an analytical ability rooted in 

mathematics, instinct, and aesthetics.12

In 1945, Langley increased the power of the 8-foot HST to generate Mach 0.95 speeds, and Whitcomb was becoming increasingly familiar with transonic aerodynamics, which helped him in his developing 

investigation into the design of supersonic aircraft. The onset of drag 

created by shock waves at transonic speeds was the primary challenge. 

John Stack, Ezra Kotcher, and Lawrence D. Bell proved that breaking 

the sound barrier was possible when Chuck Yeager flew the Bell X-1 to 

Mach 1.06 (700 mph) on October 14, 1947. Designed in the style of a .50- 

caliber bullet with straight wings, the Bell X-1 was a successful supersonic airplane, but it was a rocket-powered research airplane designed 

specifically for and limited to that purpose. The X-1 would not offer 

designers the shape of future supersonic airplanes. Operational turbojetpowered aircraft designed for military missions were much heavier and 

would use up much of their fuel gradually accelerating toward Mach 1 

to lessen transonic drag.13 The key was to get operation aircraft through 

the transonic regime, which ranged from Mach 0.9 to Mach 1.1.

A very small body of transonic research existed when Whitcomb 

undertook his investigation. British researchers W.T. Lord of the 

Royal Aeronautical Establishment and G.N. Ward of the University 

of Manchester and American Wallace D. Hayes attempted to solve the 

problem of transonic drag through mathematical analyses shortly after 

World War II in 1946. These studies generated mathematical symbols 

12. James R. Hansen, Engineer in Charge: A History of the Langley Aeronautical Laboratory, 

1917–1958, NASA SP-4305 (Washington, DC: NASA, 1987), pp. 331–332.

13. Ibid., p. 332.
Case 2 | Richard Whitcomb and the Quest for Aerodynamic Efficiency

95

2 that did not lend themselves to the design and shape of transonic and 

supersonic aircraft.14

Whitcomb’s analysis of available data generated by the NACA in 

ground and free-flight tests led him to submit a proposal for testing 

swept wing and fuselage combinations in the 8-foot HST in July 1948. 

There had been some success in delaying transonic drag by addressing 

the relationship between wing sweep and fuselage shape. Whitcomb 

believed that careful attention to arrangement and shape of the wing 

and fuselage would result in their counteracting each other. His goal was 

to reach a milestone in supersonic aircraft design. The tests, conducted 

from late 1949 to early 1950, revealed no significant decrease in drag at 

high subsonic (Mach 0.95) and low supersonic (Mach 1.2) speeds. The 

wing-fuselage combinations actually generated higher drag than their 

individual values combined. Whitcomb was at an impasse and realized 

he needed to refocus on learning more about the fundamental nature 

of transonic airflow.15

Just before Whitcomb had submitted his proposal for his wind tunnel tests, John Stack ordered the conversion of the 8-foot HST in the 

spring of 1948 to a slotted throat to enable research in the transonic 

regime. In theory, slots in the tunnel’s test section, or throat, would enable 

smooth operation at very high subsonic speeds and at low supersonic 

speeds. The initial conversion was not satisfactory because of uneven 

flow. Whitcomb and his colleagues, physicist Ray Wright and engineer 

Virgil S. Ritchie, hand-shaped the slots based on their visualization of 

smooth transonic flow. They also worked directly with Langley woodworkers to design and fabricate a channel at the downstream end of 

the test section that reintroduced air that traveled through the slots. 

Their painstaking work led to the inauguration of transonic operations 

within the 8-foot HST 7 months later, on October 6, 1950.16 Whitcomb, 

14. Hansen, Engineer in Charge, p. 341. As James R. Hansen has suggested, these were certainly 

antecedents to Whitcomb’s area rule, but it was his highly intuitive visual mind that resulted in 

something original.

15. Ibid., p. 332.

16. The NACA referred to the facility as the 8-foot Transonic Tunnel after Oct. 1950, but for the 

purposes of clarity and to avoid confusion with the follow-on 8-foot Transonic Pressure Tunnel, the 

original designation 8-foot High Speed Tunnel is used in this text. Hansen, Engineer in Charge, pp. 

327–328, 454; Steven T. Corneliussen, “The Transonic Wind Tunnel and the NACA Technical 

Culture,” in Pam E. Mack, ed., From Engineering Science to Big Science: The NACA and NASA 

Collier Trophy Research Project Winners (Washington, DC: NASA, 1998), p. 133.
NASA’s Contributions to Aeronautics

96

2

The slotted-throat test section of the 8-foot High-Speed Tunnel. NASA.

as a young engineer, was helping to refine a tunnel configuration that 

was going to allow him to realize his potential as a visionary experimental aeronautical engineer.

The NACA distributed a confidential report on the new tunnel during 

the fall of 1948, which was distributed to the military services and select 

manufacturers. By the following spring, rumors had been circulating 

about the new tunnel throughout the industry. Initially, the call for secrecy 

evolved into outright public acknowledgement of the NACA’s new transonic tunnels (including the 16-foot HST) with the awarding of the 1951 

Collier Trophy to John Stack and 19 of his associates at Langley for the 

slotted wall. The Collier Trophy specifically recognized the importance of 

a research tool, which was a first in the 40-year history of the award. The 

NACA claimed that its slotted-throat transonic tunnels gave the United 

States a 2-year lead in the design of supersonic military aircraft.17

17. Hansen, Engineer in Charge, pp. 329, 330–331.
Case 2 | Richard Whitcomb and the Quest for Aerodynamic Efficiency

97

2 With the availability of the 8-foot HST and its slotted throat, the combined use of previously available wind tunnel components—the tunnel balance, pressure orifice, tuft surveys, and schlieren photographs—resulted 

in a new theoretical understanding of transonic drag. The schlieren photographs revealed three shock waves at transonic speeds. One was the familiar shock wave that formed at the nose of an aircraft as it pushed forward 

through the air. The other two were, according to Whitcomb, “fascinating 

new types” of shock waves never before observed, in which the fuselage and 

wings met and at the trailing edge of the wing. These shocks contributed 

to a new understanding that transonic drag was much larger in proportion 

to the size of the fuselage and wing than previously believed. Whitcomb 

speculated that these new shock waves were the cause of transonic drag.18

The Path to Area Rule

Conventional high-speed aircraft design emulated Ernst Mach’s finding 

that bullet shapes produced less drag. Aircraft designers started with a 

pointed nose and gradually thickened the fuselage to increase its crosssectional area, added wings and a tail, and then decreased the diameter of the fuselage. The rule of thumb for an ideal streamlined body 

for supersonic flight was a function of the diameter of the fuselage. 

Understanding the incorporation of the wing and tail, which were added 

for practical purposes because airplanes need them to fly, into Mach’s 

ideal high-speed soon became the focus of Whitcomb’s investigation.19

The 8-foot HST team at Langley began a series of tests on various 

wing and body combinations in November 1951. The wind tunnel models featured swept, straight, and delta wings, and fuselages with varying 

amounts of curvature. The objective was to evaluate the amount of drag 

generated by the interference of the two shapes at transonic speeds. The 

tests resulted in two important realizations for Whitcomb. First, variations in fuselage shape led to marked changes in wing drag. Second, 

and most importantly, he learned that the combination of fuselage and 

wing drag had to be considered together as a synergistic aerodynamic 

system rather than separately, as they had been before.20

18. Richard T. Whitcomb and Thomas C. Kelly, “A Study of the Flow Over a 45-degree Sweptback 

Wing-Fuselage Combination at Transonic Mach Numbers,” NACA RM-L52DO1 (June 25, 1952), 

p. 1; Hansen, Engineer in Charge, p. 333.

19. Ibid., p. 333.

20. Ibid., p. 334.
NASA’s Contributions to Aeronautics

98

2 While Whitcomb was performing his tests, he took a break to 

attend a Langley technical symposium, where swept wing pioneer Adolf 

Busemann presented a helpful concept for imagining transonic flow. 

Busemann asserted that wind tunnel researchers should emulate aerodynamicists and theoretical scientists in visualizing airflow as analogous 

to plumbing. In Busemann’s mind, an object surrounded by streamlines 

constituted a single stream tube. Visualizing “uniform pipes going over 

the surface of the configuration” assisted wind tunnel researchers in 

determining the nature of transonic flow.21

Whitcomb contemplated his findings in the 8-foot HST and 

Busemann’s analogy during one of his daily thinking sessions in 

December 1951. Since his days at Worcester, he dedicated a specific 

part of his day to thinking. At the core of Whitcomb’s success in solving efficiency problems aerodynamically was the fact that, in the words 

of one NASA historian, he was the kind of “rare genius who can see 

things no one else can.”22 His relied upon his mind’s eye—the nonverbal thinking necessary for engineering—to visualize the aerodynamic 

process, specifically transonic airflow.23 Whitcomb’s ability to apply his 

findings to the design of aircraft was a clear indication that using his 

mind through intuitive reasoning was as much an analytical aerodynamic tool as a research airplane, wind tunnel, or slide rule.

With his feet propped up on his desk in his office a flash of inspiration—a “Eureka” moment, in the mythic tradition of his hero, Edison—

led him to the solution of reducing transonic drag. Whitcomb realized 

that the total cross-sectional area of a fuselage, wing, and tail caused 

transonic drag or, in his words: “transonic drag is a function of the 

longitudinal development of the cross-sectional areas of the entire 

airplane.”24 It was simply not just the result of shock waves forming 

at the nose of the airplane, but drag-inducing shock waves formed 

just behind the wings. Whitcomb visualized in his mind’s eye that if a 

designer narrowed the fuselage or reduced its cross section, where the 

wings attached, and enlarged the fuselage again at the trailing edge, 

21. Ibid., p. 334.

22. Roger D. Launius, quoted in James Schultz, Crafting Flight: Aircraft Pioneers and the Contributions of 

the Men and Women of NASA Langley Research Center (Washington, DC: NASA, 2003), p. 183.

23. Eugene S. Ferguson, Engineering and the Mind’s Eye (Boston: MIT Press, 1994), p. 41; Hansen, Engineer in Charge, p. 328.

24. Whitcomb quoted in Welch, “Whitcomb,” p. 5.
Case 2 | Richard Whitcomb and the Quest for Aerodynamic Efficiency

99

2 then the fuselage would facilitate a smoother transition from subsonic 

to supersonic speeds. Pinching the fuselage to resemble a wasp’s waist 

allowed for smoother flow of the streamlines as they traveled from the 

nose and over the fuselage, wings, and tail. Even though the fuselage 

was shaped differently, the overall cross section was the same along the 

length of the fuselage. Without the pinch, the streamlines would bunch 

and form shock waves, which created the high energy losses that prevented supersonic flight.25 The removal at the wing of those “aerodynamic anchors,” as historians Donald Baals and William Corliss called 

them, and the recognition of the sensitive balance between fuselage and 

wing volume were the key.26

Verification of the new idea involved the comparison of the data 

compiled in the 8-foot HST, all other available NACA-gathered transonic 

data, and Busemann’s plumbing concept. Whitcomb was convinced 

that his area rule made sense of the questions he had been investigating. Interestingly enough, Whitcomb’s colleagues in the 8-foot HST, 

including John Stack, were skeptical of his findings. He presented his 

findings to the Langley community at its in-house technical seminar.27

After Whitcomb’s 20-minute talk, Busemann remarked: “Some people come up with half-baked ideas and call them theories. Whitcomb 

comes up with a brilliant idea and calls it a rule of thumb.”28 The name 

“area rule” came from the combination of “cross-sectional area” with 

“rule of thumb.”29

With Busemann’s endorsement, Whitcomb set out to validate the 

rule through the wind tunnel testing in the 8-foot HST. His models featured fuselages narrowed at the waist. He had enough data by April 1952 

indicating that pinching the fuselage resulted in a significant reduction 

in transonic drag. The resultant research memorandum, “A Study of 

the Zero Lift Drag Characteristics of Wing-Body Combinations near 

25. Richard T. Whitcomb, “A Study of the Zero-Lift Drag-Rise Characteristics of Wing-Body Combinations Near the Speed of Sound,” NACA TR-1273 (1956), pp. 519, 538–539; Engineer in 

Charge, pp. 334–335.

26. Donald D. Baals and William R. Corliss, Wind Tunnels of NASA (Washington, DC: Scientific 

and Technical Information Branch, National Aeronautics and Space Administration, 1981), p. 63.

27. Hansen, Engineer in Charge, p. 336.

28. Quoted in Richard P. Hallion, Designers and Test Pilots (Alexandria, VA: Time-Life Books, 1983), 

p. 143.

29. Michael Gorn, Expanding the Envelope: Flight Research at NACA and NASA (Lexington: 

University Press of Kentucky, 2001), p. 329. 
NASA’s Contributions to Aeronautics

100

2 the Speed of Sound,” appeared the following September. The NACA 

immediately distributed it secretly to industry.30

The area rule provided a transonic solution to aircraft designers in 

four steps. First, the designer plotted the cross sections of the aircraft 

fuselage along its length. Second, a comparison was made between the 

design’s actual area distribution, which reflected outside considerations, 

such as engine diameter and the overall size dictated by an aircraft carrier’s elevator deck, and the ideal area distribution that originated in 

previous NACA mathematical studies. The third step involved the reconciliation of the actual area distribution with the ideal area distribution. 

Once again, practical design considerations shaped this step. Finally, 

the designer converted the new area distribution back into cross sections, which resulted in the narrowed fuselage that took into account 

the overall area of the fuselage and wing combination.31 A designer that 

followed those four steps would produce a successful design with minimum transonic drag.

Validation in Flight

As Whitcomb was discovering the area rule, Convair in San Diego, 

CA, was finalizing its design of a new supersonic all-weather fighterinterceptor, began in 1951, for a substantial Air Force contract. The 

YF-102 Delta Dagger combined Mach’s ideal high-speed bullet-shaped 

fuselage and delta wings pioneered on the Air Force’s Convair XF-92A 

research airplane with the new Pratt & Whitney J57 turbojet, the world’s 

most powerful at 10,000 pounds thrust. Armed entirely with air-to-air 

and forward-firing missiles, the YF-102 was to be the prototype for 

America’s first piloted air defense weapon’s system.32 Convair heard of 

the NACA’s transonic research at Langley and feared that its investment 

in the YF-102 and the payoff with the Air Force would come to naught 

if the new airplane could not fly supersonic.33 Convair’s reputation and 

a considerable Department of Defense contract were at stake.

30. Richard T. Whitcomb, “A Study of the Zero-Lift Drag-Rise Characteristics of Wing-Body Combinations Near the Speed of Sound,” NACA RM-L52H08 (Sept. 3, 1952). RM-L52H08 was superseded by TR-1273 (see note 23) when the document became unclassified in 1956.

31. Anderton, “NACA Formula Eases Supersonic Flight,” pp. 13–14.

32. Gordon Swanborough, United States Military Aircraft Since 1909 (London: Putnam, 1963), 

pp. 151, 153.

33. Hansen, Engineer in Charge, p. 337.
Case 2 | Richard Whitcomb and the Quest for Aerodynamic Efficiency

101

2 A delegation of Convair engineers visited Langley in mid-August 1952, 

where the engineers witnessed a disappointing test of an YF-102 model in 

the 8-foot HST. The data indicated, according to the NACA at least, that 

the YF-102 was unable to reach Mach 1 in level flight. The transonic drag 

exhibited near Mach 1 simply counteracted the ability of the J57 to push 

the YF-102 through the sound barrier. They asked Whitcomb what could be 

done, and he unveiled his new rule of thumb for the design of supersonic 

aircraft. The data, Whitcomb’s solution, and what was perceived as the 

continued skepticism on the part of his boss, John Stack, left the Convair 

engineers unconvinced as they went back to San Diego with their model.34

They did not yet see the area rule as the solution to their perceived problem.

Nevertheless, Whitcomb worked with Convair’s aerodynamicists to 

incorporate the area rule into the YF-102. New wind tunnel evaluations 

in May 1953 revealed a nominal decrease in transonic drag. He traveled 

to San Diego in August to assist Convair in reshaping the YF-102 fuselage. 

The NACA notified Convair that the modified design, soon be designated 

the YF-102A, was capable of supersonic flight in October.35

Despite the fruitful collaboration with Whitcomb, Convair was hedging its bets when it continued the production of the prototype YF-102 in 

the hope that it was a supersonic airplane. The new delta wing fighter with 

a straight fuselage was unable to reach its designed supersonic speeds 

during its full-scale flight evaluation and tests by the Air Force in January 

1954. The disappointing performance of the YF-102 to reach only Mach 

0.98 in level flight confirmed the NACA’s wind tunnel findings and validated 

Whitcomb’s research that led to his area rule. The Air Force realistically 

shifted the focus toward production of the YF-102A after NACA Director 

Hugh Dryden guaranteed that Chief of Staff of the Air Force Gen. Nathan 

F. Twining developed a solution to the problem and that the information 

had been made available to Convair and the rest of the aviation industry. 

The Air Force ordered Convair to stop production of the YF-102 and retool 

to manufacture the improved area rule design.36

It took Convair only 7 months to prepare the prototype YF-102A, thanks 

to the collaboration with Whitcomb. Overall, the new fighter-interceptor 

34. Baals and Corliss, Wind Tunnels of NASA, p. 62; Hansen, Engineer in Charge, p. 337.

35. Hansen, Engineer in Charge, p. 337.

36. Hansen, Engineer in Charge, pp. 337–338; Richard P. Hallion, On the Frontier: Flight

Research at Dryden, 1946–1981 (Washington, DC: NASA, 1984), p. 90; Baals and Corliss, 

Wind Tunnels of NASA, p. 63.
NASA’s Contributions to Aeronautics

102

2 was much more refined than its predecessor was, with sharper features 

at the redesigned nose and canopy. An even more powerful version of the 

J57 turbojet engine produced 17,000 pounds thrust with afterburner. The 

primary difference was the contoured fuselage that resembled a wasp’s 

waist and obvious fairings that expanded the circumference of the tail. 

With an area rule fuselage, the newly re-designed YF-102A easily went 

supersonic. Convair test pilot Pete Everest undertook the second flight 

test on December 21, 1954, during which the YF-102A climbed away 

from Lindbergh Field, San Diego, and “slipped easily past the sound 

barrier and kept right on going.” More importantly, the YF-102A’s top 

speed was 25 percent faster, at Mach 1.2.37

The Air Force resumed the contract with Convair, and the manufacturer delivered 975 production F-102A air defense interceptors, 

with the first entering active service in mid-1956. The fighter-interceptors equipped Air Defense Command and United States Air Force in 

Europe squadrons during the critical period of the late 1950s and 1960s. 

The increase in performance was dramatic. The F-102A could cruise at 

1,000 mph and at a ceiling of over 50,000 feet. It replaced three subsonic 

interceptor aircraft in the Air Force inventory—the North American 

F-86D Sabre, F-89 Scorpion, and F-94 Starfire—which were 600–650 

mph aircraft with a 45,000-foot ceiling range. Besides speed and altitude, the F-102A was better equipped to face the Soviet Myasishchev 

Bison, Tupolev Bear, and Ilyushin Badger nuclear-armed bombers 

with a full complement of Hughes Falcon guided missiles and Mighty 

Mouse rockets. Convair incorporated the F-102A’s armament in a dragreducing internal weapons bay.

When the F-102A entered operational service, the media made much 

of the fact that the F-102 “almost ended up in the discard heap” because 

of its “difficulties wriggling its way through the sound barrier.” With 

an area rule fuselage, the F-102A “swept past the sonic problem.” The 

downside to the F-102A’s supersonic capability was the noise from its 

J57 turbojet. The Air Force regularly courted civic leaders from areas 

near Air Force bases through familiarization flights so that they would 

understand the mission and role of the F-102A.38

37. Hansen, Engineer in Charge, 338; Swanborough, United States Military Aircraft Since 1909, p. 

152; Hallion, On the Frontier, pp. 90, 144 [quote]; Baals and Corliss, Wind Tunnels of NASA, p. 63.

38. Swanborough, United States Military Aircraft Since 1909, pp. 152, 154–155; Richard Witkin, “Supersonic Jets Will Defend City,” New York Times, Jan 3, 1957, p. 12.
Case 2 | Richard Whitcomb and the Quest for Aerodynamic Efficiency

103

2

The Air Force’s F-102 got a whole new look after implementing Richard Whitcomb’s area rule. 

At left is the YF-102 without the area rule, and at right is the new YF-102A version. NASA.

Convair produced the follow-on version, the F-106 Delta Dart, from 

1956 to 1960. The Dart was capable of twice the speed of the Dagger with 

its Pratt & Whitney J75 engine.39 The F-106 was the primary air defense 

interceptor defending the continental United States up to the early 1980s. 

Convair built upon its success with the F-102A and the F-106, two cornerstone aircraft in the Air Force’s Century series of aircraft, and introduced more area rule aircraft: the XF2Y-1 Sea Dart and the B-58 Hustler.40

The YF-102/YF-102A exercise was valuable in demonstrating the 

importance of the area rule and of the NACA to the aviation industry and 

the military, especially when a major contract was at stake.41 Whitcomb’s 

revolutionary and intuitive idea enabled a new generation of supersonic 

military aircraft, and it spread throughout the industry. Like Convair, 

Chance Vought redesigned its F8U Crusader carrier-based interceptor 

with an area rule fuselage. The first production aircraft appeared in 

September 1956, and deliveries began in March 1957. Four months later, 

in July 1957, Marine Maj. John H. Glenn, Jr., as part of Project Bullet, 

39. Swanborough, United States Military Aircraft Since 1909, pp. 152, 154–155.

40. Hallion, On the Frontier, p. 57.

41. Ibid., p. 96.
NASA’s Contributions to Aeronautics

104

2 made a recordbreaking supersonic transcontinental flight from Los 

Angeles to New York in 3 hours 23 minutes. Crusaders served in Navy 

and Marine fighter and reconnaissance squadrons throughout the 1960s 

and 1970s, with the last airframes leaving operational service in 1987.42

Grumman was the first to design and manufacture from the ground 

up an area rule airplane. Under contract to produce a carrier-based 

supersonic fighter, the F9F-9 Tiger, for the Navy, Grumman sent a team 

of engineers to Langley, just 2 weeks after receiving Whitcomb’s pivotal 

September 1952 report, to learn more about transonic drag. Whitcomb 

traveled to Bethpage, NY, in February 1953 to evaluate the design before 

wind tunnel and rocket-model tests were to be conducted by the NACA. 

The tests revealed that the new fighter was capable of supersonic speeds 

in level flight with no appreciable transonic drag. Grumman constructed 

the prototype, and in August 1954, with company test pilot C.H. “Corky” 

Meyer at the controls, the F9F-9 achieved Mach 1 in level flight without 

the assistance of an afterburner, which was a good 4 months before the 

supersonic flight of the F-102A.43 The Tiger, later designated the F11F1, served with the fleet as a frontline carrier fighter from 1957 to 1961 

and with the Navy’s demonstration team, the Blue Angels.44

Another aircraft designed from the ground up with an area rule 

fuselage represented the next step in military aircraft performance in 

the late 1950s. The legendary Lockheed “Skunk Works” introduced the 

F-104 Starfighter, “the missile with a man in it,” in 1954. Characterized 

by its short, stubby wings and needle nose, the production prototype 

F-104, powered by a General Electric J79 turbojet, was the first jet to 

exceed Mach 2 (1,320 mph) in flight, on April 24, 1956. Starfighters 

joined operational Air Force units in 1958. An international manufacturing scheme and sales to 14 countries in Europe, Asia, and the 

Middle East ensured that the Starfighter was in frontline use through the 

rest of the 20th century.45

42. Gordon Swanborough and Peter M. Bowers, United States Navy Aircraft Since 1911

(Annapolis: Naval Institute Press, 1990), pp. 456, 459; Barrett Tillman, MiG Master: The Story of 

the F-8 Crusader (Annapolis: Naval Institute Press, 2007), pp. 55–60.

43. Anderton, “NACA Formula Eases Supersonic Flight,” 15; Hansen, Engineer in Charge, p. 339

44. René J. Francillon, Grumman Aircraft Since 1929 (Naval Institute Press, 1989), p. 377; Swanborough and Bowers, United States Navy Aircraft Since 1911, pp. 256–257.

45. René J. Francillon, Lockheed Aircraft Since 1913 (Annapolis: Naval Institute Press, 1987), pp. 

329, 331, 342.
Case 2 | Richard Whitcomb and the Quest for Aerodynamic Efficiency

105

2

The area rule profile of the Grumman Tiger. National Air and Space Museum.

The area rule opened the way for the further refinement of supersonic aircraft, which allowed for concentration on other areas within 

the synergistic system of the airplane. Whitcomb and his colleagues continued to issue reports refining the concept and giving designers more 

options to design aircraft with higher performance. Working by himself 

and with researcher Thomas L. Fischetti, Whitcomb worked to refine 

high-speed aircraft, especially the Chance Vought F8U-1 Crusader, which 

evolved into one of the finest fighters of the postwar era.46

Spurred on by the success of the F-104, NACA researchers at 

the Lewis Flight Propulsion Laboratory in Cleveland, OH, estimated 

that innovations in jet engine design would increase aircraft speeds 

46. Richard T. Whitcomb and Thomas L. Fischetti, “Development of a Supersonic Area Rule and an 

Application to the Design of a Wing-Body Combination Having High Lift-to-Drag Ratios,” NACA 

RM-L53H31A (Aug. 18, 1953); and Richard T. Whitcomb, “Some Considerations Regarding the 

Application of the Supersonic Area Rule to the Design of Airplane Fuselages,” NACA RM-L56E23a 

(July 3, 1956).
NASA’s Contributions to Aeronautics

106

2 upward of 2,600 mph, or Mach 4, based on advanced metallurgy 

and the sophisticated aerodynamic design of engine inlets, including 

variable-geometry inlets and exhaust nozzles.47 One thing was for 

certain: supersonic aircraft of the 1950s and 1960s would have an area 

rule fuselage.

The area rule gave the American defense establishment breathing 

room in the tense 1950s, when the Cold War and the constant need to 

possess the technological edge, real or perceived, was crucial to the survival of the free world. The design concept was a state secret at a time 

when no jets were known to be capable of reaching supersonic speeds, 

due to transonic drag. The aviation press had known about it since 

January 1954 and kept the secret for national security purposes. The 

NACA intended to make a public announcement when the first aircraft 

incorporating the design element entered production. Aero Digest unofficially broke the story a week early in its September 1955 issue, when 

it proclaimed, “The SOUND BARRIER has been broken for good,” and 

declared the area rule the “first major aerodynamic breakthrough in 

the past decade.” In describing the area rule and the Grumman XF9F-9 

Tiger, Aero Digest stressed the bottom line for the innovation: the area 

rule provided the same performance with less power.48

The official announcement followed. Secretary of the Air Force Donald 

A. Quarles remarked on the CBS Sunday morning television news program 

Face the Nation on September 11, 1955, that the area rule was “the kind 

of breakthrough that makes fundamental research so very important.”49

Aviation Week declared it “one of the most significant military scientific 

breakthroughs since the atomic bomb.”50 These statements highlighted 

the crucial importance of the NACA to American aeronautics.

The news of the area rule spread out to the American public. The 

media likened the shape of an area rule fuselage to a “Coke bottle,” a 

47. Richard Witkin, “Aviation: 2,600 M.P.H.,” New York Times, Oct. 20, 1957, p. X33.

48. “Aero News Digest,” Aero Digest (Sept. 1955): p. 5. Aero Digest released the story without 

permission because publisher Fred Hamlin learned that the NACA had arranged, without his knowledge, to make the announcement in the rival journal, Aviation Week, on Sept. 19. “New Design 

Increasing Airplane Speeds Hailed,” Los Angeles Times, Sept. 12, 1955, p. 10; Alvin Shuster, 

“‘Pinch-Waist’ Plane Lifts Supersonic Speed 25%,” New York Times, Sept. 12, 1955, p. 15.

49. Quoted in Alvin Shuster, “‘Pinch-Waist’ Plane Lifts Supersonic Speed 25%,” New York Times, 

Sept. 12, 1955, p. 15.

50. Commentary by Robert Hotz, “The Area-Rule Breakthrough,” Aviation Week (Sept. 12, 1955), 

p. 152.
Case 2 | Richard Whitcomb and the Quest for Aerodynamic Efficiency

107

2 “wasp waist,” an “hourglass,” or the figure of actress Marilyn Monroe.51

While the Coke bottle description of the area rule is commonplace 

today, the NACA contended that Dietrich Küchemann’s Coke bottle and 

Whitcomb’s area rule were not the same and lamented the use of the 

term. Küchemann’s 1944 design concept pertained only to swept wings 

and tailored the specific flow of streamlines. Whitcomb’s rule applied to 

any shape and contoured a fuselage to maintain an area equivalent to 

the entire stream tube.52 Whitcomb actually preferred “indented.”53 One 

learned writer explained to readers of the Christian Science Monitor that 

an aircraft with an area rule slipped through the transonic barrier due to 

the “Huckleberry Finn technique,” which the character used to suck in 

his stomach to squeeze through a hole in Aunt Polly’s fence.54

Whitcomb quickly received just recognition from the aeronautical 

community for his 3-year development of the area rule. The National 

Aeronautics Association awarded him the Collier Trophy for 1954 for his 

creation of “a powerful, simple, and useful method” of reducing transonic 

drag and the power needed to overcome it.55 Moreover, the award citation designated the area rule as “a contribution to basic knowledge” that 

increased aircraft speed and range while reducing drag and using the same 

power.56 As Vice President Richard M. Nixon presented him the award at 

the ceremony, Whitcomb joined the other key figures in aviation history, 

including Orville Wright, Glenn Curtiss, and his boss, John Stack, in the 

pantheon of individuals crucial to the growth of American aeronautics.57

Besides the Collier, Whitcomb received the Exceptional Service Medal 

of the U.S. Air Force in 1955 and the inaugural NACA Distinguished 

Service Medal in 1956.58 At the age of 35, he accepted an honorary doctor of engineering degree from his alma mater, Worcester Polytechnic 

51. “New Plane Shape Increases Speed,” The Washington Post-Times Herald, Sept. 12, 1955, 

p. 18; “New Design Increasing Airplane Speeds Hailed,” Los Angeles Times, Sept. 12, 1955, 

p. 10; “Radial Shift in Air Design Bared by U.S.,” Christian Science Monitor, Sept. 12, 1955, p. 1.

52. “Area Rule and Coke Bottle,” Aviation Week (Sept. 12, 1955): p. 13. This source appeared 

as a sidebar in Anderton, “NACA Formula Eases Supersonic Flight.”

53. Richard Witkin, “The ‘Wasp-Waist’ Plane,” New York Times, Oct. 2, 1955, p. 20.

54. Maurice A. Garbell, “Transonic Planes Cut Drag with ‘Wasp Waist,’” Christian Science Monitor, Oct. 14, 1955, p. 5.

55. James J. Hagerty, Jr., “The Collier Trophy Winner,” Collier’s (Dec. 9, 1955): n.p.

56. “Designer to Be Honored For Pinched-Waist Plane,” New York Times, Nov. 23, 1955, p. 48.

57. Neal Stanford, “Wing Design Seeks Speed,” Christian Science Monitor, Feb. 17, 1970, p. 5.

58. “Whitcomb Receives NACA’s First DSM,” U.S. Air Services (Oct. 1956): p. 20.
NASA’s Contributions to Aeronautics

108

2

Institute, in 1956.59 Whitcomb also rose within the ranks at Langley, 

where he became head of Transonic Aerodynamics Branch in 1958.

Whitcomb’s achievement was part of a highly innovative period for 

Langley and the rest of the NACA, all of which contributed to the success 

of the second aeronautical revolution. Besides John Stack’s involvement 

in the X-1 program, the NACA worked with the Air Force, Navy, and the 

aerospace industry on the resultant high-speed X-aircraft programs. Robert 

T. Jones developed his swept wing theory. Other NACA researchers generated design data on different aircraft configurations, such as variablesweep wings, for high-speed aircraft. Whitcomb was directly involved 

in two of these major innovations: the slotted tunnel and the area rule.60

Inventing the Supercritical Wing

Whitcomb was hardly an individual content to rest on his laurels or bask 

in the glow of previous successes, and after his success with area ruling, he wasted no time in moving further into the transonic and supersonic research regime. In the late 1950s, the introduction of practical 

subsonic commercial jetliners led many in the aeronautical community 

to place a new emphasis on what would be considered the next logical 

step: a Supersonic Transport (SST). John Stack recognized the importance of the SST to the aeronautics program in NASA in 1958. As NASA 

placed its primary emphasis on space, he and his researchers would work 

on the next plateau in commercial aviation. Through the Supersonic 

Transport Research Committee, Stack and his successor, Laurence K. 

Loftin, Jr., oversaw work on the design of a Supersonic Commercial Air 

Transport (SCAT). The goal was to create an airliner capable of outperforming the cruise performance of the Mach 3 North American XB-70 

Valkyrie bomber. Whitcomb developed a six-engine arrowlike highly 

swept wing SST configuration that stood out as possessing the best liftto-drag (L/D) ratio among the Langley designs called SCAT 4.61

59. Rowes, “When You Ride Tomorrow’s Airplanes, You’ll Thank Dick Whitcomb.”

60. Joseph R. Chambers, Innovation in Flight: Research of the NASA Langley Research Center 

on Revolutionary Advanced Concepts for Aeronautics, NASA SP-2005-4539 (Washington, DC: 

NASA, 2005), p. 18.

61. Whitcomb also rejected the committee’s emphasis on variable-geometry wings as too heavy, 

which led to his ejection from the design committee by Stack. Eric M. Conway, High Speed 

Dreams: NASA and the Technopolitics of Supersonic Transportation, 1945–1999 (Baltimore: Johns 

Hopkins University Press, 2005), pp. 54–55; Becker, High Speed Frontier, pp. 55–56.
Case 2 | Richard Whitcomb and the Quest for Aerodynamic Efficiency

109

2 Manufacturers’ analyses indicated that Whitcomb’s SCAT 4 exhibited the lowest range and highest weight among a group of designs that 

would generate high operating and fuel costs and was too heavy when 

compared with subsonic transports. Despite President John F. Kennedy’s 

June 1963 commitment to the development of “a commercially successful supersonic transport superior to that being built in any other country 

in the world,” Whitcomb saw the writing on the wall and quickly disassociated himself from the American supersonic transport program in 

1963.62 Always keeping in mind his priorities based on practicality and 

what he could do to improve the airplane, Whitcomb said: “I’m going 

back where I know I can make things pay off.”63 For Whitcomb, practicality outweighed the lure of speed equated with technological progress.

Whitcomb decided to turn his attention back toward improving subsonic aircraft, specifically a totally new airfoil shape. Airfoils and wings 

had been evolving over the course of the 20th century. They reflected the 

ever-changing knowledge and requirements for increased aircraft performance and efficiency. They also represented the bright minds that developed them. The thin cambered airfoil of the Wright brothers, the thick 

airfoils of the Germans in World War I, the industry-standard Clark Y 

of the 1920s, and the NACA four- and five-digit series airfoils innovated 

by Eastman Jacobs exemplified advances in and general approaches 

toward airfoil design and theory.64

Despite these advances and others, subsonic aircraft flew at 85-percent 

efficiency.65 The problem was that, as subsonic airplanes moved toward 

their maximum speed of 660 mph, increased drag and instability developed. Air moving over the upper surface of wings reached supersonic 

speeds, while the rest of the airplane traveled at a slower rate. The plane 

had to fly at slower speeds at decreased performance and efficiency.66

When Whitcomb returned to transonic research in 1964, he specifically wanted to develop an airfoil for commercial aircraft that delayed the 

onset of high transonic drag near Mach 1 by reducing air friction and turbu62. Conway, High Speed Dreams, p. 55; Gorn, Expanding the Envelope, p. 56; Quote from 

Chambers, Innovation in Flight, p. 28.

63. Gorn, Expanding the Envelope, p. 331.

64. For more information on the history of airfoils and their theorists and designers, see Anderson, A 

History of Aerodynamics.

65. Rowes, “When You Ride Tomorrow’s Airplanes, You’ll Thank Dick Whitcomb,” p. 165.

66. Thomas Grubisich, “Fuel-Saver in Wings,” The Washington Post, July 11, 1974, p. C1.
NASA’s Contributions to Aeronautics

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2

Whitcomb inspecting a supercritical wing model in the 8-Foot TPT. NASA.

lence across an aircraft’s major aerodynamic surface, the wing. Whitcomb 

went intuitively against conventional airfoil design, in which the upper surface curved downward on the leading and trailing edges to create lift. He 

envisioned a smoother flow of air by turning a conventional airfoil upside 

down. Whitcomb’s airfoil was flat on top with a downward curved rear section.67 The shape delayed the formation of shock waves and moved them 

further toward the rear of the wing to increase total wing efficiency. The 

rear lower surface formed into deeper, more concave curve to compensate for the lift lost along the flattened wing top. The blunt leading edge 

facilitated better takeoff, landing, and maneuvering performance. Overall, 

Whitcomb’s airfoil slowed airflow, which lessened drag and buffeting, and 

improved stability.68

With the wing captured in his mind’s eye, Whitcomb turned it into 

mathematical calculations and transformed his findings into a wind tunnel model created by his own hands. He spent days at a time in the 8-foot 

Transonic Pressure Tunnel (TPT), sleeping on a nearby cot when needed, 

as he took advantage of the 24-hour schedule to confirm his findings.69

67. Gorn, Expanding the Envelope, p. 331.

68. Grubisich, “Fuel-Saver in Wings.”

69. Gorn, Expanding the Envelope, p. 331.
Case 2 | Richard Whitcomb and the Quest for Aerodynamic Efficiency

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2 Just as if he were still in his boyhood laboratory, Whitcomb stated that: 

“When I’ve got an idea, I’m up in the tunnel. The 8-foot runs on two shifts, 

so you have to stay with the job 16 hours a day. I didn’t want to drive back 

and forth just to sleep, so I ended up bringing a cot out here.”70

Whitcomb and researcher Larry L. Clark published their wind tunnel 

findings in “An Airfoil Shape for Efficient Flight at Supercritical Mach 

Numbers,” which summarized much of the early work at Langley. Their 

investigation compared a supercritical airfoil with a NASA airfoil. They 

concluded that the former developed more abrupt drag rise than the 

latter.71 Whitcomb presented those initial findings at an aircraft aerodynamics conference held at Langley in May 1966.72 He called his new innovation a “supercritical wing” by combining “super” (meaning “beyond”) 

with “critical” Mach number, which is the speed supersonic flow revealed 

itself above the wing. Unlike a conventional wing, where a strong shock 

wave and boundary layer separation occurred in the transonic regime, 

a supercritical wing had both a weaker shock wave and less developed 

boundary layer separation. Whitcomb’s tests revealed that a supercritical wing with 35-degree sweep produced 5 percent less drag, improved 

stability, and encountered less buffeting than a conventional wing at 

speeds up to Mach 0.90.73

Langley Director of Aeronautics Laurence K. Loftin believed that 

Whitcomb’s new supercritical airfoil would reduce transonic drag and 

result in improved fuel economy. He also knew that wind tunnel data alone 

would not convince aircraft manufacturers to adopt the new airfoil. Loftin 

first endorsed the independent analyses of Whitcomb’s idea at the Courant 

Institute at New York University, which proved the viability of the concept. 

More importantly, NASA had to prove the value of the new technology to 

industry by actually building, installing, and flying the wing on an aircraft.74

70. Welch, “Whitcomb,” p. 5.

71. Richard T. Whitcomb and Larry L. Clark, “An Airfoil Shape for Efficient Flight at Supercritical 

Mach Numbers,” NASA TM-X-1109 (Apr. 20, 1965).

72. For a first-person account of the development of the supercritical wing, see Richard T. Whitcomb, “Research Associated with the Langley 8-Foot Tunnels Branch: Lecture at Ames Research 

Center, October 21, 1970,” NASA TM-108686 (1970).

73. Richard T. Whitcomb, “The State of Technology Before the F-8 Supercritical Wing,” in Proceedings of the F-8 Digital Fly-By-Wire and Supercritical Wing First Flight’s 20th Anniversary, May 27, 

1992, NASA CP-3256, vol. 1 (Washington, DC: NASA, 1996), p. 81; Gorn, Expanding the 

Envelope, pp. 331–332.

74. Gorn, Expanding the Envelope, pp. 332, 394, 401.
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2 The major players met in March 1967 to discuss turning Whitcomb’s 

concept into a reality. The practicalities of manufacturing, flight characteristics, structural integrity, and safety required a flight research 

program. The group selected the Navy Chance Vought F-8A fighter as the 

flight platform. The F-8A possessed specific attributes that made it ideal 

for the program. While not an airliner, the F-8A had an easily removable 

modular wing readymade for replacement, fuselage-mounted landing 

gear that did not interfere with the wing, engine thrust capable of operation in the transonic regime, and lower operating costs than a multi-engine 

airliner. Langley contracted Vought to design a supercritical wing for the 

F-8 and collaborated with Whitcomb during wind tunnel testing beginning during the summer of 1967. Unfortunately for the program, NASA 

Headquarters suspended all ongoing contracts in January 1968 and Vought 

withdrew from the program.75

SCW Takes to the Air

Langley and the Flight Research Center entered into a joint program outlined in a November 1968 memorandum. Loftin and Whitcomb lead a 

Langley team responsible for defining the overall objectives, determining 

the wing contours and construction tolerances, and conducting wind tunnel tests during the flight program. Flight Research Center personnel determined the size, weight, and balance of the wing; acquired the F-8A airframe 

and managed the modification program; and conducted the flight research 

program. North American Rockwell won the contract for the supercritical wing and delivered it to the Flight Research Center in November 1970 

at a cost of $1.8 million. Flight Research Center technicians installed the 

new wing on a Navy surplus TF-8A trainer.76 At the onset of the flight program, Whitcomb predicted the new wing design would allow airliners to 

cruise 100 mph faster and close to the speed of sound (nearly 660 mph) at 

an altitude of 45,000 feet with the same amount of power.77

NASA test pilot Thomas C. McMurtry took to the air in the F-8 

Supercritical Wing flight research vehicle on March 9, 1971. Eighty-six 

75. Thomas C. Kelly and Richard T. Whitcomb, “Evolution of the F-8 Supercritical Wing Configuration,” in Supercritical Wing Technology—A Progress Report on Flight Evaluations, NASA SP-301 

(1972), p. 35; Gorn, Expanding the Envelope, pp. 332–333.

76. Kelly and Whitcomb, “Evolution of the F-8 Supercritical Wing Configuration,” in Supercritical 

Wing Technology, p. 35; Gorn, Expanding the Envelope, pp. 333–334.

77. Neal Stanford, “Wing Design Seeks Speed,” Christian Science Monitor, Feb. 17, 1970, p. 5.
Case 2 | Richard Whitcomb and the Quest for Aerodynamic Efficiency

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2 flights later, the program ended on May 23, 1973. A pivotal document generated during the program was Supercritical Wing Technology—A Progress 

Report on Flight Evaluations, which captured the ongoing results of the 

program. From the standpoint of actually flying the F-8, McMurtry noted 

that: “the introduction of the supercritical wing is not expected to create 

any serious problems in day-to-day transport operations.” The combined 

flight and wind tunnel tests revealed increased efficiency of commercial 

aircraft by 15 percent and, more importantly, a 2.5-percent increase in 

profits. In the high-stakes business of international commercial aviation, 

the supercritical wing and its ability to increase the range, speed, and fuel 

efficiency of subsonic jet aircraft without an increase in required power

or additional weight was a revolutionary new innovation.78

NASA went beyond flight tests with the F-8, which was a flight-test 

vehicle built specifically for proving the concept. The Transonic Aircraft 

Technology (TACT) program was a joint NASA–U.S. Air Force partnership begun in 1972 that investigated the application of supercritical wing 

technology to future combat aircraft. The program evaluated a modified 

General Dynamics F-111A variable-sweep tactical aircraft to ascertain 

its overall performance, handling qualities, and transonic maneuverability and to define the local aerodynamics of the airfoil and determine 

wake drag. Whitcomb worked directly with General Dynamics and the 

Air Force Flight Dynamics Laboratory on the concept.79 NASA worked to 

refine the supercritical wing, and its resultant theory through continued 

comparison of wind tunnel and flight tests that continued the Langley 

and Flight Research Center collaboration.80

Whitcomb developed the supercritical airfoil using his logical cutand-try procedures. Ironically, what was considered to be an unsophisticated research technique in the second half of the 20th century, 

a process John Becker called “Edisonian,” yielded the complex super78. Thomas C. McMurtry, Neil W. Matheny, and Donald H. Gatlin, “Piloting and Operational 

Aspects of the F-8 Supercritical Wing Airplane,” in Supercritical Wing Technology—A Progress 

Report on Flight Evaluations. NASA SP-301, (Washington, DC, NASA, 1972), p. 102; Gorn, 

Expanding the Envelope, pp. 335, 337.

79. Joseph Well, “Summary and Future Plans,” in Supercritical Wing Technology, pp. 127–128.

80. See Jon S. Pyle and Louis L. Steers, “Flight-Determined Lift and Drag Characteristics of an F-8 

Airplane Modified with a Supercritical Wing with Comparisons to Wind Tunnel Results,” NASA 

TM-X-3250 (Jan. 16, 1975); and Lawrence C. Montoya and Richard D. Banner, “F-8 Supercritical 

Wing Flight Pressure, Boundary-Layer, and Wake Measurements and Comparisons with Wind Tunnel Data,” NASA TM-X-3544 (June 1977).
NASA’s Contributions to Aeronautics

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2 critical airfoil. The key, once again, was the fact that the researcher, 

Whitcomb, possessed “truly unusual insights and intuitions.”81 Whitcomb 

used his intuitive imagination to search for a solution over the course 

of 8 years. Mathematicians verified his work after the fact and created a 

formula for use by the aviation industry.82 Whitcomb received patent No. 

3,952,971 for his supercritical wing in May 1976. NASA possessed the 

rights to granting licenses, and several foreign nations already had filed 

patent applications.83

The spread of the supercritical wing to the aviation industry was 

slow in the late 1970s. There was no doubt that the supercritical wing 

possessed the potential of saving the airline industry $300 million annually. Both Government experts and the airlines agreed on its new importance. Unfortunately, the reality of the situation in the mid-1970s was 

that the purchase of new aircraft or conversion of existing aircraft would 

cost the airlines millions of dollars, and it was estimated that $1.5 billion in fuel costs would be lost before the transition would be completed. The impetus would be a fuel crisis like the Arab oil embargo, 

during which the price per gallon increased from 12 to 30 cents within 

the space of a year.84

The introduction of the supercritical wing on production aircraft 

centered on the Air Force’s Advanced Medium Short Take-Off and 

Landing (STOL) Transport competition between McDonnell-Douglas 

and Boeing to replace the Lockheed C-130 Hercules in the early 1970s. 

The McDonnell-Douglas design, the YC-15, was the first large transport 

with supercritical wings in 1975. Neither the YC-15 nor the Boeing YC-14 

replaced the Hercules because of the cancellation of the competition, 

but their wings represented to the press an “exotic advance” that provided new levels of aircraft fuel economy in an era of growing fuel costs.85

During the design process of the YC-14, Boeing aerodynamicists also 

selected a supercritical airfoil for the wing. They based their decision 

on previous research with the 747 airliner wing, data from Whitcomb’s 

research at Langley, and the promising performance of a Navy T-2C 

Buckeye that North American Aviation modified with a supercritical air81. Becker, High Speed Frontier, p. 59.

82. Grubisich, “Fuel-Saver in Wings.”

83. Stacy V. Jones, “New Aircraft Wing Invented,” New York Times, May 1, 1976, p. 46.

84. Grubisich, “Fuel-Saver in Wings.”

85. Richard Witkin, “McDonnell Douglas Unveils New Cargo Jet,” New York Times, Aug. 6, 1975, p. 65.
Case 2 | Richard Whitcomb and the Quest for Aerodynamic Efficiency

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2 foil to gain experience for the F-8 wing project and undergoing flight 

tests in November 1969. Boeing’s correlation of wind tunnel and flight 

test data convinced the company to introduce supercritical airfoils on 

the YC-14 and for all of its subsequent commercial transports, including the triumphant “paperless” airplane, the 777 of the 1990s.86

The business jet community embraced the supercritical wing in the 

increasingly fuel- and energy-conscious 1970s. Business jet pioneer Bill 

Lear incorporated the new technology in the Canadair Challenger 600, 

which took to the air in 1978. Rockwell International incorporated the 

technology into the upgraded Sabreliner 65 of 1979. The extensively 

redesigned Dassault Falcon 50, introduced the same year, relied upon a 

supercritical wing that enabled an over-3,000-mile range.87

The supercritical wing program gave NASA the ability to stay in the 

public eye, as it was an obvious contribution to aeronautical technology. The program also improved public relations and the stature of both 

Langley and Dryden at a time in the 1960s and 1970s when the first “A” 

in NASA—aeronautics—was secondary to the single “S”—space. For this 

reason, historian Richard P. Hallion has called the supercritical wing 

program “Dryden’s life blood” in the early 1970s.88

Subsonic transports, business jets, STOL aircraft, and uncrewed 

aerial vehicles incorporate supercritical wing technology today.89 All 

airliners today have supercritical airfoils custom-designed and finetuned by manufacturers with computational fluid dynamics software 

programs. There is no NASA supercritical airfoil family like the significant NACA four- and five-airfoil families. The Boeing 777 wing embodies a Whitcomb heritage. This revolutionary information appeared in 

NASA technical notes (TN) and other publications with little or no fanfare and through direct consultation with Whitcomb. A Lockheed engineer and former employee of Whitcomb in the late 1960s remarked on 

his days at NASA Langley:

When I was working for Dick Whitcomb at NASA, there was hardly 

a week that went by that some industry person did not come in to see 

him. It was a time when NASA was being constantly asked for technical 

86. Chambers, Innovation in Flight, p. 183; Hallion, On the Frontier, p. 204.

87. Hallion, On the Frontier, pp. 206–207.

88. Ibid., p. 172.

89. For an overview of NASA development of supercritical airfoils up to 1990, see Charles D. Harris, 

“NASA Supercritical Airfoils—A Matrix of Family-Related Airfoils,” NASA TP-2969 (1990).
NASA’s Contributions to Aeronautics

116

2 advice, and Dick always gave that advice freely. He was always there 

when industry wanted him to help out. This is the kind of cooperation 

that makes industry want to work with NASA. As a result of that sharing, 

we have seen the influence of supercritical technology to go just about 

every corner of our industry.90

Whitcomb set the stage and the direction of contemporary 

aircraft design.

More accolades were given to Whitcomb by the Government and 

industry during the years he worked on the supercritical wing. From 

NASA, he received the Medal for Exceptional Scientific Achievement in 

1969, and 5 years later, NASA Administrator James Fletcher awarded 

Whitcomb $25,000 in cash for the invention of the supercritical wing 

from NASA in June 1974. The NASA Inventions and Contributions Board 

recommended the cash prize to recognize individual contributions to the 

Agency’s programs. It was the largest cash award given to an individual 

at NASA.91 In 1969, Whitcomb accepted the Sylvanus Albert Reed Award 

from the American Institute of Aeronautics and Astronautics, the organization’s highest honor for achievement in aerospace engineering. In 

1973, President Richard M. Nixon presented him the highest honor for 

science and technology awarded by the U.S. Government, the National 

Medal of Science.92 The National Aeronautics Association bestowed 

upon Whitcomb the Wright Brothers Memorial Trophy in 1974 for his 

dual achievements in developing the area rule and supercritical wing.93

Winglets—Yet Another Whitcomb Innovation

Whitcomb continued to search for ways to improve the subsonic airplane beyond his work on supercritical airfoils. The Organization of the 

Petroleum Exporting Countries (OPEC) oil embargo of 1973–1974 dramatically affected the cost of airline operations with high fuel prices.94 NASA 

implemented the Aircraft Energy Efficiency (ACEE) program as part of 

90. Blackwell, “Influence on Today’s Aircraft,” p. 114.

91. “Dr. Whitcomb to Receive $25,000 Award from NASA,” NASA Release No. 74-148 (June 4, 

1974): pp. 1, 3, File CW-463000-01, National Air and Space Museum Archives; Gorn, Expanding the Envelope, p. 337.

92. Grubisich, “Fuel-Saver in Wings.”

93. “Richard Travis Whitcomb: Distinguished Research Associate,” NASA Langley Research Center, 

Apr. 1983.

94. Welch, “Whitcomb,” p. 5.
Case 2 | Richard Whitcomb and the Quest for Aerodynamic Efficiency

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2 the national energy conservation effort in the 1970s. At this time, Science

magazine featured an article discussing how soaring birds used their tip 

feathers to control flight characteristics. Whitcomb immediately shifted 

focus toward the wingtips of an aircraft—specifically flow phenomena 

related to induced drag—for his next challenge.95

Two types of drag affect the aerodynamic efficiency of a wing: 

profile drag and induced drag. Profile drag is a two-dimensional phenomenon and is clearly represented by the iconic airflow in the slipstream 

image that represents aerodynamics. Induced drag results from threedimensional airflow near the wingtips. That airflow rolls up over the tip 

and produces vortexes trailing behind the wing. The energy exhausted 

in the wingtip vortex creates induced drag. Wings operating in high-lift, 

low-speed performance regimes can generate large amounts of induced 

drag. For subsonic transports, induced drag amounts to as much as 50 

percent of the total drag of the airplane.96

As part of the program, Whitcomb chose to address the wingtip vortex, the turbulent air found at the end of an airplane wing. These vortexes 

resulted from differences in air pressure generated on the upper and lower 

surfaces of the wing. As the higher-pressure air forms along the lower 

surface of the wing, it creates its own airflow along the length of the wing. 

At the wingtip, the airflow curls upward and forms an energy-robbing 

vortex that trails behind. Moreover, wingtip vortexes create enough 

turbulent air to endanger other aircraft that venture into their wake.

Whitcomb sought a way to control the wingtip vortex with a new 

aeronautical structure called the winglet. Winglets are vertical wing-like 

surfaces that extend above and sometimes below the tip of each wing. 

A winglet designer can balance the relationship between cant, the angle 

the winglet bends from the vertical, and toe, the angle the winglet deviates from airflow, to produce a lift force that, when placed forward of 

the airfoils, generates thrust from the turbulent wingtip vortexes. This 

phenomenon is akin to a sailboat tacking upwind while, in the words of 

aviation observer George Larson: “the keel squeezes the boat forward 

like a pinched watermelon seed.”97

95. Joseph R. Chambers, Concept to Reality: Contributions of the Langley Research Center to U.S. 

Civil Aircraft of the 1990s (Washington, DC: NASA, 2003), p. 35.

96. Ibid., p. 35.

97. George Larson, “Winglets,” Air & Space Magazine (Sept. 01, 2001), http://www.

airspacemag.com/flight-today/wing.html (Accessed Feb. 20, 2009).
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2 There were precedents for the use of what Whitcomb would call a 

“nonplanar,” or nonhorizontal, lifting system. It was known in the burgeoning aeronautical community of the late 1800s that the induced drag 

of wingtip vortexes degraded aerodynamic efficiency. Aeronautical pioneer Frederick W. Lanchester patented vertical surfaces, or “endplates,” 

to be mounted at an airplane’s wingtips, in 1897. His research revealed 

that vertical structures reduced drag at high lift. Theoretical studies 

conducted by the Army Air Service Engineering Division in 1924 and 

the NACA in 1938 in the United States and by the British Aeronautical 

Research Committee in 1956 investigated various nonplanar lifting systems, including vertical wingtip surfaces.98 They argued that theoretically, 

these structures would provide significant aerodynamic improvements 

for aircraft. Experimentation revealed that while there was the potential of reducing induced drag, the use of simple endplates produced too 

much profile drag to justify their use.99

Whitcomb and his research team investigated the drag-reducing 

properties of winglets for a first-generation, narrow-body subsonic jet 

transport in the 8-foot TPT from 1974 to 1976. They used a semispan 

model, meaning it was cut in half and mounted on the tunnel wall 

to enable the use of a larger test object that would facilitate a higher 

Reynolds number and the use of specific test equipment. He compared 

a wing with a winglet and the same wing with a straight extension to 

increase its span. The constant was that both the winglet and extension 

exerted the same structural load on the wing. Whitcomb found that winglets

reduced drag by approximately 20 percent and doubled the improvement 

in the lift-to-drag ratio to 9 percent compared with the straight wing extension. Whitcomb published his findings in “A Design Approach and Selected 

Wind-Tunnel Results at High Subsonic Speeds for Wing-Tip Mounted 

Winglets.”100 It was obvious that the reduction in drag generated by a pair 

of winglets boosted performance by enabling higher cruise speeds. 

98. See F. Nagel, Wings With End Plates. Memo. Rep. 130, Eng. Div., McCook Field, Nov. 4, 

1924; W. Mangler, “The Lift Distribution of Wings With End Plates,” NACA TM-856 (1938); J. Weber, 

Theoretical Load Distribution on a Wing with Vertical Plates. R. & M. No. 2960, British A.R.C., 1956.

99. Richard T. Whitcomb, “A Design Approach and Selected Wind-Tunnel Results at High Subsonic 

Speeds for Wing-Tip Mounted Winglets,” NASA TN-D-8260 (July 1976), p. 1; Chambers, Concept to Reality, p. 35.

100. Whitcomb, “A Design Approach and Selected Wind-Tunnel Results at High Subsonic Speeds 

for Wing-Tip Mounted Winglets,” NASA TN-D-8260 (July 1976), pp. 13–14.
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119

2 With the results, Whitcomb provided a general design approach for 

the basic design of winglets based on theoretical calculations, physical 

flow considerations, and emulation of his overall approach to aerodynamics, primarily “extensive exploratory experiments.” What made a 

winglet rather than a simple vertical surface attached to the end of a 

wing was the designer’s ability to use well-known wing design principles to incorporate side forces to reduce lift-induced inflow above the 

wingtip and outflow below the tip to create a vortex diffuser. The placement and optimum height of the winglet reflected both aerodynamic 

and structural considerations in which the designer had to take into 

account the efficiency of the winglet as well as its weight. For practical 

operational purposes, the lower portion of the winglet could not hang 

down far below the wingtip for fear of damage on the ground. The fact 

that the ideal airfoil shape for a winglet was NASA’s general aviation airfoil made it even easier to incorporate winglets into an aircraft design.101

Whitcomb’s basic rules provided that foundation.

Experimental wind tunnel studies of winglets in the 8-foot TPT 

continued through the 1970s. Whitcomb and his colleagues Stuart G. 

Flechner and Peter F. Jacobs concentrated next on the effects of winglets on a representative second-generation jet transport—the semispan 

model vaguely resembled a Douglas DC-10—at high subsonic speeds, 

specifically Mach 0.7 to 0.83. They concluded that winglets significantly 

reduced the induced drag coefficient while lowering overall drag. The 

smoothing out of the vortex behind the wingtip by the winglet accounted 

for the reduction in induced drag. As in the previous study, they saw that 

winglets generated a small increase in lift. The researchers calculated that 

winglets reduced drag better than simple wingtip extensions did, despite 

a minor increase in structural bending moments.102

Another benefit derived from winglets was the increase in the aspect 

ratio of wing without compromising its structural integrity. The aspect 

101. Ibid., pp. 1–2, 5, 13–14. Whitcomb also suggested consultation of the following two 

references regarding winglet design: John E. Lamar, “A Vortex-Lattice Method for the Mean Camber 

Shapes of Trimmed Noncoplanar Platforms with Minimum Vortex Drag,” NASA TN-D-8090 (1976) 

and M.I. Goldhammer, “A Lifting Surface Theory for the Analysis of Nonplanar Lifting Systems,” 

AIAA Paper No. 76-16 (Jan. 1976).

102. Stuart G. Flechner, Peter F. Jacobs, and Richard T. Whitcomb, “A High Subsonic Wind Tunnel 

Investigation of Winglets on a Representative Second-Generation Jet Transport Wing,” NASA TN8264 (July 1976), pp. 1, 13.
NASA’s Contributions to Aeronautics

120

2 ratio of a wing is the relationship between span—the distance from tip 

to tip—and chord—the distance between the leading and trailing edge. 

A long, thin wing has a high aspect ratio, which produces longer range 

at a certain cruise speed because it does not suffer from wingtip vortexes 

and the corresponding energy losses as badly as a short and wide chord 

low aspect ratio wing. The drawback to a high aspect ratio wing is that 

its long, thin structure flexes easily under aerodynamic loads. Making 

this type of wing structurally stable required strengthening that added 

weight. Winglets offered increased aspect ratio with no increase in wingspan. For every 1-foot increase in wingspan, meaning aspect ratio, there 

was an increase in wing-bending force. Wings structurally strong enough 

to support a 2-foot span increase would also support 3-foot winglets while 

producing the same gain in aspect ratio.103

NASA made sure the American aviation industry was aware of the 

results of Whitcomb’s winglet studies and its part in the ACEE program. 

Langley organized a meeting focusing on advanced technologies developed by NASA for Conventional Take-Off and Landing (CTOL) aircraft, 

primarily airliners, business jets, and personal aircraft, from February 28 

to March 3, 1978. During the session dedicated to advanced aero-dynamic 

controls, Flechner and Jacobs summarized the results of wind tunnel 

results on winglets applied to a Boeing KC-135 aerial tanker, Lockheed 

L-1011 and McDonnell-Douglas DC-10 airliners, and a generic model 

with high aspect ratio wings.104 Presentations from McDonnell-Douglas 

and Boeing representatives revealed ongoing industry work done under 

contract with NASA. Interest in winglets was widespread at the conference and after as manufacturers across the United States began to 

consider their use and current and future designs.105

Whitcomb’s winglets first found use on general aviation aircraft at 

the same time he and his colleagues at Langley began testing them on air 

transport models and a good 4 years before the pivotal CTOL conference. 

Another visionary aeronautical engineer, Burt Rutan, adopted them for 

his revolutionary designs. The homebuilt Vari-Eze of 1974 incorporated 

103. Larson, “Winglets.”

104. See also Stuart G. Flechner and Peter F. Jacobs, “Experimental Results of Winglets on First, 

Second, and Third Generation Jet Transports,” NASA TM-72674 (1978).

105. For these articles, see Conventional Take-off and Landing (CTOL) Transport Technology 1978: 

Proceedings of a Conference Held at Langley Research Center, Hampton, VA, Feb. 28–Mar. 3, 1978, 

NASA CP-2036, Parts I and II (Washington, DC: NASA, 1978); Chambers, Concept to Reality, p. 38.
Case 2 | Richard Whitcomb and the Quest for Aerodynamic Efficiency

121

2 winglets combined with vertical control surfaces. The airplane was an 

overall innovative aerodynamic configuration with its forward canard, 

high aspect ratio wings, low-weight composite materials, a lightweight 

engine, and pusher propeller, Whitcomb’s winglets on Rutan’s Vari-Eze 

offered private pilots a stunning alternative to conventional airplanes. 

His nonstop world-circling Voyager and the Beechcraft Starship of 1986 

also featured winglets.106

The business jet community was the first to embrace winglets and 

incorporate them into production aircraft. The first jet-powered airplane to enter production with winglets was the Learjet Model 28 in 

1977. Learjet was in the process of developing a new business jet, the 

Model 55, and built the Model 28 as a testbed to evaluate its new proprietary high aspect ratio wing and winglet system, called the Longhorn. 

The manufacturer developed the system on its own initiative without 

assistance from Whitcomb or NASA, but it was clear where the winglets 

came from. The comparison flight tests of the Model 28 with and without winglets showed that the former increased its range by 6.5 percent. 

An additional benefit was improved directional stability. Learjet exhibited the Model 28 at the National Business Aircraft Association convention and put it into production because of its impressive performance 

and included winglets on its successive business jets.107 Learjet’s competitor, Gulfstream, also investigated the value of winglets to its aircraft 

in the late 1970s. The Gulfstream III, IV, and V aircraft included winglets 

in their designs. The Gulfstream V, able to cruise at Mach 0.8 for a distance of 6,500 nautical miles, captured over 70 national and world flight 

records and received the 1997 Collier Trophy. Records aside, the ability 

to fly business travelers nonstop from New York to Tokyo was unprecedented after the introduction of the Gulfstream V in 1995.108

Actual acceptance on the part of the airline industry was mixed in the 

beginning. Boeing, Lockheed, and Douglas each investigated the possibility 

of incorporating winglets into current aircraft as part of the ACEE program. 

Winglets were a fundamental design technology, and each manufacturer 

106. Chambers, Concept to Reality, p. 41. During the takeoff for the world flight, one of Voyager’s 

winglets broke off, and pilot Dick Rutan had to severely maneuver the aircraft to break the other one 

off before the journey could continue.

107. Ibid., pp. 41–43.

108. Gulfstream, “The History of Gulfstream: 1958–2008,” 2009, http://www.

gulfstream.com/history (Accessed Feb. 15, 2009).
NASA’s Contributions to Aeronautics

122

2

The KC-135 winglet test vehicle in flight over Dryden. NASA.

had to design them for the specific airframe. NASA awarded contracts to 

manufacturers to experiment with incorporating them into existing and 

new designs. Boeing concluded in May 1977 that the economic benefits 

of winglets did not justify the cost of fabrication for the 747. Lockheed 

chose to extend the wingtips for the L-1011 and install flight controls 

to alleviate the increased structural loads. McDonnell-Douglas immediately embraced winglets as an alternative to increasing the span of a 

wing and modified a DC-10 for flight tests.109

The next steps for Whitcomb and NASA were flight tests to demonstrate the viability of winglets for first and second transport and airliner generations. Whitcomb and his team chose the Air Force’s Boeing 

KC-135 aerial tanker as the first test airframe. The KC-135 shared with 

its civilian version, the pioneering 707, and other early airliners and 

transports an outer wing that exhibited elliptical span loading with high 

loading at the outer panels. This wingtip loading was ideal for winglets. 

Additionally, the Air Force wanted to improve the performance and fuel 

efficiency of the aging aerial tanker. Whitcomb and this team designed 

the winglet, and Boeing handled the structural design and fabrication 

of winglets for an Air Force KC-135. NASA and the Air Force performed 

the flights tests at Dryden Flight Research Center in 1979 and 1980. 

The tests revealed a 20-percent reduction in drag because of lift, with a 

109. Chambers, Concept to Reality, p. 38.
Case 2 | Richard Whitcomb and the Quest for Aerodynamic Efficiency

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2 7-percent gain in the lift-to-drag ratio at cruise, which confirmed 

Whitcomb’s findings at Langley.110

McDonnell-Douglas conducted a winglet flight evaluation program 

with a DC-10 airliner as part of NASA’s Energy Efficient Transport (EET) 

program within the larger ACEE program in 1981. The DC-10 represented 

a second-generation airliner with a wing designed to produce nonelliptic 

loading to avoid wingtip pitch-up characteristics. As a result, the wing 

bending moments and structural requirements were not as dramatic as 

those found on a first-generation airliner, such as the 707. Whitcomb and 

his team conducted a preliminary wind tunnel examination of a DC-10 

model in the 8-foot TPT. McDonnell-Douglas engineers designed the 

aerodynamic and structural shape of the winglets and manufacturing 

personnel fabricated them. The company performed flights tests over 16 

months, which included 61 comparison flights with a DC-10 leased from 

Continental Airlines. These industry flight tests revealed that the addition 

of winglets to a DC-10, combined with a drooping of the outboard ailerons, produced a 3-percent reduction in fuel consumption at passengercarrying distances, which met the bottom line for airline operators.111

The DC-10 did not receive winglets because of the prohibitive cost 

of Federal Aviation Administration (FAA) recertification. Nevertheless, 

McDonnell-Douglas was a zealous convert and used the experience and 

design data for the advanced derivative of the DC-10, the MD-11, when 

that program began in 1986. The first flight in January 1990 and the grueling 10-month FAA certification process that followed validated the use 

of winglets on the MD-11. The extended range version could carry almost 

300 passengers at distances over 8,200 miles, which made it one of the farther flying aircraft in history and ideal for expanding Pacific air routes.112

Despite its initial reluctance, Boeing justified the incorporation of 

winglets into the new 747-400 in 1985, making it the first large U.S. commercial transport to incorporate winglets. The technology increased 

110. KC-1935 Winglet Program Review: Proceedings of a Symposium Held At Dryden Flight 

Research Center, Sept. 16, 1981, NASA CP-2211 (Washington, DC: NASA, 1982), pp. 1, 

11–12; Chambers, Concept to Reality, pp. 38–39. In the end, the Air Force chose not to equip its 

KC-135 aerial tankers with winglets, opting for new engines instead.

111. Staff of Douglas Aircraft Company, DC-10 Winglet Flight Evaluation, NASA CR-3704 (June 

1983), pp. v, 115–116; Chambers, Concept to Reality, pp. 38, 39, 41, 43.

112. Chambers, Concept to Reality, p. 43; “Winglets for the Airlines,” n.p., n.d.; The Boeing 

Company, “Commercial Airplanes: MD-11 Family,” 2009, http://www.boeing.com/commercial/

md-11family/index.html (Accessed Mar. 1, 2009).
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2 the new airplane’s range by 3 percent, enabling it to fly farther and 

with more passengers or cargo. The Boeing winglet differed from the 

McDonnell-Douglas design in that it did not have a smaller fin below 

the wingtip. Boeing engineers felt the low orientation of the 747 wing, 

combined with the practical presence of airport ground-handling equipment, made the deletion necessary.113

It was clear that Boeing included winglets on the 747-400 for 

improved performance. Boeing also offered winglets as a customer 

option for its 737 series aircraft and adopted blended winglets for its 

737 and the 737-derivative Business Jet provided by Aviation Partners, 

Inc., of Seattle in the early 1990s. The specialty manufacturer introduced 

its proprietary “blended winglet” technology—the winglet is joined to 

the wing via a characteristic curve—and started retrofitting them to 

Gulfstream II business jets. The performance accessory increased fuel 

efficiency by 7 percent. That work lead to commercial airliner accounts. 

Winglets for the 737 offered fuel savings and reduced noise pollution. The 

relationship with Boeing lead to a joint venture called Aviation Partners 

Boeing, which now produces winglets for the 757 and 767 airliners. By 

2003, there were over 2,500 Boeing jets flying with blended winglets. 

The going rate for a set of the 8-foot winglets in 2006 was $600,000.114

Whitcomb’s winglets found use on transport, airliner, and business 

jet applications in the United States and Europe. Airbus installed them 

on production A319, A320, A330, and A340 airliners. It was apparent that 

regardless of national origin, airlines chose a pair of winglets for their 

aircraft because they offered a savings of 5 percent in fuel costs. Rather 

than fly at the higher speeds made possible by winglets, most airline 

operators simply cruised at their pre-winglet speeds to save on fuel.115

Whitcomb’s aerodynamic winglets also found a place outside aeronautics, as they met the hydrodynamic needs of the international yacht 

racing community. In preparation for the America’s Cup yacht race in 

1983, Australian entrepreneur Alan Bond embraced Whitcomb’s work on 

113. Chambers, Concept to Reality, pp. 38, 43.

114. Aviation Partners Boeing, “Winglets,” 2006, http://www.aviationpartnersboeing.com (Accessed 

Mar. 27, 2009); Stephen O. Andersen and Durwood Zaelke, Industry Genius: Inventions and People 

Protecting the Climate and Fragile Ozone Layer (Sheffield, UK: Greenleaf Publishing, 2003), pp. 

32–52; Aviation Partners Boeing, “Winglets Save Airlines Money: An Interview with Joe Clark and 

Jason Paur,” 2006, http://www.aviationpartnersboeing.com/interview.html (Mar. 27, 2009).

115. Welch, “Whitcomb,” p. 5; Larson, “Winglets.”
Case 2 | Richard Whitcomb and the Quest for Aerodynamic Efficiency

125

2 spiraling vortex drag and believed it could be applied to racing yachts. He 

assembled an international team that designed a winged keel, essentially 

a winglet tacked onto the bottom of the keel, for Australia II. Stunned 

by Australia II’s upsetting the American 130-year winning streak, the 

international yachting community heralded the innovation as the key 

to winning the race. Bond argued that the 1983 America’s Cup race was 

instrumental to the airline industry’s adoption of the winglet and erroneously believed that McDonnell-Douglas engineers began experimenting with winglets during the summer of 1984.116

Of the three triumphant innovations pioneered by Whitcomb, the 

area rule fuselage, the supercritical wing, and the winglet, perhaps it 

is the last that is the most easily recognizable for everyday air travelers and aviation observers. Engineer and historian Joseph R. Chambers 

remarked that: “no single NASA concept has seen such widespread use 

on an international level as Whitcomb’s winglets.” The application to 

commercial, military, and general aviation aircraft continues.117

Whitcomb and History

Aircraft manufacturers tried repeatedly to lure Whitcomb away from 

NASA Langley with the promise of a substantial salary. At the height of 

his success during the supercritical wing program, Whitcomb remarked: 

“What you have here is what most researchers like—independence. In 

private industry, there is very little chance to think ahead. You have to 

worry about getting that contract in 5 or 6 months.”118 Whitcomb’s independent streak was key to his and the Agency’s success. His relationship 

with his immediate boss, Laurence K. Loftin, the Chief of Aerodynamic 

Research at Langley, facilitated that autonomy until the late 1970s. 

When ordered to test a laminar flow concept that he felt was impractical in the 8-foot TPT, which was widely known as “Whitcomb’s tunnel,” 

he retired as head of the Transonic Aerodynamics Branch in February 

1980. He had worked in that organization since coming to Hampton 

from Worcester 37 years earlier, in 1943.119

Whitcomb’s resignation was partly due to the outside threat to his 

independence, but it was also an expression of his practical belief that 

116. David Devoss, “The Race to Recover the Cup,” Los Angeles Times, Aug. 31, 1986, p. X9.

117. Chambers, Concept to Reality, p. 44.

118. Grubisich, “Fuel-Saver in Wings.”

119. Bert, “Winged Victory”; Welch, “Whitcomb,” p. 4.
NASA’s Contributions to Aeronautics

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2

his work in aeronautics was finished. He was an individual in touch 

with major national challenges and having the willingness and ability to 

devise solutions to help. When he made the famous quote “We’ve done 

all the easy things—let’s do the hard [emphasis Whitcomb’s] ones,” he 

made the simple statement that his purpose was to make a difference.120

In the early days of his career, it was national security, when an innovation such as the area rule was a crucial element of the Cold War tensions between the United States and the Soviet Union. The supercritical 

wing and winglets were Whitcomb’s expression of making commercial 

aviation and, by extension, NASA, viable in an environment shaped by 

world fuel shortages and a new search for economy in aviation. He was 

a lifelong workaholic bachelor almost singularly dedicated to subsonic 

aerodynamics. While Whitcomb exhibited a reserved personality outside 

the laboratory, it was in the wind tunnel laboratory that he was unrestrained in his pursuit of solutions that resulted from his highly intuitive and individualistic research methods.

With his major work accomplished, Whitcomb remained at Langley 

as a part-time and unpaid distinguished research associate until 1991. 

With over 30 published technical papers, numerous formal presentations, and his teaching position in the Langley graduate program, he was 

a valuable resource for consultation and discussion at Langley’s numerous technical symposiums. In his personal life, Whitcomb continued his 

involvement in community arts in Hampton and pursued a new quest: 

an alternative source of energy to displace fossil fuels.121

Whitcomb’s legacy is found in the airliners, transports, business jets, 

and military aircraft flying today that rely upon the area rule fuselage, 

supercritical wings, and winglets for improved efficiency. The fastest, 

highest-flying, and most lethal example is the U.S. Air Force’s Lockheed 

Martin F-22 Raptor multirole air superiority fighter. Known widely as 

the 21st Century Fighter, the F-22 is capable of Mach 2 and features an 

area rule fuselage for sustained supersonic cruise, or supercruise, performance and a supercritical wing. The Raptor was an outgrowth of the 

Advanced Tactical Fighter (ATF) program that ran from 1986 to 1991. 

Lockheed designers benefited greatly from NASA work in fly-by-wire 

120. Hallion, On the Frontier, p. 202.

121. Bert, “Winged Victory”; NASA History Office, “Richard T. Whitcomb,” 2008, http://history.

nasa.gov/naca/bio.html (Accessed Feb. 27, 2009); “Richard Travis Whitcomb: Distinguished 

Research Associate,” NASA Langley Research Center, Apr. 1983.
Case 2 | Richard Whitcomb and the Quest for Aerodynamic Efficiency

127

2 control, composite materials, and stealth design to meet the mission of 

the new aircraft. The Raptor made its first flight in 1997, and production aircraft reached Air Force units beginning in 2005.122

Whitcomb’s ideal transonic transport also included an area rule 

fuselage, but because most transports are truly subsonic, there is no 

need for that design feature for today’s aircraft.123 The Air Force’s C-17 

Globemaster III transport is the most illustrative example. In the early 

1990s, McDonnell-Douglas used the knowledge generated with the YC-15 

to develop a system of new innovations—supercritical airfoils, winglets, 

advanced structures and materials, and four monstrous high-bypass turbofan engines—that resulted in the award of the 1994 Collier Trophy. 

After becoming operational in 1995, the C-17 is a crucial element in the 

Air Force’s global operations as a heavy-lift, air-refuelable cargo transport.124 After the C-17 program, McDonnell-Douglas, which was absorbed 

into the Boeing Company in 1997, combined NASA-derived advanced 

blended wing body configurations with advanced supercritical airfoils 

and winglets with rudder control surfaces in the 1990s.125

Unfortunately, Whitcomb’s tools are in danger of disappearing. Both 

the 8-foot HST and the 8-foot TPT are located beside each other on 

Langley’s East Side, situated between Langley Air Force Base and the 

Back River. The National Register of Historic Places designated the 

Collier-winning 8-foot HST a national historic landmark in October 

1985.126 Shortly after Whitcomb’s discovery of the area rule, the NACA 

suspended active operations at the tunnel in 1956. As of 2006, the 

Historic Landmarks program designated it as “threatened,” and its future 

122. James Blackwell, “Influence on Today’s Aircraft,” in Proceedings of the F-8 Digital Fly-ByWire and Supercritical Wing First Flight’s 20th Anniversary, May 27, 1992, pp. 96–97, 100; 

U.S. Air Force, “F-22 Raptor,” Mar. 2009, http://www.af.mil/information/factsheets/factsheet.

asp?fsID=199 (May 21, 2009).

123. Hallion, On the Frontier, p. 206.

124. Langley Research Center, “NASA Contributions to the C-17 Globemaster III,” FS-1996-05-06-

LaRC (May 1996): p. 2.

125. Chambers, Innovation in Flight, p. 79.

126. The National Register also recognized two other important Langley wind tunnels: the 

Variable-Density Tunnel of 1922 and the Full-Scale Tunnel of 1931. National Park Service, “From 

Sand Dunes to Sonic Booms: List of Sites,” n.d., http://www.nps.gov/nr/travel/aviation/sitelist.

htm (Accessed Mar. 15, 2009); NASA, “Langley Research Center National Historic Landmarks,” 

1992, http://www.nasa.gov/centers/langley/news/factsheets/Landmarks.html#8FT (Accessed 

Mar. 15, 2009).
NASA’s Contributions to Aeronautics

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2

The Boeing C-17 Globemaster III. U.S. Air Force.

disposition was unclear.127 The 8-foot TPT opened in 1953. He validated 

the area rule concept and conducted his supercritical wing and winglet research through the 1950s, 1960s, and 1970s in this tunnel, which 

was located right beside the old 8-foot HST. The tunnel ceased operations in 1996 and has been classified as “abandoned” by NASA.128 In 

the early 21st century, the need for space has overridden the historical 

importance of the tunnel, and it is slated for demolition.

Overall, Whitcomb and Langley shared the quest for aerodynamic efficiency, which became a legacy for both. Whitcomb flourished working in his tunnel, limited only by the wide boundaries of his 

intellect and enthusiasm. One observer considered him to be “flight 

127. National Park Service, “Eight-Foot High Speed Tunnel,” n.d., http://www.nps.gov/nr/

travel/aviation/8ft.htm (Accessed Mar. 5, 2009); National Park Service, “National Historic Landmarks Program: Eight-Foot High Speed Tunnel,” 2006, http://tps.cr.nps.gov/nhl/detail.cfm?

ResourceId=1916&ResourceType=Structure, (Accessed Mar. 5, 2009).

128. Welch, “Whitcomb,” p. 4; NASA, “Audit of Wind Tunnel Utilization,” 2003, oig.nasa.gov/

audits/reports/FY03/pdfs/ig-03-027.pdf (Accessed Mar. 17, 2009).
Case 2 | Richard Whitcomb and the Quest for Aerodynamic Efficiency

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2

A 3-percent scale model of the Boeing Blended Wing Body 450 passenger subsonic transport 

in the Langley 14 x 22 Subsonic Tunnel. NASA.

theory personified.”129 More importantly, Whitcomb was the ultimate 

personification of the importance of the NACA and NASA to American 

aeronautics during the second aeronautical revolution. The NACA and 

NASA hired great people, pure and simple, in the quest to serve American 

aeronautics. These bright minds made up a dynamic community that 

created innovations and ideas that were greater than the sum of their 

parts. Whitcomb, as one of those parts, fostered innovations that proved 

to be of longstanding value to aviation.

129. Welch, “Whitcomb,” p. 4.
NASA’s Contributions to Aeronautics

130

2 Recommended Additional Readings

Reports, Papers, Articles, and Presentations: 

John D. Anderson, Jr., “Richard Whitcomb and the Area Rule,” in U.S. 

Air Force: A Complete History, Dik A. Daso, ed. (New York: Hugh 

Lauter Levin Associates, 2006).

David A. Anderton, “NACA Formula Eases Supersonic Flight,” Aviation 

Week 63 (Sept. 12, 1955).

Clay Blair, Jr., “The Man Who Put the Squeeze on Aircraft Design,” Air 

Force Magazine 39 (1956).

Conventional Take-off and Landing (CTOL) Transport Technology 1978: 

Proceedings of a Conference Held at Langley Research Center, Hampton, 

VA, Feb. 28–Mar. 3, 1978, NASA CP-2036 (1978).

Douglas Aircraft Company, DC-10 Winglet Flight Evaluation, NASA 

CR-3704 (1983).

Stuart G. Flechner and Peter F. Jacobs, “Experimental Results of Winglets 

on First, Second, and Third Generation Jet Transports,” NASA 

TM-72674 (1978).

Stuart G. Flechner, Peter F. Jacobs, and Richard T. Whitcomb, “A High 

Subsonic Wind Tunnel Investigation of Winglets on a Representative 

Second-Generation Jet Transport Wing,” NASA TN-8264 (1976).

M.I. Goldhammer, “A Lifting Surface Theory for the Analysis of 

Nonplanar Lifting Systems,” AIAA Paper No. 76-16 (1976).

Charles D. Harris, “NASA Supercritical Airfoils—A Matrix of FamilyRelated Airfoils,” NASA TP-2969 (1990).

KC-1935 Winglet Program Review: Proceedings of a Symposium Held at 

Dryden Flight Research Center, Sept. 16, 1981, NASA CP-2211 (1982).
Case 2 | Richard Whitcomb and the Quest for Aerodynamic Efficiency

131

2 John E. Lamar, “A Vortex-Lattice Method for the Mean Camber Shapes 

of Trimmed Noncoplanar Platforms with Minimum Vortex Drag,” 

NASA TN-D-8090 (1976).

Langley Research Center, “NASA Contributions to the C-17 Globemaster 

III,” FS-1996-05-06-LaRC (1996).

Lawrence C. Montoya and Richard D. Banner, “F-8 Supercritical Wing 

Flight Pressure, Boundary-Layer, and Wake Measurements and 

Comparisons with Wind Tunnel Data,” NASA TM-X-3544 (1977).

Proceedings of the F-8 Digital Fly-By-Wire and Supercritical Wing First 

Flight’s 20th Anniversary, May 27, 1992, NASA CP-3256 (1996).

Jon S. Pyle and Louis L. Steers, “Flight-Determined Lift and Drag 

Characteristics of an F-8 Airplane Modified with a Supercritical 

Wing with Comparisons to Wind Tunnel Results,” NASA TM-X3250 (1975).

Supercritical Wing Technology—A Progress Report on Flight Evaluations, 

NASA SP-301 (1972).

Richard T. Whitcomb, “A Design Approach and Selected Wind-Tunnel 

Results at High Subsonic Speeds for Wing-Tip Mounted Winglets,” 

NASA TN-D-8260 (1976).

Richard T. Whitcomb, “A Study of the Zero-Lift Drag-Rise Characteristics 

of Wing-Body Combinations Near the Speed of Sound,” NACA 

RM-L52H08 (1952).

Richard T. Whitcomb, “A Study of the Zero-Lift Drag-Rise Characteristics 

of Wing-Body Combinations Near the Speed of Sound,” NACA 

TR-1273 (1956).

Richard T. Whitcomb, “Research Associated with the Langley 8-Foot 

Tunnels Branch: Lecture at Ames Research Center, October 21, 1970,” 

NASA TM-108686 (1970).
NASA’s Contributions to Aeronautics

132

2 Richard T. Whitcomb, “Some Considerations Regarding the Application 

of the Supersonic Area Rule to the Design of Airplane Fuselages,” 

NACA RM-L56E23a (1956).

Richard T. Whitcomb and Larry L. Clark, “An Airfoil Shape for Efficient 

Flight at Supercritical Mach Numbers,” NASA TM-X-1109 (1965).

Richard T. Whitcomb and Thomas L. Fischetti, “Development of a 

Supersonic Area Rule and an Application to the Design of a WingBody Combination Having High Lift-to-Drag Ratios,” NACA 

RM-L53H31A (1953).

Richard T. Whitcomb and Thomas C. Kelly, “A Study of the Flow Over 

a 45-degree Sweptback Wing-Fuselage Combination at Transonic 

Mach Numbers,” NACA RM-L52DO1 (1952).

Books and Monographs:

John D. Anderson, Jr., A History of Aerodynamics and its Impact on Flying 

Machines (New York: Cambridge University Press, 1997).

Donald D. Baals and William R. Corliss, Wind Tunnels of NASA 

(Washington, DC: NASA, 1981).

John Becker, The High Speed Frontier: Case Histories of Four NACA 

Programs 1920–1950, NASA SP-445 (Washington, DC: NASA, 1980).

Joseph R. Chambers, Concept to Reality: Contributions of the Langley 

Research Center to U.S. Civil Aircraft of the 1990s (Washington, DC: 

NASA, 2003).

Joseph R. Chambers, Innovation in Flight: Research of the NASA Langley 

Research Center on Revolutionary Advanced Concepts for Aeronautics, 

NASA SP-2005-4539 (Washington, DC: NASA, 2005).

Eric M. Conway, High Speed Dreams: NASA and the Technopolitics of 

Supersonic Transportation, 1945–1999 (Baltimore: Johns Hopkins 

University Press, 2005).
Case 2 | Richard Whitcomb and the Quest for Aerodynamic Efficiency

133

2 Eugene S. Ferguson, Engineering and the Mind’s Eye (Boston: MIT Press, 

1994).

Michael Gorn, Expanding the Envelope: Flight Research at NACA and 

NASA (Lexington: University Press of Kentucky, 2001).

Richard P. Hallion, Designers and Test Pilots (Alexandria, VA: Time-Life 

Books, 1983).

Richard P. Hallion, On the Frontier: Flight Research at Dryden, 1946–

1981 (Washington, DC: NASA, 1984).

James R. Hansen, Engineer in Charge: A History of the Langley Aeronautical 

Laboratory, 1917–1958, NASA SP-4305 (Washington, DC: NASA, 

1987).

Pam E. Mack, ed., From Engineering Science to Big Science: The NACA 

and NASA Collier Trophy Research Project Winners (Washington, DC: 

NASA, 1998).

James Schultz, Crafting Flight: Aircraft Pioneers and the Contributions of 

the Men and Women of NASA Langley Research Center (Washington, 

DC: NASA, 2003).
134

The Bell XV-1 Convertiplane. NASA.
135

3

NACA–NASA and the 

Rotary Wing Revolution

John F. Ward

CASE

3

The NACA and NASA have always had a strong interest in promoting 

Vertical/Short Take-Off and Landing (V/STOL) flight, particularly those systems that make use of rotary wings: helicopters, autogiros, and tilt rotors. 

New structural materials, advanced propulsion concepts, and the advent of 

fly-by-wire technology influenced emergent rotary wing technology. Work by 

researchers in various Centers, often in partnership with the military, enabled 

the United States to achieve dominance in the design and development of 

advanced military and civilian rotary wing aircraft systems, and continues 

to address important developments in this field.

I

F WORLD WAR I LAUNCHED THE FIXED WING AIRCRAFT INDUSTRY, the 

Second World War triggered the rotary wing revolution and sowed the 

seeds of the modern American helicopter industry. The interwar years 

had witnessed the development of the autogiro, an important short takeoff 

and landing (STOL) predecessor to the helicopter, but one incapable of true 

vertical flight, or hovering in flight. The rudimentary helicopter appeared 

at the end of the interwar era, both in Europe and America. In the United 

States, the Sikorsky R-4 was the first and only production helicopter used 

in United States’ military operations during the Second World War. R-4 

production started in 1943 as a direct outgrowth of the predecessor, VS-300, 

the first practical American helicopter, which Igor Sikorsky had refined by 

the end of 1942. That same year, the American Helicopter Society (AHS) 

was chartered as a professional engineering society representing the rotary 

wing industry. Also in 1943, the Civil Aeronautics Administration (CAA), 

forerunner of the Federal Aviation Administration (FAA), issued Aircraft 

Engineering Division Report No. 32, “Proposed Rotorcraft Airworthiness.” 

Thus was America’s rotary wing industry birthed.1

1. Russell E. Lee, “Famous Firsts in Helicopter History,” in Walter J. Boyne and Donald S. Lopez, eds., 

Vertical Flight: The Age of the Helicopter (Washington: Smithsonian Institution Press, 1984), p. 248; 

Don Fertman, “The Helicopter History of Sikorsky Aircraft,” Vertiflite, vol. 30, no. 4 (May/June 1984), 

p. 16; Mike Debraggio, “The American Helicopter Society—A Leader for 40 Years,” Vertiflite, vol. 

30, no. 4 (May/June 1984), p. 56.
NASA’s Contributions to Aeronautics

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3

Igor Sikorsky flying the experimental VS-300. Sikorsky.

As a result of the industry’s growth spurred by continued military 

demand during the Korean war and the Vietnam conflict, interest in helicopters grew almost exponentially. As a result of the boost in demand 

for helicopters, Sikorsky Aircraft, Bell Helicopter, Piasecki Helicopter 

(which evolved into Vertol Aircraft Corporation in 1956, becoming the 

Vertol Division of the Boeing Company in 1960), Kaman Aircraft, Hughes 

Helicopter, and Hiller Aircraft entered design evaluations and prototype 

production contracts with the Department of Defense. Over the past 65 

years, the rotary wing industry has become a vital sector of the world aviation system. Types of private, commercial and military utilization abound 

using aircraft designs of increasing capability, efficiency, reliability, and 

safety. Helicopters have now been joined by the military V-22, the first 

operational tilt rotor, and emerging rotary wing unmanned aerial vehicles 

(UAV), with both successful rotary wing concepts having potential civil 

applications. Over the past 78 years, the National Advisory Committee for 

Aeronautics (NACA) and its successor, the National Aeronautics and Space 

Administration (NASA), have made significant research and technology 

contributions to the rotary wing revolution, as evidenced by numerous 

technical publications on rotary wing research testing, database analysis, 

and theoretical developments published since the 1930s. These technical 
Case 3 | NACA–NASA and the Rotary Wing Revolution

137

3

resources have made significant contributions to the Nation’s aircraft 

industry, military services, and private and commercial enterprises.

The Research Culture

As part of the broad scope of aeronautics research, the rotary wing efforts 

spanned the full range of research activity, including theoretical study, 

wind tunnel testing, and ground-based simulation. Flight-test NACA 

rotary wing research began in the early 1920s with exploratory wind tunnel tests of simple rotor models as the precursor to the basic research 

undertaken in the 1930s. The Langley Memorial Aeronautical Laboratory, 

established at Hampton, VA, in 1917, purchased a Pitcairn PCA-2 autogiro in 1931 for research use.2

 The National Advisory Committee for 

Aeronautics had been formed in 1915 to “supervise and direct scientific study the problems of flight, with a view to their practical solution.” 

Rotary wing research at Langley proceeded under the direction of the 

Committee with annual inspection meetings by the full Committee to 

review aeronautical research progress. In the early 1940s, the Ames 

Aeronautical Laboratory, now known as the Ames Research Center, 

opened for research at Moffett Field in Sunnyvale, CA. Soon after, the 

Aircraft Engine Research Laboratory, known for many years as the Lewis 

Research Center and now known as the Glenn Research Center, opened 

in Cleveland, OH. Each NACA Center had unique facilities that accommodated rotary wing research needs. Langley Research Center played a 

major role in NACA–NASA rotary wing research until 1976, when Ames 

Research Center was assigned the lead role.

The rotary wing research is carried out by a staff of research engineers, scientists, technical support specialists, senior management, and 

administrative personnel. The rotary wing research staff draws on the 

expertise of the technical discipline organizations in areas such as aerodynamics, structures and materials, propulsion, dynamics, acoustics, 

and human factors. Key support functions include such activities as 

test apparatus design and fabrication, instrumentation research and 

development (R&D), and research computation support. The constant 

instrumentation challenge is to adapt the latest technology available to 

acquiring reliable research data. Over the years, the related challenge 

for computation tasks is to perform data reduction and analysis for the 

2. F.B. Gustafson, “A History of NACA Research on Rotating-Wing Aircraft,” Journal of the American Helicopter Society, vol. 1, no. 1 (Jan. 1956), p. 16.
NASA’s Contributions to Aeronautics

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3

increasing sophistication and scope of theoretical investigations and 

test projects. In the NACA environment, the word “computers” actually referred to a large cadre of female mathematicians. They managed 

the test measurement recordings, extracted the raw data, analyzed the 

data using desktop electromechanical calculators, and hand-plotted the 

results. The NASA era transformed this work from a tedious enterprise 

into managing the application of the ever-increasing power of modern 

electronic data recording and computing systems. 

The dissemination of the rotary wing research results, which form 

the basis of NACA–NASA contributions over the years, takes a number 

of forms. The effectiveness of the contributions depends on making 

the research results and staff expertise readily available to the Nation’s 

Government and industry users. The primary method has traditionally been the formal publication of technical reports, studies, and compilations that are available for exploitation and use by practitioners. 

Another method that fosters immediate dialogue with research peers 

and potential users is the presentation of technical papers at conferences and technical meetings. These papers are published in the conference proceedings and are frequently selected for broader publication 

as papers or journal articles by technical societies such as the Society of 

Automotive Engineers (SAE)–Aerospace and the American Institute of 

Aeronautics and Astronautics (AIAA). Since 1945, NACA–NASA rotary 

wing research results have been regularly published in the Proceedings 

of the American Helicopter Society Annual Forum and the Journal of 

the AHS. During this time, 30 honorary awards have been presented to 

NACA and NASA researchers at the Annual Forum Honors Night ceremonies. These awards were given to individual researchers and to technical teams for significant contributions to the advancement of rotary 

wing technology. 

Over the years, the technical expertise of the personnel conducting 

the ongoing rotary wing research at NACA–NASA has represented a valuable national resource at the disposal of other Government organizations 

and industry. Until the Second World War, small groups of rotary wing 

specialists were the prime source of long-term, fundamental research. In 

the late 1940s, the United States helicopter industry emerged and established technical teams focused on more near-term research in support 

of their design departments. In turn, the military recognized the need 

to build an in-house research and development capability to guide their 

major investments in new rotary wing fleets. The Korean war marked 
Case 3 | NACA–NASA and the Rotary Wing Revolution

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Case 3 | NACA–NASA and the Rotary Wing Revolution

the beginning of the U.S. Army’s long-term commitment to the utilization of rotary wing aircraft. In 1962, Gen. Hamilton H. Howze, the first 

Director of Army Aviation, convened the U.S. Army Tactical Mobility 

Requirements Board (Howze Board).3 This milestone launched the emergence of the Air Mobile Airborne Division concept and thereby the steady 

growth in U.S. military helicopter R&D and production. The working 

relationship among Government agencies and industry R&D organizations has been close. In particular, the availability of unique facilities 

and the existence of a pool of experienced rotary wing researchers at 

NASA led to the United States Army’s establishing a “special relationship” with NASA and an initial research presence at the Ames Research 

Center in 1965. This was followed by the creation of co-located and integrated research organizations at the Ames, Langley, and Glenn Research 

Centers in the early 1970s. The Army organizations were staffed by specialists in key disciplines such as unsteady aerodynamics, aeroelasticity, acoustics, flight mechanics, and advanced design. In addition, Army 

civilian and military engineering and support personnel were assigned 

to work full time in appropriate NASA research facilities and theoretical 

analysis groups. These assignments included placing active duty military test pilots in the NASA flight research organizations. Over the long 

term, this teaming arrangement facilitated significant research activity. 

In addition to Research and Technology Base projects, it made it possible to perform major jointly funded and managed rotary wing Systems 

Technology and Experimental Aircraft programs. The United States Army 

partnership was augmented by other research teaming agreements with 

the United States Navy, FAA, the Defense Advanced Research Projects 

Agency (DARPA), academia, and industry.

NACA 1930–1958: Establishing Fundamentals

While the helicopter industry did not emerge until the 1950s, the NACA 

was engaged in significant rotary wing research starting in the 1930s at 

the Langley Memorial Aeronautical Laboratory (LMAL), now the NASA 

3. Edgar C. Wood, “The Army Helicopter, Past, Present and Future,” Journal of the American 

Helicopter Society, vol. 1, no. 1 (Jan. 1956), pp 87–92; Lt. Gen. John J. Tolson, Airmobility, 

1961-1971, a volume in the U.S. Army Vietnam Studies series (Washington, DC: Army, 1973), pp. 

16–24; and J. A. Stockfisch, The 1962 Howze Board and Army Combat Developments, Monograph Report MR-435-A (Santa Monica: The RAND Corporation, 1994).
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Pitcairn PCA-2 Autogiro. NASA.

Langley Research Center.4 The early contributions were the result of 

studies of the autogiro. The focus was on documenting flight characteristics, performance prediction methods, comparison of flight-test and 

wind tunnel test results, and theoretical predictions. In addition, fundamental operating problems definition and potential solutions were 

addressed. In 1931, the NACA made its first direct purchase of a rotary 

wing aircraft for flight test investigations, a Pitcairn PCA-2 autogiro. 

(With few exceptions, future test aircraft were acquired as short-term 

loan or long-term bailment from the military aviation departments.) 

The Pitcairn was used over the next 5 years in flight-testing and tests of 

the rotor in the Langley 30- by 60-foot Full-Scale Tunnel. Formal publications of greatest permanent value received “report” status, and the 

Pitcairn’s first study, NACA Technical Report 434, was the first authoritative information on autogiro performance and rotor behavior.5

4. This case study has drawn upon two major sources covering the period 1930 through 1984 published in Vertiflite, the quarterly magazine of the American Helicopter Society: Frederic B. Gustafson, 

“History of NACA/NASA Rotating-Wing Aircraft Research, 1915–1970,” Vertiflite, Reprint VF-70, 

(Apr. 1971), pp. 1–27; and John Ward, “An Updated History of NACA/NASA Rotary-Wing Aircraft Research 1915-1984,” Vertiflite, vol. 30, No. 4 (May/June 1984), pp. 108–117. The author 

(who wrote the second of those two) has extended the coverage beyond the original 1984 end date. 

5. J.B. Wheatley, “Lift and Drag Characteristics and Gliding Performance of an Autogiro as Determined In Flight,” NACA Report No. 434 (1932).
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The mid-1930s brought visiting autogiros and manufacturing personnel to Langley Research Center. In addition, analytical and wind 

tunnel work was carried out on the “Gyroplane,” which incorporated 

a rotor without the usual flapping or lead-lag hinges at the blade root. 

This was the first systematic research documented and published for 

what is now called the “rigid” or “hingeless” rotor. This work was the 

forerunner of the hingeless rotor’s reappearance in the 1950s and 1960s 

with extensive R&D effort by industry and Government. The NACA’s 

early experience with the Gyroplane rotor suggested that “designing 

toward flexibility rather than toward rigidity would lead to success.” In 

the 1950s, the NACA began to encourage this design approach to those 

expressing interest in hingeless rotors.

While the NACA worked to provide the fundamentals of rotary 

wing aerodynamics, the autogiro industry experienced major changes. 

Approximately 100 autogiros were built in the United States and hundreds 

more worldwide. Problems in smaller autogiros were readily addressed, 

but those in larger sizes persisted. They included stick vibration, heavy control forces, vertical bouncing, and destructive out-of-pattern blade behavior known as ground resonance. Private and commercial use underwent 

a discouraging stage. However, military interest grew in autogiro utility 

capabilities for safe flight at low airspeed. In an early example of cooperation with the military, the NACA’s research effort was linked to the needs 

of the Army Air Corps (AAC), predecessor of the Army Air Forces (AAF). 

In quick succession, Langley Laboratory conducted flight and/or wind tunnel tests on a series of Kellett Autogiros, including the KD-1, YG-1, YG-1A, 

YG-1B, and the Pitcairn YG-2. The NACA provided control force and performance measurements, and pilot assessments of the YG-1. In addition, 

recommendations were provided on maneuver limitations and redesign 

for better military serviceability. This led to the NACA providing recommendations and pilot training to enable the Army Air Corps to begin conducting its own rotary wing aircraft experimental and acceptance testing. 

In the fall of 1938, international events required that the NACA’s emphasis turn to preparedness. The United States required fighters and bombers with superior performance. In the next few years, experimental rotary 

wing research declined, but important basic groundwork was conducted. 

Limited effort began on the potentially catastrophic phenomena of ground 

resonance or coupled rotor-fuselage mechanical instability. Photographs 

were taken of the rotor-blade out-of-pattern behavior by mounting a camera high on the Langley Field balloon (airship) hangar while an autogiro 
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was operated on the ground. Exploratory flight tests were done using a hubmounted camera. In these tests blade motion studies were conducted to 

document the pattern of rotor-blade stalling behavior. In the closing years 

of the 1930s, analytical progress was also made in the creation of a new 

theory of rotor aerodynamics that became a classic reference and formed 

the basis for NACA helicopter experimentation in the 1940s.6 In these years, 

the top leadership of the NACA engaged in visible participation in the formal dialogue with the rotating wing community. In 1938, Dr. George W. 

Lewis, the NACA Headquarters Director of Aeronautical Research, served 

as Chairman of the Research Programs session of the pioneering RotatingWing Aircraft Meeting at the Franklin Institute in Philadelphia. In 1939, 

Dr. H.J.E. Reid, Director of Langley Laboratory, the NACA’s only laboratory, served as Chairman of the session in Dr. Lewis’s absence.7

The early 1940s continued a period of only modest NACA effort on 

rotary wing research. However, military interest in the helicopter as a 

new operational asset started to grow with attention to the need for special missions such as submarine warfare and the rescue of downed pilots. 

As noted in the introduction to this chapter, the need was met by the 

Sikorsky R-4 (YR-4B), which was the only production helicopter used in 

United States military operations during the Second World War. The R-4 

production started in 1943 as a direct outgrowth of the Sikorsky VS-300. 

As the helicopter industry emerged, the NACA rotary wing community 

enjoyed a productive contact through the interface provided by the NACA 

Rotating Wing (later renamed Helicopter) Subcommittee. It was in these 

technical subcommittees that experts from Government, industry, and 

academia spelled out the research needs and set priorities to be addressed 

by the NACA rotary wing research specialists. The NACA committee and

subcommittee roles were marked by a strong supervisory tone, as called 

for in the NACA charter. The members lent a definite direction to NACA 

research based on their technical needs. They also attended annual inspection tours of the three NACA Centers to review the progress on the assigned 

6. J.B. Wheatley, “A Aerodynamic Analysis of the Autogiro Rotor With Comparison Between Calculated and Experimental Results,” NACA Report No. 487 (1934).

7. Anon., “Proceedings of Rotating-Wing Aircraft Meeting of the Franklin Institute, Philadelphia, 

Pennsylvania, Oct. 20–29, 1938,” Philadelphia Section, Institute of the Aeronautical Sciences 

(IAS); Anon., “Proceedings of the Second Annual Rotating-Wing Aircraft Meeting of the Franklin 

Institute, Philadelphia, Pennsylvania, Nov. 30–Dec. 1, 1939,” Philadelphia Section, Institute of the 

Aeronautical Sciences (IAS).
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Sikorsky YR-4B tested in the Langley 30 x 60 ft. wind tunnel. NASA.

research efforts. In the NASA era, the committees and subcommittees 

evolved into a more advisory function: commenting upon and ranking 

the merits of projects proposed by the research teams.

NACA Report 716, published in 1941, constituted a particularly significant contribution to helicopter theory, for it provided simplified methods and charts for determining rotor power required and blade motion.8

For the first time, design studies could be performed to begin to assess 

the impacts of blade-section stalling and tip-region compressibility effects. 

Theoretical work continued throughout the 1940s to extend the simple 

theory into the region of more extreme operating conditions. Progress 

began to be made in unraveling the influence of airfoil selection, high bladesection angles of attack, and high tip Mach numbers. The maximum 

Mach number excursion occurred as the tip passed through the region 

where the rotor rotational velocity and the forward airspeed combined. 

Flight research was begun with the first production helicopter, the 

Sikorsky YR-4B. This work produced a series of comparisons of flighttest results with theoretical predictions utilizing the new methodology 

8. F.J. Bailey, Jr., “A Simplified Theoretical Method of Determining the Characteristics of a Lifting 

Rotor in Forward Flight,” NACA Report No. 716 (1941).
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for rotor performance and blade motion. The results of the comparisons validated the basic theoretical methods for hover and forward flight 

in the range of practical steady-state operating conditions. The YR-4B 

helicopter was also tested in the Langley 30 by 60 tunnel.

This facilitated rotor-off testing to provide fuselage-only lift and 

drag measurements. This in turn enabled the flight measurements to 

be adjusted for direct comparison with rotor theory. 

With research progressing in flight test, wind tunnel test and theory 

development, a growing, well-documented open rotary wing database 

was swiftly established. At the request of industry, Langley airfoil specialists designed and tested airfoils specifically tailored to operating in the 

challenging unsteady aerodynamic environment of the helicopter rotor. 

However, the state-of-the-art of airfoil development required that the airfoil be designed on the basis of a single, steady airflow condition. Selecting 

this artful compromise between rapid excursions into the high angle of 

attack stall regions and the zero-lift conditions was daunting.9 Database 

buildup also included the opportunity offered by the YR-4B 30x60 wind 

tunnel test setup. This provided the opportunity to document a database 

from hovering tests on six sets of rotor blades of varying construction and 

geometry. The testing included single, coaxial, and tandem rotor configurations. Basic single rotor investigations were conducted of rotor-blade pressure distribution, Mach number effects, and extreme operation conditions. 

In 1952, Alfred Gessow and Garry Myers published a comprehensive textbook for use by the growing helicopter industry.10 The authors’ 

training and experience had been gained at Langley Laboratory, and 

the experimental and theoretical work done by laboratory personnel 

over the previous 15 years (constituting over 70 published documents) 

served as the basis of the aerodynamic material developed in the book. 

The Gessow-Myers textbook remains to this day a classic introduction 

to helicopter design. 

Significant contributions were made in rotor dynamics. The principal contributions addressed the lurking problem of ground resonance, or 

self-excited mechanical instability—the coupling of in-plane rotor-blade 

9. F.B. Gustafson, “Effects on Helicopter Performance of Modifications in Profile-Drag Characteristics 

of Rotor-Blade Airfoil Sections,” NACA WR-L-26 [formerly NACA Advanced Confidential Report 

ACR L4H05] (1944).

10. Alfred Gessow and Garry C. Myers, Jr., Aerodynamics of the Helicopter (New York: The Macmillan Company, 1952; reissued by Frederick Ungar Publishing Co., 1967).
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oscillations with the rocking motion of the fuselage on its landing gear. 

First encountered in some autogiro designs, the potential for a catastrophic outcome also existed for the helicopter.11 Theory developed 

and disseminated by the NACA enabled the understanding and analysis of ground resonance. This capability was considered essential to the 

successful design, production, and general use of rotary wing aircraft. 

Langley pioneered the use of scaled models for the study of dynamic 

problems such as ground resonance, blade flutter, and control coupling.12

This contribution to the contemporary state-of-the-art was a forerunner 

of the all-encompassing development and use of mathematical modeling throughout the modern rotary wing technical community.

As the helicopter flight-testing experience evolved, the research pilots 

observed problems in holding to steady, precision flight to enable data 

recording. Frequent control input adjustments were required to prevent 

diverging into attitudes that were difficult to recover from. Investigation 

of these flying quality characteristics led to devising standard piloting 

techniques to produce research-quality data. Deliberate, sharp-step 

and pulse-control inputs were made, and the resulting aircraft pitch, 

roll, and yaw responses were recorded for a few seconds. Out of this 

work came the research specialties of rotary wing flying qualities, stability and control, and handling qualities. Standard criteria for defining required flying qualities specifications gradually emerged from the 

NACA flight research. The results of this work supported the development of Navy helicopter specifications in the early 1950s and eventually 

for all military helicopters in 1956. In 1957, research at the NACA Ames 

Research Center produced a systematic protocol for pilots to assess aircraft handling qualities.13 The importance of damping of angular velocity 

and control power, and their interrelation, was investigated in Langley 

flight-testing. The results provided the basis for a major portion of formal flying-qualities criteria.14 After modification in 1969 based on exten11. R.P. Coleman, “Theory of Self-Excited Mechanical Oscillations of Hinged Rotor Blades,” NACA 

WR-L-308 [formerly NACA Advanced Restricted Report 3G29] (1943). 

12. G.W. Brooks, “The Application of Models to Helicopter Vibration and Flutter Research,” Proceedings of the ninth annual forum of the American Helicopter Society (May 1953).

13. George E. Cooper, “Understanding and Interpreting Pilot Opinion,” Aeronautical Engineering 

Review, vol. 16, no. 3, (Mar. 1957), p. 47–51.

14. S. Salmirs and R.J. Tapscott, “The Effects of Various Combinations of Damping and Control 

Power on Helicopter Handling Qualities During Both Instrument and Visual Flight,” NASA TN-D58 (1959).
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sive study of in-flight and simulation tasks at Ames, the Cooper-Harper 

Handling Qualities Rating Scale was published. It remains the standard 

for evaluating aircraft flying qualities, including rotary wing vehicles.15

In the late 1950s, the Army expanded the use of helicopters. The 

rotary wing industry grew to the point that manufacturers’ engineering departments included research and development staff. In addition, 

the Army established an aviation laboratory (AVLABS), now known 

as the Aviation Applied Technology Directorate (AATD), at the Army 

Transportation School, Fort Eustis, VA. This organization was able to 

sponsor and publish research conducted by the manufacturers. Fort 

Eustis was situated within 25 miles of the NACA’s Langley Research 

Center in Hampton on the Virginia peninsula. A majority of the key 

AVLABS personnel were experienced NACA rotary wing researchers. 

As it turned out, this personnel relocation, amounting to an unplanned 

“contribution” of expertise to the Army, was the forerunner of significant, long-term, co-located laboratory teaming agreements between the 

Army and NASA. 

NASA 1958–1970: A Time of Transition

The transformation of the NACA into NASA in 1958 was marked by 

an inevitable subordination of the NACA’s aeronautical research charter to NASA’s mandated space mission work. The assigned aeronautics staff dropped over 80 percent, from 7,100 to 1,400, as the space 

program gained momentum in the early 1960s. In the new spacefocused environment, aeronautics needed to be product-oriented to 

attract budget allocation support. In these circumstances, helicopter research lost ground as the focus shifted to new nonrotor Vertical 

Take-Off and Landing (VTOL) and Short Take-Off and Landing aircraft. In many cases, the rotary wing work formed the base for VTOL 

investigations. In the case of NACA–NASA rotor-flow studies, experimental and theoretical studies on rotor-time-averaged inflow led to 

extensive work on establishing wind tunnel jet-boundary layer (wall 

interference) correction methodology for other VTOL, as well as rotorborne, lifting systems.16

15. G.E. Cooper and R.P. Harper, Jr., “The Use of Pilot Rating in the Evaluation of Aircraft Handling 

Qualities,” NASA TN-D-5153 (1969).

16. Harry H. Heyson and S. Katzoff, “Induced Velocities Near a Lifting Rotor with Nonuniform Disk 

Loading,” NACA Report 1319 (1957).
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In a sense, it became the U.S. Army’s turn to bolster NASA rotary 

wing endeavors in support of the Army’s need for continued helicopter development. In 1965, the Army was granted permission to reactivate, staff, and utilize the Ames 7- by 10-foot Tunnel No. 2. In addition, 

the Army provided personnel to assist Ames in carrying out projects of 

interest to the Army. A group of about 45 people was established by the 

Army and identified as the Army Aeronautical Activity at Ames (AAA–

A).17 In 1970, the working relationship between NASA and the Army was 

significantly enhanced. Co-located Army research organizations were 

established at Ames, Langley, and Lewis (now Glenn) Research Centers. 

They focused on the respective Center’s specialty of aeroflight dynamics, 

structures, and propulsion. This teaming laid the solid groundwork for 

major rotary wing programs that NASA and the Army jointly planned, 

executed, and funded in the 1970s and 1980s that influenced both military and civilian rotary wing aircraft development. 

One of the unique research facilities authorized in 1939 and operated by the NACA, and then NASA, was the 40- by 80-foot Full-Scale 

Tunnel at Ames. This research facility also provided the opportunity to 

work directly with industry on vehicle development programs. In the 

case of rotary wing aircraft, the tunnel was utilized for investigating new 

vehicle and rotor system concepts and for thoroughly documenting the 

basic aerodynamic behavior of prototype and production articles. By the 

1960s, numerous in-house and industry full-scale rotary wing hardware 

were tested. Examples include the Bell XV-1 “convertiplane” in 1953–

1954, followed by many other projects, including a modified production 

rotor incorporating leading edge camber and boundary-layer control; 

the Bell UH-1 “Huey” helicopter (tested to assist in the development of 

a high-performance flight-test helicopter); a folded rotor with test data 

obtained in start-stop and folding conditions at forward speeds; and a 

four-bladed rotor investigation with extensive rotor-blade pressure measurements taken as a followup to prior flight test measurements made 

at Langley Research Center.18

17. Edwin P. Hartman, Adventures in Research, A History of Ames Research Center,1940–1965, 

NASA SP-4302 (Washington, DC: NASA, 1970), p. 411.

18. William Warmbrodt, Charles Smith, and Wayne Johnson, “Rotorcraft Research Testing in the 

National Full-Scale Aerodynamics Complex at NASA Ames Research Center,” NASA TM-86687 

(May 1985); J. Sheiman and L.H. Ludi, “Qualitative Evaluation of Effect of Helicopter Rotor Blade 

Tip Vortex on Blade Airloads,” NASA TN-D-1637 (1963).
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The pressure-instrumented blade used in the latter tests had an 

extremely limited operating life of only 10 hours. This was because of 

the installation of nearly 50 miniature differential pressure transducers inside the rotor blade. This required that a total of almost 100 small 

holes be drilled in the upper and lower surface of the primary structure D-spar—normally an absolute “safety of flight” violation. The conservative 10-hour limit was based upon conservative crack-growth-rate 

limits determined from blade specimen cyclic load tests. The earlier flight 

test investigation of blade pressure distributions produced a very significant contribution as a primary database for the understanding of basic 

rotor unsteady aerodynamics. The tabulated pressure data provided time 

histories of individual differential pressures and simultaneous blade 

bending moments around the rotor azimuth in a wide assortment of 

steady and maneuvering flight conditions.19 This database became the 

standard experimental data reference source for advancing theoretical 

comparison work for many years. As an aside, in working with the original 

flight data to hand-digitize the detailed recordings of differential pressure 

time-history traces, it became possible, in time, to visually recognize the 

specific flight-test condition by the periodic pressure trace signature.20 It was 

possible to identify the rotor’s actual flight condition relative to the surrounding airmass. This still raises the question of the possibility of applying 

modern signal recognition technology to provide on-board safety-of-flight 

and noise abatement operating boundary displays for the pilot. 

Flying qualities flight investigations emphasized the importance 

of ample damping of angular velocity and of control power (rotorgenerated aircraft pitch and roll control moments) and their interaction. 

This work at Langley and similar work at Ames provided a significant portion of the helicopter flying qualities criteria. This early work was extended 

to the use of in-flight simulation using Langley’s YHC-1A tandem rotor 

helicopter with special onboard computing and recording equipment.21

19. James Sheiman, “A Tabulation of Helicopter Rotor-Blade Differential Pressures, Stresses, and 

Motions As Measured In Flight,” NASA TM-X-952 (1964).

20. John F. Ward, “Helicopter Rotor Periodic Differential Pressures and Structural Response Measured in Transient and Steady-State Maneuvers,” Journal of the American Helicopter Society, vol. 16, 

no. 1 (Jan. 1971).

21. F. Garren, J.R. Kelly, and R.W. Summer, “VTOL Flight Investigation to Develop a Decelerating 

Instrument Approach Capability,” Society of Automotive Engineers Paper No. 690693 (1969), 

presented at the Aeronautics and Space Engineering and Manufacturing Meeting, Los Angeles, CA, 

Oct. 6–10, 1969.
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Tilt rotor semi-span dynamic model in the Langley Transonic Dynamics Tunnel. NASA.

The flight operations of most interest were terminal area instrument 

flight on steep approaches to vertical touchdown landings. The results 

of this work were initially oriented to nonrotor VTOL operations, but 

the results were found to be equally applicable to helicopters. 

In the area of structural dynamics, investigations addressing the 

problems of aeroelastic stability of rotor-powered aircraft were conducted utilizing new analytical methods and experimental studies by 

Langley and Ames researchers. Emphasis was placed on tilt rotor and 

tilt propeller (i.e., tilt wing) aircraft concepts. Two-degree-of-freedom 

“air resonance” (akin to rotor-fuselage “ground resonance”) and proprotor/propeller whirl instability were among the problems investigated.22

Rotor-pylon-wing aeroelastic instability problems for tilt rotor designs 

were explored in the Ames 40 by 80 Full-Scale Tunnel in this period. 

The aeroelastic stability problems of the tilt rotor and tilt-stopped rotor 

designs were also investigated at model scale in the unique Freon atmosphere of the Langley Transonic Dynamics Tunnel, which provided fullscale Mach number and Reynolds number scaling.23 These research 

22. Wilmer H. Reed, III, “Review of Propeller-Rotor Whirl Flutter,” NASA TR-R-264 (1967).

23. William T. Yeager, Jr., and Raymond G. Kvaternik, “A Historical Overview of Aeroelasticity 

Branch and Transonic Dynamics Tunnel Contributions to Rotorcraft Technology and Development,” 

NASA TM-2001-211054 / ARL-TR-2564, (Aug. 2001).
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investigations resulted in significant contributions to the development 

of the validated design tools for advanced rotorcraft.

With the increased interest in hingeless rotor concepts, NASA 

obtained and quickly accomplished flight research with a copy of an 

experimental Bell Helicopter three-bladed hingeless rotor installed on 

an H-13 helicopter.24 Early experience with “rigid” rotors had led the 

NACA to encourage interest in exploring the possibilities of removing 

conventional blade-root hinges and substituting instead blade structural flexibility. Another manufacturer, Lockheed Aircraft, made a major 

commitment to the hingeless rotor concept coupled to a mast-mounted 

mechanical gyro introduced into the pitch control linkage.25 The root 

regions of the blades in this innovative design were “matched stiffness” 

or “soft in-plane,” which meant that the blade chord-wise, or horizontal, structural bending stiffness was matched to the flap-wise, or vertical, 

bending stiffness. Dynamic model tests of this concept were conducted in 

the Langley 30 by 60 Full-Scale Tunnel and in the Freon atmosphere of 

the Langley Transonic Dynamics Tunnel. The use of Freon gas facilitated 

the testing of the 10-foot-diameter rotor model at full-scale Reynolds 

number and Mach numbers. This work began the establishment of a 

documented database for hingeless rotor design. These dynamic model 

tests were part of a cooperative NASA–Army AVLABS program. 

To further explore the problems and practical means for realizing 

the potential of the hingeless rotor concept, Langley Research Center 

purchased the Lockheed XH-51N, a high-speed research helicopter. 

The flight investigationfocused on the tendency for hingeless rotors to 

encounter high in-plane blade loads in roll maneuvers, coupling between 

the response to longitudinal and lateral control input, ride quality, and 

pilot handling qualities. In general, it was demonstrated with the flight 

tests and model tests that the hingeless rotor system was different from 

the conventional hinged systems. Inherently, the hingeless designs produced increased control moments, quicker response to pilot input and 

superior handling qualities. It turned out that later rotor designs incorporating elastomeric bearings to replace conventional hinges could provide a practical option to some of the fully hingeless designs.

24. R.J. Huston, “An Exploratory Investigation of Factors Affecting the Handling Qualities of a 

Rudimentary Hingeless Rotor Helicopter,” NASA TN-D-3418 (May 1966).

25. I.H. Culver and J.E. Rhodes, “Structural Coupling in the Blades of a Rotating Wing Aircraft,” 

IAS Paper No. 62-33 (1962).
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NASA 1970–1990: Joint Program Momentum Peaks 

During the early 1970s, the Ames Flight Simulator for Advanced Aircraft 

(FSAA) became operational and the first tilt rotor simulations were successfully accomplished. By 1975, the Army decided to augment the rotary 

wing flight dynamics research at Ames as NASA initiated the fabrication of the Vertical Motion Simulator (VMS). This simulator, with very 

large vertical and horizontal motion capability, was a national asset well 

suited for rotary wing research.

At Langley, a major instrument flight rules (IFR) investigation was 

conducted under the VTOL Approach and Landing Technology (VALT) 

program. The VALT Boeing-Vertol CH-47 Chinook helicopter was the primary research vehicle for exploring the control/display/task relationships. 

In addition, the Sikorsky SH-3 Sea King helicopter was used as a testbed 

for exploring the merits and defining the electro-optical parameter requirements associated with advanced “real-world” display concepts. The objective was to identify systems that might be capable of providing a pilot an 

“out-the-window display” during IFR flight conditions through the use of 

fog-cutting sensors or advanced computer-generated visual situation displays. The VALT CH-47 flights were conducted at the Wallops Flight Center, 

where the NASA Aeronautical Research Radar Complex provided omnidirectional tracking coverage. This facility permitted the investigation of 

a wide variety of approach trajectories and selection of any desired wind 

direction relative to the final approach heading. Computer-graphic displays were generated on the ground and transmitted via video link to the 

aircraft for presentation in the pilots’ instrument panel. The integrated 

flight-test system permitted manual, augmented, or completely automatic 

control for executing flight trajectories that could be optimized from the 

standpoints of fuel, time, airspace utilization, ride qualities, noise abatement, or air traffic control considerations. Many concepts were explored 

in the IFR program, including flight director control/display concepts and 

signal smoothing techniques, which proved valuable in achieving fully 

automatic approach and landing capability.26 Extensive flight demonstrations were conducted at Wallops Flight Center with the VALT CH-47 

aircraft for Government and industry groups to demonstrate the new 

progress achieved in IFR approach and landing technology.

26. J.R. Kelly, F.R. Niessen, J.J. Thibodeaux, K.R. Yenni, and J.F. Garren, Jr., “Flight Investigation of 

Manual and Automatic VTOL Decelerating Instrument Approaches and Landings,” NASA TN-D-7524 

(July 1974).
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In structures technology, one of the important outcomes of the space 

program was the development and implementation of comprehensive 

computational finite element analyses. State-of-the-art finite element 

methodology was collected from among the large aerospace companies and unified into the NASA Structural Analysis (NASTRAN) computer program. The basic development contract was managed by NASA’s 

Goddard Space Flight Center and then by Langley for improvements 

and distribution to approximately 260 installations. During the early 

1980s, Langley played a key role in bringing advanced structural design 

capability into the helicopter industry. The contribution here was the 

onsite assignment of an experienced structural dynamics specialist at 

a prime manufacturer’s facility to guide the integration of the preliminary static structural design methodology with rotor dynamic analysis 

methodology.27 This avoided the tedious process of repeatedly freezing an airframe structural design effort and each time doing a separate 

dynamic analysis to determine if an acceptable dynamic response criterion was achieved.

During this period, the Army added to its already extensive helicopter crash-test activities by joining with NASA to crash-test the Boeing 

Vertol CH-47C helicopter in the Impact Dynamics Research Facility at 

Langley, which accommodated aircraft up to 30,000 pounds.28 The facility had been converted from a Lunar Landing Research Facility to a 

center for the study of crash effects on aircraft. A unique feature of this 

massive gantry structure was the capability to impact full-scale aircraft 

under free-flight conditions with precise control of attitude and velocity.

The ongoing rotary wing research began to expand in scope with 

the establishment of the Army co-located research groups at the three 

NASA Centers. At Ames, full-scale rotor wind tunnel testing continued 

at an increased pace in the 40- by 80-foot tunnel. In the 1970s, the wind 

tunnel tests included the Sikorsky Advancing Blade Concept (ABC) rotor. 

This rotor concept incorporated two counter-rotating coaxial rotors. The 

hingeless blades were very stiff to allow the advancing blades on both 

sides of the rotor disk to balance the opposing rolling moments thereby 

27. R.G. Kvaternik and W.G. Walton, Jr., “A Formulation of Rotor-Airframe Coupling for the Design 

Analysis of Vibrations of Helicopter Airframes,” NASA RP-1089 (June 1982).

28. Karen Jackson, Richard L. Boitnott, Edwin L. Fasanella, Lisa E. Jones, and Karen H. Lyle, “A 

Summary of DOD-Sponsored Research Performed at NASA Langley’s Impact Dynamics Research 

Facility,” Journal of the American Helicopter Society, vol. 51, no. 1 (June 2004).
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The Sikorsky XH-59A Advancing Blade Concept helicopter was a joint test program between 

the Army, Navy, NASA, and Air Force. NASA

maintaining aircraft trim as airspeed is increased. Forward thrust is 

supplied by auxiliary propulsion rather than by forward tilt of the main 

rotor as in conventional helicopter designs. 

NASA also tested a full-scale semispan wing-pylon-rotor of the Bell 

Helicopter tilt rotor design.29 This test was followed by a similar entry 

of a semispan setup of a Boeing Vertol tilt rotor concept. During this 

period, improvements were made in the 40- by 80-foot Full-Scale Tunnel 

to upgrade the research capability. Its online data capability was augmented by introducing a new Dynamic Analysis System for real-time 

analysis of critical test parameters. A new Rotor Test Apparatus (RTA) 

was added to facilitate full-scale rotor testing. With this new equipment in place, a Kaman Controllable Twist Rotor (CTR) was first investigated in 1975.

In the early 1970s, the modest in-house research funding level for 

rotary wing projects led to seeking other sources within the new, more 

elaborate financial system of NASA. It turned out that contracting outof-house research had become a staple of the rapidly growing procure29. H.K. Edenborough, T.M. Gaffey, and J.A. Weiberg, “Analysis and Tests Confirm Design of 

Proprotor Aircraft,” AIAA Paper No. 72-803 (1972).
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ment system.30 This offered the opportunity to begin to solicit, select, 

and fund small supporting research contracts to augment the in-house 

rotary wing work categorized as Research and Technology Base. In the 

Flight Research Branch at Langley between 1969 and 1974, over 77 contractor reports (CR) and related technical papers were published. The 

performing organizations included industry and university research 

departments. The research topics included analytical and experimental 

investigations of rotor-blade aeroelastic stability, blade-tip vortex aerodynamics, rotor-blade structural loads prediction, free-wake geometry 

prediction, nonuniform swash-plate dynamic analysis program, rotorblade dynamic stall, composite blade structures, and variable geometry rotor concepts, In the mid 1970s, this entry into contracted research 

to augment in-house work was further augmented by teaming of NASA 

and Army rotary wing research at the three NASA Centers. Finally, projects between NASA, the Army, and contractors evolved into major joint 

efforts in Systems Technology and Experimental Aircraft during the 

following decade. 

The mid-1970s brought two major rotary wing experimental aircraft 

programs, both jointly funded and managed by NASA and the Army. 

At Langley, the Rotor Systems Research Aircraft (RSRA) program was 

launched. This was a new approach to conducting flight research on 

helicopter rotor systems.31 Two vehicles were designed and fabricated by 

Sikorsky Aircraft. The basic airframe, propulsion, and control systems 

of the two RSRA vehicles were those of the Sikorsky S-61 Sea King helicopter. In addition, the RSRA incorporated a unique rotor force balance 

system and isolation system, a programmable electronic control system, 

a variable incidence wing with a force balance system, drag brakes, and 

two TF34 auxiliary thrust turbofan engines. As a unique safety feature, 

the three-member-crew ejection system incorporated automatic balanced sequencing of explosive separation of the test rotor-blades as the 

first step in permitting the rapid ejection of the pilot, copilot, and test 

engineer. After design and fabrication at Sikorsky, the first of two RSRA 

vehicles made its first flight in 1976. After initial tests of the helicopter 

configuration, flight-testing was continued at the NASA Wallops Flight 

30. James R. Hansen, Spaceflight Revolution, Langley Research Center From Sputnik to Apollo, 

NASA SP-4308 (Washington, DC: NASA 1995), pp. 81–111.

31. A.W. Linden and M.W. Hellyer, “The Rotor Systems Research Aircraft,” AIAA Paper No. 74-

1277 (1974).
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Center with the Langley–Army project team and contractor onsite support. Acceptance testing was completed by the Langley team, which was 

then joined by Ames flight-test representatives in anticipation of pending transfer of the RSRA program to Ames.

At Ames, a NASA–Army program of equal magnitude was launched 

to design and fabricate two XV-15 Tilt Rotor Research Aircraft (TRRA). 

In this case, the program focused on a proof-of-concept flight investigation. This concept, pursued by rotary wing designers since the early 20th 

century, employs a low-disk-loading rotor at each wingtip that can tilt 

its axis from vertical, providing lift, to horizontal, providing propulsive 

thrust in wing-borne forward flight. The TRRA contract was awarded 

to Bell Helicopter Textron. Late in the program, as the XV-15 reached 

flight status, the United States Navy added funding for special missionsuitability testing. Eventually, XV-15 testing gave confidence to tilt rotor 

advocates who successfully pushed for development of an operational 

system, which emerged as the V-22 Osprey.

The RSRA and TRRA experimental aircraft programs together represented a total initial investment of approximately $90 million, ($337 

million in 2009 dollars), shared equally by NASA and the Army. The size 

and scope of these programs were orders of magnitude beyond previous NACA–NASA rotary wing projects. This represented a new level of 

The NASA–Army Sikorsky S-72 Rotor Systems Research Aircraft in flight at NASA’s Ames 

Research Center. NASA.
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resources in rotary wing research for NASA and with it came considerably more day-to-day visibility within the NASA aeronautics program.

The bicentennial year of 1976 also marked a year of major organizational change in NASA rotary wing research. As part of an overall Agency reassessment of the roles and missions of each Center, the 

Ames Research Center was assigned the lead Center responsibility for 

helicopter research. An objective of the lead Center concept was to consolidate program lead in one Center and, wherever possible, combine 

research efforts of similar nature. As a result, all rotary wing flight test, 

guidance, navigation, and terminal area research were consolidated at 

Ames, which brought these research activities together with the extensive simulation and related flight research facilities. Langley retained 

supporting research roles in structures, noise, dynamics, and aeroelasticity. The realignment of responsibilities and transfer of flight 

research aircraft caused unavoidable turbulence in the day-to-day 

conduct of the rotary wing program from 1976 to 1978. However, the 

momentum of the program gradually returned, and the program grew 

to new levels with NASA and Army research teams at Ames, Langley, and 

Glenn working to carry out their responsibilities in rotary wing research.

At Ames, the testing of full-scale rotor systems continued at 

an increasing pace in the 40 by 80 Full-Scale Tunnel. In 1976, the 

Controllable Twist Rotor concept was tested again, this time with multicyclic control. “Two-per-rev” (two control cycles per one rotor revolution), “three-per-rev,” and “four-per-rev” cyclic control was added to the 

CTR’s servo flap system to evaluate the effectiveness in reducing blade 

stresses and vibration of the fuselage module. Both favorable effects 

were achieved with only minor effect on the rotor power requirements. 

The Sikorsky S-76 rotor system was tested in 1977 in a joint NASA–

Sikorsky investigation of tip shapes. This was followed by a joint NASA–

Bell investigation of the Bell Model 222 fuselage drag characteristics. In 

1978, the NASA–Army XV-15 Tilt Rotor Research Aircraft arrived from 

Bell Helicopter for full-scale wind tunnel tests prior to initiation of its 

own flight tests. The wind tunnel tests revealed a potential tail structural vibration problem that would be further explored in flight following the strengthening of the empennage attachment structure. The next 

rotor test was the Kaman Circulation Control Rotor (CCR) in 1978.32

32. Jack N. Nielsen and James C. Biggers, “Recent Progress in Circulation Control Aerodynamics,” 

AIAA Paper No. 87-0001 (1987).
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A new concept was introduced based on technology developed at the 

David Taylor Ship Research and Development Center (since 1992 the 

Carderock Division of the Naval Surface Weapons Center). The Kaman 

rotor utilized elliptical-shaped airfoils with trailing edge slots. Lift was 

augmented by blowing compressed air from these slots. The need for 

mechanical cyclic blade feathering to provide rotor control was eliminated replaced by cyclic blowing. The wind tunnel testing investigated 

the amount of blowing control necessary to maintain forward flight. 

In 1979, the Lockheed X-Wing Stoppable Rotor was tested in the 40 by 

80 Full-Scale Tunnel. This concept, funded by the Defense Advanced 

Research Projects Agency, also incorporated a circulation control concept. The X-Wing rotor was designed to be stoppable (and startable) at 

high forward flight speed while still carrying lift. Since two of the four 

blade trailing edges become leading edges when stopped, provisions were 

made to provide for separate blowing systems for the leading and trailing edges of the blades. When operating as a fixed X-Wing aircraft, aircraft roll and pitch control were provided by differential blowing from 

the aft edges of opposing, nonrotating blades serving as swept forward 

and aft wings. The wind tunnel tests of the 25-foot-diameter rotor successfully demonstrated the ability to start and stop the rotor at speeds 

of approximately 180 knots (maximum tunnel speed).

The Boeing Vertol Bearingless Main Rotor (BMR) was tested in 

1980.33 The BMR used elastic materials in the construction of the rotor 

hub rather than mechanical bearings for articulation. Such designs have 

aeroelastic stability characteristics different from conventional mechanical systems. Therefore, the wind tunnel tests investigated the degree of 

stability present and established appropriate boundaries for safe flight. 

In addition, in 1980, the Sikorsky Advancing Blade Concept (ABC) coaxial rotor was again tested in the 40 by 80 Full-Scale Tunnel.34 In this 

entry, the full-scale rotor was tested atop a configuration replica of the 

actual XH-59A flight-test aircraft. This testing focused on an investigation of the drag characteristics of the rotor shaft and hubs of the coaxial 

rotors. In an effort to reduce the drag, tests were made with the actual 

fuselage modeled and the actual flight demonstrator hardware compo33. W. Warmbrodt and J.L. McCloud, II, “A Full-Scale Wind Tunnel Investigation of a Helicopter 

Bearingless Main Rotor,” NASA TM-81321 (1981).

34. M. Mosher and R.L. Peterson, “Acoustic Measurements of a Full-Scale Coaxial Helicopter,” 

AIAA Paper No. 83-0722 (1983). 
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nents utilized to explore several inter-rotor fairing configurations. (In 

2008, Sikorsky Aircraft unveiled a new technology demonstrator aircraft 

incorporating the advancing blade concept identified as the X2. In this 

design forward thrust is provided by a pusher propeller installation.) 

In 1984, Ames shut down the 40- by 80-foot facility for tunnel modification to upgrade the 40- by 80-foot section to a speed capability of 

250 knots and add a new 80 by 120 leg to the tunnel facility capable of 

speeds to 80 knots. The upgraded facility, known as the National FullScale Aerodynamics Complex (NFAC), reopened in 1987 and would 

have been operated by NASA until 2010. However, budgetary pressures 

forced its closure in 2003. Four years later, in 2007, the United States 

Air Force’s Arnold Engineering Development Center (AEDC) upgraded 

key operating systems and reopened the facility under a 25-year lease 

with NASA. The anticipated majority customer for this national asset 

was seen to be the United States Army, in collaboration with NASA, in 

support of rotary wing research.

A Helicopter Transmission Technology program was initiated at the 

Glenn Research Center to foster the application of an extensive technology base in bearings, seals, gears, and new concepts specifically to 

helicopter propulsion systems.35 Research continued at a growing pace. 

In order to upgrade the analytical methods for large spiral bevel gears, 

NASA supported the development and validation testing of finite element method computer programs by Boeing Vertol. The opportunity was 

taken to utilize the available aft transmission hardware assets, available 

from the canceled XCH-62 Heavy Lift Helicopter Program, for analytical methods validation data. Another program at Glenn was the joint 

NASA–DARPA Convertible Engine Systems Technology (CEST) program. This program involved the modification of a TF34 turbofan engine 

to a fan/shaft engine configuration for use as a research test engine to 

investigate the performance, control, noise, and transient characteristics. The potential application of CEST was to the X-Wing vehicle concept by using a single-core engine to provide shaft power to a rotor in 

hover and low speed, and conversion capability to provide fan thrust 

for high speed, stopped rotor mode, and flight propulsion. 

Ongoing research in helicopter handling qualities continued and 

expanded at the Ames Research Center. In 1978, one of these programs 

35. Robert C. Ball, “Summary Highlights of the Advanced Rotor Transmission (ART) Program,” AIAA 

Paper No. 92-3362 (1992). 
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provided essential simulation data on the effects of large variations in rotor 

design parameters on handling qualities and agility in helicopter nap-ofthe-Earth (NOE) flight. The parameters investigated including flapping 

hinge offset, flapping hinge restraint, rotor blade inertia, and blade pitchflap coupling. Experiments were carried out on the Ames piloted simulators to systematically study stability and control augmentation systems 

designed to improve NOE flying and handling qualities characteristics. 

New efforts in computational analysis to increase rotor efficiency 

began at Ames. An analytical procedure was developed to predict rotor 

performance trends in relation to changes in the shape of the blade tips. 

The analytical procedure utilized two full potential flow-field computer 

programs developed for computation of the transonic flow field about 

fixed wings and airfoils. The analytical procedure rapidly became a useful tool for predicting aerodynamic performance improvements that may 

be achieved by modifying blade geometry. The procedure was guided by 

design studies and reduced the experimental testing required to select 

blade configurations. NASA continued the long-established tradition of furnishing excellent references for technical practice when, in 1980, research 

scientist Wayne Johnson, a member of the Army–NASA research team 

at Ames, published his book Helicopter Theory, a comprehensive state-ofthe-art coverage of the fundamentals of helicopter theory and engineering 

analysis. The extensive bibliography of cited literature included an extensive listing of rotary wing technical publications authored by researchers 

from the NACA, NASA, the Army, industry, and academia.36

Research accelerated on advancing the ability of a helicopter to 

execute a radar approach. Civil weather/mapping radar could be used 

to provide approach guidance under instrument meteorological conditions (IMC) to select safe landing environments. Onboard radar systems were widely used by helicopter operators to provide approach 

guidance to offshore oil rigs without the need for electronic navigation aids at the landing site. For use over the water, the radar provided 

guidance and ensures obstacle awareness and avoidance, but involved 

very high pilot workload and limited guidance accuracy. For use over 

land, the ground clutter return made these approaches infeasible without more advanced radar systems. Two programs at Ames resulted 

from major advances in radar approaches. One program involved the 

36. Wayne Johnson, Helicopter Theory (Princeton: Princeton University Press, 1980).
NASA’s Contributions to Aeronautics

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The NASA/Army/Bell XV-15 Tilt Rotor Research Aircraft in flight. NASA.

use of a video data processor in conjunction with the weather radar 

for overwater approaches. This system automatically tracked a designated radar target and displayed a pilot-selected approach course. 

The second radar program involved the development of an innovative 

technique to suppress ground clutter radar returns in order to locate 

simple, low-cost radar reflectors near the landing site. This program 

was extended to provide the pilot with precision localizer and glideslope information using airborne weather radar and a ground-based 

beacon or reflector array.

The 1980s brought several major accomplishments in the tilt rotor 

program.37 The second XV-15 aircraft was brought to flight status and 

accepted by the Government after check flights and acceptance ceremonies at NASA’s Dryden Flight Research Center on October 28, 1980. 

It was then used for flight tests aimed at verifying aeroelastic stability, evaluating fatigue load reduction modifications, and expanding the 

maneuver envelope. Subsequently, this aircraft was ferried to Ames, 

where tests continued in the areas of handling qualities, flight control, and expansion of the landing approach envelope. The first XV-15 

aircraft was brought to flight status in late 1980, and initial work was 

37. D.C. Dugan, R.G. Erhart, and L.G. Schroers, “The XV-15 Tilt Rotor Research Aircraft,” NASA 

TM-81244 / AVRADCOM Technical Report 80-A-15 (1980).
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done on a ground tiedown rig to measure the downwash field and noise 

environment. Meanwhile, the second XV-15 participated in the Paris 

Air Show. The aircraft performed daily, on schedule, and received wide 

acclaim as a demonstration of new aeronautical technological achievement. The XV-15 crew concluded each daily performance with a courteous “bow,” the hovering tilt rotor ceremoniously dipping its nose to 

the audience. After the flight demonstration in France and subsequent 

flights in Farnborough, England, the aircraft was returned to Ames for 

continued flight demonstration and proof-of-concept testing. The two 

vehicles achieved a high level of operational reliability, not the usual 

attribute of highly specialized research aircraft. One of the vehicles 

was returned to Bell Helicopter under a cooperative arrangement that 

made the aircraft available to the contractor at no cost in exchange for 

doing a number of program flight-test tasks, particularly in the mission suitability category. The overall success of the NASA–Army XV-15 

(with a rotor diameter of 25 feet and a gross weight of 13,428 pounds) 

proof-of-concept program contribution is reflected in the application of the proven technology to the design and production of the new 

joint-service V-22 Osprey, (rotor diameter: 38 feet; gross weight: 52,000 

pounds). The classic claim of research results having to endure a 20-year 

shelf life before actual engineering design application begins did not 

apply. It took only 5 years to move from achieving proof-of-concept 

with the XV-15 research aircraft to initiation of preliminary design of 

the operational V-22 Osprey.

There has been over a half century of an unbroken series of NACA–

NASA research contributions to tilt rotors since early XV-3 flight assessments and wind tunnel testing in the mid-1950s.38 Since that beginning, 

NACA–NASA researchers have pursued many subject areas, including tilt rotor analytical investigations to solve a rotor/pylon aeroelastic stability problem, dynamic model aeroelastic testing in the Langley 

Transonic Dynamics Tunnel, analytical method development and verification, wind tunnel tests of full-scale rotor/wing/pylon assembles, XV-15 

vehicle wind tunnel tests and flight tests, and detailed investigation of 

many other potential problem areas. This sustained effort and the robust 

demonstration and advocacy of the technology’s potential resulted in 

the XV-15 program being cited in 1993 as “the program that wouldn’t 

38. Martin D. Maisel, Demo J. Giulianetti, and Daniel C. Dugan, The History of The XV-15 Tilt Rotor 

Research Aircraft From Concept to Flight, NASA SP-2000-4517 (Washington, DC: NASA, 2000).
NASA’s Contributions to Aeronautics

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die” in a University of California at Berkeley School of Engineering case 

study in a course on “The Political Process in Systems Architecture.”39

During the early 1980s, the rotary wing activity at Glenn Research 

Center increased with the addition of new transmission test facilities 

rated at 500 and 3,000 horsepower. Research progressed on traction 

drive, hybrid drive, and other advanced technology concepts. The problem of efficient engine operation at partial power settings was addressed 

with initial studies indicating turbine bypass engine concepts offered 

potential solutions. Similar studies on contingency power for oneengine-inoperative (OEI) emergency operation focused on water injection 

and cooling flow modulation. Renewed efforts in aircraft icing included 

rotary wing icing research. A broad scope program was launched to 

study the icing environment, develop basic ice accretion prediction 

methods, acquiring in-flight icing data for comparison with wind tunnel data from airfoil icing tests to verify rotor performance prediction 

methods. In addition, flight tests of a pneumatic deicing boot system 

were conducted using the Ottawa spray rig and the United States Army 

CH-47 in-flight icing spray system. In 1983, research testing began on 

the NASA–DARPA Convertible Engine System Technology program.40

TF34 fan/shaft engine hardware with variable fan inlet guide vanes for 

thrust modulation was used to evaluate fan hub design and map the 

steady-state and transient performance and stability of the concept. New 

rotary wing efforts were also started in the areas of transmission noise, 

and flight/propulsion control integration technology. 

Langley Research Center activity in rotary wing research increased 

substantially within the Structures Directorate, with focused programs 

in acoustics, dynamics, structural materials, and crashworthiness. This 

research was carried out in close association with the Army Structures 

Laboratory, now known as the Vehicle Technology Directorate (VTD). 

NASA and Army joint use of the Langley 4- by 7-meter tunnel for aerodynamic and acoustic model testing became an important feature of 

the rotary wing program. Confirmed progress was achieved in airframe 

dynamic analysis methodology addressing the engineering management and execution of the efficient use of finite element methods for 

39. Brenda Forman, “The V-22 Tiltrotor ‘Osprey:’ The Program That Wouldn’t Die,” Vertiflite, vol. 

39, no. 6, (Nov./Dec. 1993), pp. 20–23.

40. Jack G. McArdle, “Outdoor Test Stand Performance of a Convertible Engine with Variable Inlet 

Guide Vanes for Advanced Rotorcraft Propulsion,” NASA TM-88939 (1986).
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simultaneous tasks of static and dynamic airframe preliminary design.41

These techniques were demonstrated, publicly documented, and verified 

by comparison with shake test data for the CH-47 helicopter airframe. 

Other research related to helicopter dynamics included participation with 

the Army in a program to demonstrate the use of closed-loop multicyclic 

control of rotor-blade pitch motion for vibration reduction. The program 

involved flight-testing of an Army OH-6 helicopter by Hughes Helicopters.42

One of the more innovative approaches to research teaming was 

developed in the area of rotary wing noise. In 1982, discussions between 

the American Helicopter Society and NASA addressed the industry concern that the proposed rulemaking by Federal Aviation Administration 

would place the helicopter industry at a considerable disadvantage. The 

issue was based on the point that the state-of-the-art noise prediction 

did not allow the prediction of noise for new designs with acceptable 

confidence levels. As a result, NASA and the Society, joined by the FAA 

and the Helicopter Association International (HAI)—an organization 

of helicopter commercial operators—embarked on a joint program in 

noise research. Through the AHS, American helicopter manufacturers 

pooled their research with that of NASA under a 5-year plan leading to 

improved noise prediction capability. All research results were shared 

among the Government and industry participants in periodic technical exchanges. Langley managed the program with full participation by 

Ames and Glenn Research Centers in their areas of expertise.

After delivery of the two RSRA vehicles to the Ames Research Center 

in the late 1970s, the helicopter and compound (with wing and TF34 

turbofan engines installed) configurations were involved in an extended 

period of ground- and flight-testing to document the characteristics of 

the basic vehicles. This included extensive calibrations of the onboard 

load measurement systems for the rotor forces and moments; wing lift, 

drag, and pitching moment; and TF34 engine thrust. This work was followed by the initiation the research flight program utilizing the delivered S-61 rotor system. In 1983, NASA and DARPA launched a major 

research program to design, fabricate and flight-test an X-Wing rotor on 

the new RSRA. The RSRA was ideally suited to the testing of new rotor 

41. Raymond G. Kvaternik, “The NASA/Industry Design Analysis Methods for Vibration (DAMVIBS) 

Program—A Government Overview,” AIAA Paper No. 92-2200 (1992).

42. B.P. Gupta, A.H. Logan, and E.R. Wood, “Higher Harmonic Control for Rotary Wing Aircraft,” 

AIAA Paper No. 84-2484 (1984).
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concepts, being specifically design for the purpose. One RSRA vehicle 

was returned to Sikorsky Aircraft for installation of an X-Wing rotor. 

This aircraft was eventually moved to NASA Dryden Flight Research 

Center at Edwards Air Force Base, CA, where final preparations were 

made for flight-testing. The second vehicle embarked on fixed-wing flight 

testing at the Dryden Center to expand and document the flight envelope of the RSRA beyond 200 knots, the speed range of interest in the 

start-stop conversion testing for the X-Wing rotor.

Contributions were beginning to emerge from the NASA–American 

Helicopter Society Rotorcraft Noise Prediction Program, the joint 

Government-industry effort initiated in 1983.43 The four major thrusts 

were: noise prediction, database development, noise reduction, and criteria development. Fundamental experimental and analytical studies were 

started in-house and under grants to universities. In order to obtain highquality noise data for comparison with evolving prediction capability, a 

wind tunnel testing program was initiated. This NASA-sponsored program was performed in 1986 in the Dutch-German wind tunnel (DuitsNederlandse wind tunnel, DNW) using a model-scale Bo 105 main rotor. 

This program was performed with the support of the Federal Aviation 

Administration and the collaboration of the German aerospace research 

establishment. In these tests and in subsequent tests of the model in the 

DNW tunnel in 1988, researchers gained valuable insight into the aeroacoustic mechanism of blade vortex interaction (BVI) noise.

In regard to rotor external noise reduction, Langley researchers 

investigated the possibility of BVI noise reduction using active control 

of blade pitch. A model-scale wind tunnel test was conducted in the 

Langley Transonic Dynamics Tunnel (TDT) using the Aeroelastic Rotor 

Experimental System (ARES).44 Results were encouraging and demonstrated noise level reductions up to 5 decibels (dB) for low and moderate 

forward speeds. A major contribution of the NASA–AHS program was 

the development of a comprehensive rotorcraft system noise prediction 

capability. The primary objective of this capability, the computer code 

named ROTONET, was to provide industry with a reliable predictor for 

43. Ruth M. Martin, “NASA/AHS Rotorcraft Noise Reduction Program: NASA Langley Acoustics 

Division Contributions,” Vertiflite, vol. 35, no. 4, (May/June 1989), pp. 48–52.

44. W.R. Mantay, W.T. Yeager, Jr., M.N. Hamouda, R.G. Cramer, Jr., and C.W. Langston, 

“Aeroelastic model Helicopter Testing in the Langley TDT,” NASA TM-86440 / USAAVSCOM 

TM-85-8-5 (1985). 
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use in design evaluation and noise certification efforts. ROTONET was 

developed in several phases, with each phase released to Noise Reduction 

Program participants for testing and evaluation. Validation data from 

flight test of production and experimental rotorcraft constituted a vital 

element of the program. The first was of the McDonnell-Douglas 500E 

helicopter, tested at NASA’s Wallops Flight Facility. The second flighttest effort at Wallops, a joint NASA–Army program, was performed in 

1987 using an Aerospatiale Dauphine helicopter, which had a relatively 

advanced blade design and a Fenestron-type (ducted) tail rotor. The year 

1988 saw a joint NASA–Bell Helicopter effort in flight investigation of 

the noise characteristics the NASA–Army XV-15 Tilt Rotor Research 

Aircraft. The results indicated that while the aircraft seemed very quiet 

in the airplane mode, significant blade-vortex interaction noise was evident in the helicopter mode of flight. NASA benefited from the interaction with and participation in the variety of industry noise programs, 

which helped set the groundwork for subsequent joint participation in 

rotary wing research.45

NASA 1990–2007: Coping with Institutional and Resource Challenges

Over the next decade and a half, the NASA rotary wing program’s available organizational and financial resources were significantly impacted 

by NASA and supporting Agency organizational, mission, and budget 

management decisions. These decisions were driven by changes in program priorities in the face of severe budget pressures and reorganization 

mandates seeking to improve operational efficiency. NASA leaders were 

being tasked with more ambitious space missions and with recovering 

from two Shuttle losses. In the face of these challenges, the rotary wing 

program, among others, was adjusted in the effort to continue to make 

notable research contributions. Examples of the array of real impacts 

on the rotary wing program over this period were: (1) termination of the 

NASA–DARPA RSRA–X-Wing program; (2) stopping the NASA–Army 

flight operations of the only XV-15 TRRA aircraft and the two RSRA vehicles; (3) transfer of all active NASA research aircraft to Dryden Flight 

Research Center, which essentially closed NASA rotary wing flight operations; (4) elimination of vehicle program offices at NASA Headquarters; 

(5) closing the National Full-Scale Aerodynamic Complex wind tunnel at 

45. Robert J. Huston, Robert A. Golub, and James C. Yu, “Noise Considerations for Tilt Rotor,” 

AIAA Paper 89-2359 (1989).
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Ames in 2003 (reopened under a lease to the United States Air Force in 

2007); (6) converting to full-cost accounting, which represented a new 

burden on vehicle research funding allocations; and (7) the imposition 

of a steady and severe decline in aeronautics budget requests, staring 

in the late 1990s. Overshadowing this retrenching activity in the 1990s 

was the total reorientation, and hence complete transformation, of the 

Ames Research Center from an Aeronautics Research Mission Center 

to a Science Mission Center with the new lead in information technology (IT).46 Responsibility for Ames’s aerodynamics and wind tunnel 

management was assigned to Langley Research Center. The persistent 

turbulence in the NASA rotary wing research community presented a 

growing challenge to the ability to generate research contributions. Here 

is where the established partnership with the United States Army and 

co-located laboratories at Ames, Langley, and Glenn Research Centers 

made it possible to maximize effectiveness by strengthening the combined efforts. In the case of Ames, this was done by creating a new combined Army–NASA Rotorcraft Division. The center of gravity of NASA 

rotary wing research thus gradually shifted to the Army. 

The decision to ground and place in storage the only remaining 

XV-15 TRRA in 1994 was fortunately turned from a real setback to an 

unplanned contribution. Bell Helicopter, having lost the other XV-15, 

N702NA, in an accident in 1992, requested bailment of the Ames aircraft, N703NA, in 1994 to continue its own tilt rotor research, demonstrations, and applications evaluations in support of the ongoing (and 

troubled) V-22 Osprey program. The NASA and Army management 

agreed. As part of the extended use, on April 21, 1995, the XV-15 became 

the first tilt rotor to land at the world’s first operational civil vertiport 

at the Dallas Convention Center Heliport/Vertiport. After its long and 

successful operation and its retirement in 2003, this aircraft is on permanent display at the Smithsonian Institution’s Udvar-Hazy Center at 

Washington Dulles International Airport, Chantilly, VA.

With the military application of proven tilt rotor technology well 

underway with the procurement of the V-22 Osprey by the Marine Corps 

and Air Force, the potential for parallel application of tilt rotor technology to civil transportation was also addressed by NASA. Early studies, 

funded by the FAA and NASA, indicated that the concept had potential 

46. Glenn E. Bugos, Atmosphere of Freedom, Sixty Years at the Ames Research Center, NASA SP4314 (Washington, DC: NASA 2000), pp. 211–246.
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for worldwide application and could be economically viable.47 In late 

1992, Congress directed the Secretary of Transportation to establish a 

Civil Tilt Rotor Development Advisory Committee (CTRDAC) to examine the technical, operational, and economic issues associated with integrating the civil tilt rotor (CTR) into the Nation’s transportation system. 

The Committee was also charged with determining the required additional research and development, the regulatory changes required, and 

the estimated cost of the aircraft and related infrastructure development. In 1995, the Committee issued the findings. The CTR was determined to be technically feasible and could be developed by the United 

States’ industry. It appeared that the CTR could be economically viable 

in heavily traveled corridors. Additional research and development and 

infrastructure planning were needed before industry could make a production decision. In response to this finding, elements of work suggested 

by the CTRDAC were included in the NASA rotorcraft program plans. 

Significant advances in several technological areas would be required 

to enable the tilt rotor concept to be introduced into the transportation 

system. In 1994, researchers at Ames, Langley, and Glenn Research 

Centers launched the Advanced Tiltrotor Transport Technology (ATTT) 

program to develop the new technologies. Because of existing funding limitations, initial research activity was focused on the primary 

concerns of noise and safety. The noise research activity included the 

development of refined acoustic analyses, the acquisition of wind tunnel prop-rotor noise data to validate the analytical method, and flight 

tests to determine the effect of different landing approach profiles on 

terminal area and community noise. The safety effort was related to 

the need to execute approaches and departures at confined urban vertiports. For these situations the capability to operate safely with oneengine-inoperative in adverse weather conditions was required. This area 

was addressed by conducting engine design studies to enable generating high levels of emergency power in OEI situations without adversely 

impacting weight, reliability, maintenance, or normal fuel economy. 

Additional operational safety investigations were carried out on the 

Ames Vertical Motion Simulator to assess crew station issues, control 

law variations, and assign advanced configurations such as the variable diameter tilt rotor. The principal American rotary wing airframe 

47. Maisel, et al., History of the XV-15 Tilt-Rotor, pp. 110–114.
NASA’s Contributions to Aeronautics

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and engine manufacturers participated in the noise and safety investigations, which assured that proper attention was given to the practical 

application of the new technology.48 An initial step in civil tilt rotor aircraft development was taken by Bell Helicopter in September 1998, by 

teaming with Agusta Helicopter Company of Italy, to design, manufacture, and certify a commercial version of the XV-15 aircraft design designated the BA 609. 

Despite the institutional and resource turbulence overshadowing 

rotary wing activity, the NASA and Army researchers persisted in conducting base research. They continued to make contributions to advance 

the state of rotary wing technology applicable to civil and military needs, 

a typical example being the analysis of the influence of the vortex ring 

state (VRS) flight in rapid, steep descents, brought to the forefront by 

initial operating problems experienced by the V-22 Osprey.49 The current NASA Technical Report Server (NTRS) Web site has posted over 

2,200 NASA rotary wing technical reports. Of these, approximately 800 

entries have been posted since 1991—the peak year, with 143 entries. 

These postings facilitate public access to the formal documentation of 

NASA contributions to rotary wing technology. The annual postings gradually declined after 1991. In what may be a mirror image of the state of 

NASA’s realigned rotary wing program, since 2001 the annual totals of 

posted rotary wing reports are in the 20–40 range, with an increasing 

percentage reflecting contributions by Army coauthors.

As the Army and NASA rotary wing research was increasingly linked 

in mutually supporting roles at the co-located centers, outsourcing, 

cooperation, and partnerships with industry and academia also grew. 

In 1995, the Army and NASA agreed to form the National Rotorcraft 

Technology Center (NRTC) occupying a dedicated facility at Ames 

Research Center. This jointly funded and managed organization was 

created to provide central coordination of rotary wing research activities 

of the Government, academia, and industry. Government participation 

included Army, NASA, Navy, and the FAA. The academic laboratories’ 

participation was accomplished by NRTC having acquired the responsibility to manage the Rotorcraft Centers of Excellence (RCOE) program 

48. William J. Snyder, John Zuk, and Hans Mark, “Tilt Rotor Technology Takes Off,” AIAA Paper 

89-2359 (1989). 

49. Wayne Johnson, “Model for Vortex Ring State Influence on Rotorcraft Flight Dynamics,” NASA 

TP-2005-213477 (2005).
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that had been in existence since 1982 under the Army Research Office. In 

1996, the periodic national competition resulted in establishing Georgia 

Institute of Technology, the University of Maryland at College Park, and 

Pennsylvania State University as the three RCOE sites.

The Rotorcraft Industry Technology Association (RITA), Inc., was 

also established in 1996. Principal members of RITA included the 

United States helicopter manufacturers Bell Helicopter Textron, the 

Boeing Company, Sikorsky Aircraft Corporation, and Kaman Aerospace 

Corporation. Supporting members included rotorcraft subsystem manufacturers and other industry entities. Associate Members included a 

growing number of American universities and nonprofit organizations. 

RITA was governed by a Board of Directors supported by a Technical 

Advisory Committee that guided and coordinated the performance 

of the research projects. This industry-led organization and NRTC 

signed a unique agreement to be partners in rotary wing research. The 

Government would share the cost of annual research projects proposed by RITA and approved by NRTC evaluation teams. NASA and the 

Army each contributed funds for 25 percent of the cost of each project—together they matched the industry-member share of 50 percent. 

Over the first 5 years of the Government-industry agreement, the total 

annual investment averaged $20 million. The RITA projects favored 

mid- and near-term research efforts that complemented mid- and longterm research missions of the Army and NASA. Originally, there was 

concern that the research staff of industry competitors would be reluctant to share project proposal information and pool results under the 

RITA banner. This concern quickly turned out to be unfounded as the 

research teams embarked on work addressing common technical problems faced by all participants. 

NRTC was not immune to the challenges posed by limited NASA 

budgets, which eventually caused some cutbacks in NRTC support of 

RITA and the RCOE program. In 2005, the name of the RITA enterprise was changed to the Center for Rotorcraft Innovation (CRI), and 

the principal office was relocated from Connecticut to the Philadelphia 

area.50 Accomplishments posted by RITA–CRI include cost-effective 

integrated helicopter design tools and improved design and manufacturing practices for increased damage tolerance. The area of rotorcraft 

50. The Center for Rotorcraft Innovation (CRI) Web site is: http://www.irotor.org.
NASA’s Contributions to Aeronautics

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operations accomplishments included incorporating developments in 

synthetic vision and cognitive decision-making systems to enhance the 

routine performance of critical piloting tasks and enabling changes 

in the air traffic management system that will help rotorcraft become 

a more-significant participant in the civil transportation system. The 

American Helicopter Society International recognized RITA for one of 

its principal areas of research effort by awarding the Health and Usage 

Monitoring Project Team the AHS 1998 Grover E. Bell Award for “fostering and encouraging research and experimentation in the important 

field of helicopters.” 

As previously noted, in the mid-1990s, NASA Ames’s entire aircraft 

fleet was transferred some 300 miles south to Dryden Flight Research 

Center at Edwards Air Force Base, CA. This inventory included a number of NASA rotary wing research aircraft that had been actively engaged 

since the 1970s.51 However, the U.S. Army Aeroflightdynamics Directorate, 

co-located at Ames since 1970, chose to retain their research aircraft. In 

1997, after several years of negotiation, NASA Headquarters signed a 

directive that Ames would continue to support the Army’s rotorcraft airworthiness research using three military helicopters outfitted for special 

flight research investigations. The AH-1 Cobra had been configured as 

the Flying Laboratory for Integrated Test and Evaluation (FLITE). One 

UH-60 Blackhawk was configured as the Rotorcraft Aircrew Systems 

Concepts Airborne Laboratory (RASCAL) and remained as the focus 

for advanced controls and was utilized by the NASA–Army Rotorcraft 

Division to develop programmable, fly-by-wire controls for nap-of-theEarth maneuvering studies. This aircraft was also used for investigating noise-abatement, segmented approaches using local differential 

Global Positioning System (GPS) guidance. The third aircraft, another 

UH-60 Blackhawk, had been extensively instrumented for the conduct 

of the UH-60 Airloads Program. The principal focus of the program was 

the acquisition of detailed rotor-blade pressure distributions in a wide 

array of flight conditions to improve and validate advanced analytical 

methodology. The last NACA–NASA rotor air-loads flight program of 

this nature had been conducted over three decades earlier, before the 

advent of the modern digital data acquisition and processing revolu51. David D. Few, “A Perspective on 15 Years of Proof-of-Concept Aircraft Development and Flight 

Research at Ames—Moffett by the Rotorcraft and Powered-Lift Flight Projects Division, 1970-1985,” 

NASA RP-1187 (1987).
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tion.52 Again, the persistence of the NASA–Army researchers met the 

institutional and resource challenges and pressed on with fundamental research to advance rotary wing technology.

On December 20, 2006, the White House issued Executive Order 13419 

establishing the first National Aeronautics Research and Development 

Policy. The Executive order was accompanied by the policy statement prepared by the National Science and Technology Council’s Committee on 

Technology. This 13-page document included recommendations to clarify, focus, and coordinate Federal Government aeronautics R&D activities. Of particular note for NASA’s rotary wing community was Section V 

of the policy statement: “Stable and Long-Term Foundational Research 

Guidelines.” The roles and responsibilities of the executive departments 

and agencies were addressed, noting that several executive organizations 

should take responsibility for specific parts of the national foundational 

(i.e., fundamental) aeronautical research program. Specifically, “NASA 

should maintain a broad foundational research effort aimed at preserving the intellectual stewardship and mastery of aeronautics core competencies.” In addition, “NASA should conduct research in key areas related 

to the development of advanced aircraft technologies and systems that 

support DOD, FAA, the Joint Planning and Development Office (JPDO) 

and other executive departments and agencies.53 NASA may also conduct such research to benefit the broad aeronautics community in its 

pursuit of advanced aircraft technologies and systems. . . . ” In supporting research benefiting the broad aeronautics community, care is to be 

taken “to ensure that the government is not stepping beyond its legitimate purpose by competing with or unfairly subsidizing commercial ventures.” There is a strong implication that the new policy may lead NASA’s 

aeronautics role in a return to the more modest, but successful, ways of 

NASA’s predecessor, the National Advisory Committee for Aeronautics, 

with a primary focus on fundamental research, with the participation of 

52. Edwin W. Aiken, Robert A. Jacobson, Michelle M. Eshow, William S. Hindson, and Douglas 

H. Doane, “Preliminary Design Features of the RASCAL—A NASA/Army Rotorcraft In-Flight Simulator,” AIAA Paper 92-4175 (1992); Robert T.N. Chen, William S. Hindson, and Arnold W. Mueller, “Acoustic Flight Tests of Rotorcraft Noise-Abatement Approaches Using Local Differential GPS 

Guidance,” NASA TM-110370 (1995); Robert M. Kufeld and Paul C. Loschke, “UH-60 Airloads 

Program—Status and Plans,” AIAA Paper 91-3142 (1991).

53. In 2003, Congress authorized the Joint Planning and Development Office (JPDO) coordinating 

the activities of multiple Federal agencies in planning Next Generation Air Transportation System to 

implement the transformation of the national airspace system.
NASA’s Contributions to Aeronautics

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academia, and the cooperative research support for systems technology 

and experimental aircraft program investments by the DOD, the FAA, and 

industry. In the case of rotary wing research, since the 1990s, NASA management decisions had moved the residual effort in this direction under 

the pressure of limited resources.

As charged, 1 year after the Executive order and policy statement 

were issued, the National Science and Technology Council issued the 

“National Plan For Aeronautics Research and Development and Related 

Infrastructure.” Rotary wing R&D is specifically identified as being 

among the aviation elements vital to national security and homeland 

defense with a goal of “Developing improved lift, range, and mission 

capability for rotorcraft.” Future NASA rotary wing foundational research 

contributions may also contribute to other goals and objective of the plan. 

For example, under Energy Efficiency and Environment Protection, is 

Goal 2: Advance development of technologies and operations to enable 

significant increases in energy efficiency of the aviation system, and Goal 

3: Advance development of technologies and operational procedures to 

decrease the significant environmental impacts of the aviation system.

Perhaps the most important long-term challenge for the rotary wing 

segment of aviation is the need for focused attention on improved safety. 

In this regard, Goal 2 under the plan section titled “Aviation Safety is 

Paramount” appears to embrace the rotary wing need in calling for developing technologies to reduce accidents and incidents through enhanced 

aerospace vehicle operations on the ground and in the air. The opportunity for making significant contributions in this arena may exist through 

enhanced teaming of NASA and the rotary wing community under the 

International Helicopter Study Team (IHST).54 The goal of the ambitious 

IHST is to work to reduce helicopter accident rates by 80 percent in 

10 years. The participating members of the organization include technical societies, helicopter and engine manufacturers, commercial operator 

and public service organizations, the FAA, and NASA. Past performance 

suggests that the timely application of NASA rotary wing fundamental 

research expertise and unique facilities to this international endeavor 

would spawn significant contributions and accomplishments.

54. Mark Liptak, “International Helicopter Study Team (IHST) Overview Briefing,” presented at 

Helicopter Association International HELI EXPO Meeting, Houston, TX, Feb. 21–23, 2009 (see 

http://www.ihst.org).
Case 3 | NACA–NASA and the Rotary Wing Revolution

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Recommended Additional Readings

Reports, Papers, Articles, and Presentations: 

Edwin W. Aiken, Robert A. Jacobson, Michelle M. Eshow, William S. 

Hindson, and Douglas H. Doane, “Preliminary Design Features of 

the RASCAL—A NASA/Army Rotorcraft In-Flight Simulator,” AIAA 

Paper 92-4175 (1992).

F.J. Bailey, Jr., “A Simplified Theoretical Method of Determining the 

Characteristics of a Lifting Rotor in Forward Flight,” NACA Report 

716 (1941).

Robert C. Ball, “Summary Highlights of the Advanced Rotor Transmission 

(ART) Program,” AIAA Paper 92-3362 (1992).

G.W. Brooks, “The Application of Models to Helicopter Vibration and 

Flutter Research,” Proceedings of the ninth annual forum of the 

American Helicopter Society (May 1953).

Robert T.N. Chen, William S. Hindson, and Arnold W. Mueller, “Acoustic 

Flight Tests of Rotorcraft Noise-Abatement Approaches Using Local 

Differential GPS Guidance,” NASA TM-110370 (1995). 

R.P. Coleman, “Theory of Self-Excited Mechanical Oscillations of 

Hinged Rotor Blades,” NACA WR-L-308 [formerly NACA Advanced 

Restricted Report 3G29] (1943). 

G.E. Cooper, “Understanding and Interpreting Pilot Opinion,” 

Aeronautical Engineering Review, vol. 16, no. 3, (Mar. 1957), p. 47–51.

G.E. Cooper and R.P. Harper, Jr., “The Use of Pilot Rating in the 

Evaluation of Aircraft Handling Qualities,” NASA TN-D-5153 (1969).

I.H. Culver and J.E. Rhodes, “Structural Coupling in the Blades of a 

Rotating Wing Aircraft,” IAS Paper 62-33 (1962). 

Mike Debraggio, “The American Helicopter Society—A Leader for 40 

Years,” Vertiflite, vol. 30, no. 4 (May–June 1984).
NASA’s Contributions to Aeronautics

174

3

D.C. Dugan, R.G. Erhart, and L.G. Schroers, “The XV-15 Tilt Rotor 

Research Aircraft,” NASA TM-81244 / AVRADCOM Technical Report 

80-A-15 (1980).

H.K. Edenborough, T.M. Gaffey, and J.A. Weiberg, “Analysis and Tests 

Confirm Design of Proprotor Aircraft,” AIAA Paper 72-803 (1972).

Don Fertman, “The Helicopter History of Sikorsky Aircraft,” Vertiflite, 

vol. 30, no. 4 (May–June 1984).

David D. Few, “A Perspective on 15 Years of Proof-of-Concept Aircraft 

Development and Flight Research at Ames–Moffett by the Rotorcraft 

and Powered-Lift Flight Projects Division, 1970–1985,” NASA 

RP-1187 (1987).

Brenda Forman, “The V-22 Tiltrotor ‘Osprey:’ The Program That Wouldn’t 

Die,” Vertiflite, vol. 39, no. 6, (Nov.–Dec. 1993), pp. 20–23.

F. Garren, J.R. Kelly, and R.W. Summer, “VTOL Flight Investigation to 

Develop a Decelerating Instrument Approach Capability,” Society 

of Automotive Engineers Paper No. 690693 (1969).

B.P. Gupta, A.H. Logan, and E.R. Wood, “Higher Harmonic Control for 

Rotary Wing Aircraft,” AIAA Paper 84-2484 (1984).

F.B. Gustafson, “Effects on Helicopter Performance of Modifications 

in Profile-Drag Characteristics of Rotor-Blade Airfoil Sections,” 

NACA WR-L-26 [Formerly NACA Advanced Confidential Report 

ACR L4H05] (1944).

F.B. Gustafson, “A History of NACA Research on Rotating-Wing Aircraft,” 

Journal of the American Helicopter Society, vol. 1, no. 1 (Jan. 1956), 

p. 16.

F.B. Gustafson, “History of NACA/NASA Rotating-Wing Aircraft Research, 

1915–1970,” Vertiflite, Reprint VF-70 (Apr. 1971), pp. 1–27.

Harry H. Heyson and S. Katzoff, “Induced Velocities Near a Lifting 

Rotor with Nonuniform Disk Loading,” NACA Report 1319 (1957).
Case 3 | NACA–NASA and the Rotary Wing Revolution

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3

R.J. Huston, “An Exploratory Investigation of Factors Affecting the 

Handling Qualities of a Rudimentary Hingeless Rotor Helicopter,” 

NASA TN-D-3418 (1966).

Robert J. Huston, Robert A. Golub, and James C. Yu, “Noise 

Considerations for Tilt Rotor,” AIAA Paper 89-2359 (1989).

Karen Jackson, Richard L. Boitnott, Edwin L. Fasanella, Lisa E. Jones, 

and Karen H. Lyle, “A Summary of DOD-Sponsored Research 

Performed at NASA Langley’s Impact Dynamics Research Facility,” 

Journal of the American Helicopter Society, vol. 51, no. 1 (June 2004).

Wayne Johnson, “Model for Vortex Ring State Influence on Rotorcraft 

Flight Dynamics,” NASA TP-2005-213477 (2005).

J.R. Kelly, F.R. Niessen, J.J. Thibodeaux, K.R. Yenni, and J.F. Garren, Jr., 

“Flight Investigation of Manual and Automatic VTOL Decelerating 

Instrument Approaches and Landings,” NASA TN-D-7524 (1974).

Robert M. Kufeld and Paul C. Loschke, “UH-60 Airloads Program—

Status and Plans,” AIAA Paper 91-3142 (1991).

R.G. Kvaternik and W.G. Walton, Jr., “A Formulation of Rotor-Airframe 

Coupling for the Design Analysis of Vibrations of Helicopter 

Airframes,” NASA RP-1089 (1982).

R.G. Kvaternik, “The NASA/Industry Design Analysis Methods for 

Vibration (DAMVIBS) Program—A Government Overview,” AIAA 

Paper 92-2200 (1992).

A.W. Linden and M.W. Hellyer, “The Rotor Systems Research Aircraft,” 

AIAA Paper No. 74-1277 (1974).

Mark Liptak, “International Helicopter Study Team (IHST) Overview 

Briefing,” Helicopter Association International HELI EXPO Meeting, 

Houston, TX, Feb. 21–23, 2009.
NASA’s Contributions to Aeronautics

176

3

W.R. Mantay, W.T. Yeager, Jr., M.N. Hamouda, R.G. Cramer, Jr., and C.W. 

Langston, “Aeroelastic model Helicopter Testing in the Langley TDT,” 

NASA TM-86440 / USAAVSCOM TM-85-8-5 (1985).

Ruth M. Martin, “NASA/AHS Rotorcraft Noise Reduction Program: 

NASA Langley Acoustics Division Contributions,” Vertiflite, vol. 35, 

no. 4 (May–June 1989), pp. 48–52.

J.G. McArdle, “Outdoor Test Stand Performance of a Convertible Engine 

with Variable Inlet Guide Vanes for Advanced Rotorcraft Propulsion,” 

NASA TM-88939 (1986).

M. Mosher and R.L. Peterson, “Acoustic Measurements of a Full-Scale 

Coaxial Helicopter,” AIAA Paper 83-0722 (1983). 

Jack N. Nielsen and James C. Biggers, “Recent Progress in Circulation 

Control Aerodynamics,” AIAA Paper 87-0001 (1987).

Wilmer H. Reed, III, “Review of Propeller-Rotor Whirl Flutter,” NASA 

TR-R-264 (1967).

S. Salmirs and R.J. Tapscott, “The Effects of Various Combinations 

of Damping and Control Power on Helicopter Handling Qualities 

During Both Instrument and Visual Flight,” NASA TN-D-58 (1959).

James Sheiman, “A Tabulation of Helicopter Rotor-Blade Differential 

Pressures, Stresses, and Motions As Measured In Flight,” NASA 

TM-X-952 (1964).

J. Sheiman and L.H. Ludi, “Qualitative Evaluation of Effect of Helicopter 

Rotor Blade Tip Vortex on Blade Airloads,” NASA TN-D-1637 (1963).

William J. Snyder, John Zuk, and Hans Mark, “Tilt Rotor Technology 

Takes Off,” AIAA Paper 89-2359 (1989).

John F. Ward, “Helicopter Rotor Periodic Differential Pressures and 

Structural Response Measured in Transient and Steady-State Maneuvers,” 

Journal of the American Helicopter Society, vol. 16, no. 1 (Jan. 1971).
Case 3 | NACA–NASA and the Rotary Wing Revolution

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3

John F. Ward, “An Updated History of NACA/NASA Rotary-Wing Aircraft 

Research 1915-1984,” Vertiflite, vol. 30, no. 4 (May–June 1984), pp. 

108–117. 

W. Warmbrodt and J.L. McCloud, II, “A Full-Scale Wind Tunnel 

Investigation of a Helicopter Bearingless Main Rotor,” NASA 

TM-81321 (1981).

William Warmbrodt, Charles Smith, and Wayne Johnson, “Rotorcraft 

Research Testing in the National Full-Scale Aerodynamics Complex 

at NASA Ames Research Center,” NASA TM-86687 (1985).

J.B. Wheatley, “Lift and Drag Characteristics and Gliding Performance 

of an Autogiro as Determined In Flight,” NACA Report 434 (1932).

J.B. Wheatley, “A Aerodynamic Analysis of the Autogiro Rotor With 

Comparison Between Calculated and Experimental Results,” NACA 

Report 487 (1934).

Edgar C. Wood, “The Army Helicopter, Past, Present and Future,” Journal 

of the American Helicopter Society, vol. 1, no. 1 (Jan. 1956), pp 87–92.

William T. Yeager, Jr., and Raymond G. Kvaternik, “A Historical 

Overview of Aeroelasticity Branch and Transonic Dynamics Tunnel 

Contributions to Rotorcraft Technology and Development,” NASA 

TM-2001-211054, ARL-TR-2564, (2001).

Books and Monographs:

Walter J. Boyne and Donald S. Lopez, eds., Vertical Flight: The Age of 

the Helicopter (Washington: Smithsonian Institution Press, 1984).

Glenn E. Bugos, Atmosphere of Freedom, Sixty Years at the Ames Research 

Center, NASA SP-4314 (Washington, DC: NASA, 2000).

Alfred Gessow and Garry C. Myers, Jr., Aerodynamics of the Helicopter

(New York: The Macmillan Company, 1952; reissued by Frederick 

Ungar Publishing Co., 1967).
NASA’s Contributions to Aeronautics

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3

Michael H. Gorn, Expanding the Envelope: Flight Research at NACA and 

NASA (Lexington: The University Press of Kentucky, 2001).

Richard P. Hallion and Michael H. Gorn, On the Frontier: Experimental 

Flight at NASA Dryden (Washington, DC: Smithsonian Books, 2003).

James R. Hansen, Engineer in Charge: A History of the Langley Aeronautical 

Laboratory, 1917–1958, NASA SP-4305 (Washington, DC: NASA, 

1987).

James R. Hansen, Spaceflight Revolution: Langley Research Center From 

Sputnik to Apollo, NASA SP-4308 (Washington, DC: NASA, 1995).

Edwin P. Hartman, Adventures in Research, A History of Ames Research 

Center, 1940–1965, NASA SP-4302 (Washington, DC: NASA, 1970).

Michael J. Hirschberg, The American Helicopter: An Overview of Helicopter 

Developments in America, 1908–1999 (Arlington, VA: ANSER, 1999).

Wayne Johnson, Helicopter Theory (Princeton: Princeton University 

Press, 1980).

Martin D. Maisel, Demo J. Giulianetti, and Daniel C. Dugan, The History 

of The XV-15 Tilt Rotor Research Aircraft From Concept to Flight, 

NASA SP-2000-4517 (Washington, DC: NASA, 2000).

J.A. Stockfisch, The 1962 Howze Board and Army Combat Developments, 

Monograph Report MR-435-A (Santa Monica: The RAND 

Corporation, 1994).

Lt. Gen. John J. Tolson, Airmobility, 1961–1971, a volume in the U.S. 

Army Vietnam Studies series (Washington, DC: U.S. Army, 1973).
THIS PAGE INTENTIONALLY BLANK
180

Aerodynamic model of NASA’s SCAT-15F supersonic transport design attached for a subsonic 

wind tunnel test in 1969. NASA.
181

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Softening the Sonic Boom: 

50 Years of NASA Research

Lawrence R. Benson

The advent of practical supersonic flight brought with it the shattering shock of the sonic boom. From the onset of the supersonic age 

in 1947, NACA–NASA researchers recognized that the sonic boom 

would work against acceptance of routine overland supersonic aircraft operation. In concert with researchers from other Federal and military organizations, they developed flight-test programs and innovative 

design approaches to reshape aircraft to minimize boom effects while 

retaining desirable high-speed behavior and efficient flight performance.

4

CASE

AFTER ITS FORMATION IN 1958, the National Aeronautics and Space 

Administration (NASA) began devoting most of its resources to the 

Nation’s new civilian space programs. Yet 1958 also marked the 

start of a program in the time-honored aviation mission that the Agency 

inherited from the National Advisory Committee for Aeronautics (NACA). 

This task was to help foster an advanced passenger plane that would fly 

at least twice the speed of sound. 

Because of economic and political factors, developing such an aircraft 

became more than a purely technological challenge. One of the major barriers 

to producing a supersonic transport involved a phenomenon of atmospheric 

physics barely understood in the late 1950s: the shock waves generated by 

supersonic flight. Studying these “sonic booms” and learning how to control them became a specialized and enduring field of NASA research for the 

next five decades. During the first decade of the 21st century, all the study, 

testing, and experimentation of the past finally began to reap tangible 

benefits in the same California airspace where supersonic flight began.1

1. The author is grateful to Karl Bender of NASA’s Dryden Research Library for helping to gather source 

materials. For a concise introduction to sonic boom theory, see Kenneth J. Plotkin and Domenic J. 

Maglieri, “Sonic Boom Research: History and Future,” American Institute of Aeronautics and Astronautics 

(AIAA), Paper 2003-3575, June 23, 2003.
NASA’s Contributions to Aeronautics

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From Curiosity to Controversy

In 1947, Muroc Army Airfield, CA, was a small collection of aircraft hangars and other austere buildings adjoining the vast Rogers Dry Lake in 

the high desert of the Antelope Valley, across the San Gabriel Mountains 

from the Los Angeles basin. Because of the airfield’s remoteness and clear 

skies, a small team of Air Force, the NACA, and contractor personnel 

was using Muroc for a secret project to explore the still unknown territory of supersonic flight. On October 14, more than 40,000 feet over 

the little desert town of Boron, visible only by its contrail, Capt. Chuck 

Yeager’s 31-foot-long rocket-propelled Bell XS-1 successfully “broke” the 

fabled sound barrier.2 The sonic boom from his little experimental airplane—the first to fly supersonic in level flight—probably did not reach 

the ground on that historic day.3 Before long, however, the acoustical 

signature of the shock waves generated by XS-1s and other supersonic 

aircraft became a familiar sound at and around the isolated airbase. 

In the previous century, an Austrian physicist-philosopher, Ernst 

Mach, was the first to explain the phenomenon of supersonic shock 

waves, which he displayed visually in 1887 with a cleverly made photograph showing those formed by a high-velocity projectile, in this case a 

bullet. The speed of sound, he also determined, varied in relation to the 

density of the medium though which it passed, such as air molecules. 

(At sea level, the speed of sound is 760 mph.) In 1929, Jakob Ackeret, 

a Swiss fluid dynamicist, named this variable “Mach number” in his 

honor. This guaranteed that Ernst would be remembered by future generations, especially after it became known that the 700 mph speed of 

Yeager’s XS-1, flying at 43,000 feet, was measured as Mach 1.06.4

Humans have long been familiar with and often frightened by 

natural sonic booms in the form of thunder, i.e., sudden surges of air 

2. For its development and testing, see Richard P. Hallion, Supersonic Flight: Breaking the Sound 

Barrier and Beyond: The Story of the Bell X-1 and Douglas D-558 (New York: Macmillan, 1977). 

3. Some of the personnel stationed at Muroc when Yeager broke the sound barrier later recalled 

hearing a sonic boom, but these may have been memories of subsequent flights at higher speeds. 

One of NASA’s top sonic boom experts has calculated that at Mach 1.06 and 41,000 feet above 

ground level, atmospheric refraction and absorption of the shock waves would almost certainly have 

dissipated the XS-1’s sonic boom before it could reach the surface. E-mail, Edward A. Haering, 

Dryden Flight Research Center, to Lawrence R. Benson, Apr. 8, 2009.

4. “Ernst Mach,” Stanford Encyclopedia of Philosophy, Mar. 21, 2008, http://plato.stanford.

edu/entries/Ernst-mach; Jeff Scott, “Ernst Mach and Mach Number,” Nov. 9, 2003, http://www.

aerospaceweb.org/question/history/q0149.shmtl. 
Case 4 | Softening the Sonic Boom: 50 Years of NASA Research

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Bell XS-1—the first aircraft to exceed Mach 1 in level flight, October 14, 1947. U.S. Air Force.

pressure caused when strokes of lightning instantaneously heat contiguous columns of air molecules. Perhaps the most awesome of sonic 

booms—heard only rarely—have been produced by large meteoroid 

fireballs speeding through the atmosphere. On a much smaller scale, 

the first acoustical shock waves produced by human invention were the 

modest cracking noises from the snapping of a whip. The high-power 

explosives perfected in the latter half of the 19th century were able—as 

Mach explained—to propel projectiles faster than the speed of sound. 

Their acoustical shock waves would be among the cacophony of fearsome sounds heard by millions of soldiers during the two World Wars.5

On a Friday evening, September 8, 1944, an explosion blew out a 

large crater in Stavely Road, west of London. The first German V-2 ballistic missile aimed at England had announced its arrival. “After the 

explosion came a double thunderclap caused by the sonic boom catch5. By the end of World War II, ballistic waves were well understood, e.g., J.W.M. Dumond, et al., 

“A Determination of the Wave Forms and Laws of Propagation and Dissipation of Ballistic Shock 

Waves,” Journal of the Acoustical Society of America (hereinafter cited as JASA), vol. 18, no. 1 

(Jan. 1946), pp. 97–118.
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ing up with the fallen rocket.”6 For the next 7 months, millions of people would hear these sounds, which would become known as “sonic 

bangs” in Britain, from more than 3,000 V-2s launched at England as 

well as liberated portions of France, Belgium, and the Netherlands. Their 

sound waves would always arrive too late to warn any of those unfortunate enough to be near the missiles’ points of impact.7

 After World War 

II, these strange noises faded into memory for several years—until the 

arrival of new jet fighter planes.

In November 1949, the NACA designated its growing detachment 

at Muroc as the High-Speed Flight Research Station (HSFRS), 1 month 

before the Air Force renamed the installation Edwards Air Force Base 

(AFB).8 By the early 1950s, the desert and mountains around Edwards 

reverberated with the occasional sonic booms of experimental and prototype aircraft, as did other flight-test locations in the United States 

and United Kingdom. Scientists and engineers had been familiar with 

the “axisymmetric” ballistic shock waves of projectiles such as artillery shells (referred to scientifically as bodies of revolution).9

 This was 

one reason the fuselage of the XS-1 was shaped like a 50-caliber bullet. But these new acoustic phenomena—many of which featured a 

double-boom sound—hinted that they were more complex. In late 

1952, the editors of the world’s oldest aeronautical weekly stated with 

some hyperbole that “the ‘supersonic bang’ phenomenon, if only by 

reason of its sudden incidence and the enormous public interest it has 

aroused, is probably the most spectacular and puzzling occurrence in the 

history of aerodynamics.”10

6. David Darling: The Complete Book of Spaceflight: From Apollo 1 to Zero Gravity (Hoboken, 

NJ: John Wiley and Sons, 2003), p. 457. See also “Airpower: Missiles and Rockets in Warfare,” 

http://www.centennialofflight.gov/essay/Air_Power/Missiles/AP29.htm; and Bob Ward, Dr. 

Space: The Life of Wernher von Braun (Annapolis: Naval Institute, 2005), p. 43.

7. The definitive biography, Van Braun: Dreamer of Space, Engineer of War, by Michael J. Neufeld 

(New York: Alfred A. Knopf, 2007), pp. 133–136, leaves open the question of whether the 

Germans at Peenemünde heard the first manmade sonic booms in 1942 when their A-4 test rockets 

exceeded Mach 1 about 25 seconds after launch.

8. For the authoritative history of the NACA/NASA mission at Edwards AFB, see Richard P. Hallion 

and Michael H. Gorn, On the Frontier: Experimental Flight at NASA Dryden (Washington, DC: 

Smithsonian, 2003).

9. Plotkin and Maglieri, “Sonic Boom Research,” pp. 1–2. 

10. Introduction to “The Battle of the Bangs,” Flight and Aircraft Engineer, vol. 61, no. 2289 (Dec. 

5, 1952), p. 696, http://www.flightglobal.com/pdfarchive/view/1952/%203457.
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A young British graduate student, Gerald B. Whitham, was the 

first to analyze thoroughly the abrupt rise in air pressure upon arrival 

of a supersonic vehicle’s “bow wave,” followed by a more gradual but deeper fall in pressure for a fraction of a second, and then a 

recompression with the passing of the vehicle’s tail wave. As shown in a 

simplified fashion by Figure 1, this can be illustrated graphically by an 

elongated capital “N” (the solid line) transecting a horizontal axis (the 

dashed line) representing ambient air pressure during a second or less 

of elapsed time. For Americans, the pressure change is usually expressed 

in pounds per square foot (psf—also abbreviated as lb/ft2).

Because a jet fighter (or a V-2 missile) is much longer than an artillery shell is, the human ear could detect a double boom if its tail shock 

wave arrived a tenth of a second or more after its bow shock wave. 

Whitham was first to systematically examine the more complex shock 

waves, which he called the F-function, generated by “nonaxisymmetrical” (i.e., asymmetrical) configurations, such as airplanes.11

The number of these double booms multiplied in the mid-1950s as 

the Air Force Flight Test Center (AFFTC) at Edwards (assisted by the 

HSFRS) began putting a new generation of Air Force jet fighters and 

interceptors, known as the Century Series, through their paces. The 

remarkably rapid advance in aviation technology (and priorities of the 

Cold War “arms race”) is evident in the sequence of their first flights at 

Edwards: YF-100 Super Sabre, May 1953; YF-102 Delta Dagger, October 

1953; XF-104 Starfighter, February 1954; F-101 Voodoo, September 

1954; YF-105 Thunderchief, October 1955; and F-106 Delta Dart, 

December 1956.12

With the sparse population living in California’s Mojave Desert 

region during the 1950s, disturbances caused by the flight tests of new jet 

aircraft were not a serious issue. But even in the early 1950s, the United 

11. G.B. Whitham, “The Flow Pattern of a Supersonic Projectile,” Communications on Pure 

and Applied Mathematics, vol. 5, no. 3 (1952), pp. 301–348 (available at http://www3.

interscience.wiley.com/journal/113395160/issue) and “On the Propagation of Weak Shock 

Waves,” Journal of Fluid Dynamics, vol. 1, No. 3 (Sept. 1956), pp. 290–318 (available at 

http://journals.cambridge.org/action/displayJournal?jid=JFM), and described in Larry J. Runyan, 

et al., Sonic Boom Literature Survey, vol. II, Capsule Summaries, (Seattle: Boeing Commercial 

Airplane Co. for the FAA), Sept. 1973, pp. 6–8, 59–60. Whitham later taught at both the 

Massachusetts and California Institutes of Technology.

12. Air Force Flight Test Center History Office, Ad Inexplorata: The Evolution of Flight Testing at 

Edwards Air Force Base (Edwards AFB: AFFTC, 1996), Appendix B, p. 55.
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Figure 1. Simplified N-shaped sonic boom signature. NASA.

States Air Force (USAF) became concerned about their future impact. 

In November 1954, for example, its Aeronautical Research Laboratory 

at Wright-Patterson AFB, OH, submitted a study to the Air Force Board 

of top generals on early findings regarding the still somewhat mysterious nature of sonic booms. Although concluding that low-flying aircraft flying at supersonic speeds could cause considerable damage, the 

report optimistically predicted the possibility of supersonic flight without booms at altitudes over 35,000 feet.13

As the latest Air Force and Navy fighters went into full production and began flying from bases throughout the Nation, much of the 

American public was exposed to jet noise for the first time. This included 

the thunderclap-like thuds characteristic of sonic booms—often accompanied by rattling windowpanes. Under certain conditions, as the U.S. 

armed services and British Royal Air Force (RAF) had learned, even 

maneuvers below Mach 1 (e.g., accelerations, dives, and turns) could 

generate and focus transonic shock waves in such a manner as to cause 

strong sonic booms.14 Indeed, residents of Southern California began 

hearing such booms in the late 1940s, when North American Aviation 

was flight-testing its new F-86 Sabre. The first civilian claim against the 

13. John G. Norris, “AF Says ‘Sonic Boom’ Can Peril Civilians,” Washington Post and Times Herald (hereinafter cited as Washington Post), Nov. 9, 1954, pp. 1, 12.

14. One of the first studies on focused booms was G.M. Lilley, et al., “Some Aspects of Noise from 

Supersonic Aircraft,” Journal of the Royal Aeronautical Society, vol. 57 (June 1953), pp. 396–414, 

as described in Runyan, Sonic Boom Capsule Summaries, p. 54. AFFTC used F-100s to conduct 

the first in-flight boom measurements: Marshall E. Mullens, “A Flight Test Investigation of the Sonic 

Boom,” AFFTC TN-56-20, May 1956. 
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USAF for sonic boom damage was apparently filed at Eglin AFB, FL, in 

1951, when only subsonic jet fighters were assigned there.15 Additionally, 

as shown in 1958 by Frank Walkden, another English mathematician, 

the lift effect of airplane wings could magnify the strength of sonic booms 

more than previously estimated.16

Sonic boom claims against the Air Force first became statistically 

significant in 1957, reflecting its growing inventory of Century fighters and the type of maneuvers they sometimes performed, which could 

focus acoustical rays into what became called “super booms.” (It was 

found that these powerful but localized booms had a U-shaped signature, with the tail shock wave as well as that from the nose of the airplane being above ambient air pressure.) Most claims involved broken 

windows or cracked plaster, but some were truly bizarre, such as the 

death of pets or the insanity of livestock. In addition to these formal 

claims, Air Force bases, local police switchboards, and other agencies 

received an uncounted number of phone calls about booms, ranging 

from merely inquisitive to seriously irate.17 Complaints from constituents 

also became an issue for the U.S. Congress.18 Between 1956 and 1968, 

some 38,831 claims were submitted to the Air Force, which approved 

14,006 in whole or in part—65 percent for broken glass, 21 percent for 

cracked plaster (usually already weakened), 8 percent for fallen objects, 

and 6 percent for other reasons.19

The military’s problem with sonic boom complaints seems to have 

peaked in the 1960s. One reason was the sheer number of fighter-type 

aircraft stationed around the Nation (over three times as many as 

today). Secondly, many of these aircraft’s missions were air defense. 

This often meant flying at high speed over populated areas for training in 

15. History of the 3201 Air Base Group, Eglin AFB, Jul.–Sept. 1951, Abstract from Information Retrieval and Indexing System (IRIS) No. 438908, Air Force Historical Research Center, Maxwell AFB, AL.

16. F. Walkden, “The Shock Pattern of a Wing-Body Combination Far from the Flight Path,” Aeronautical Quarterly, vol. 9, pt. 2 (May 1958), pp. 164–194; described in Runyan, Sonic Boom 

Capsule Summaries, 8–9. Both Walkden and Whitman did their pioneering studies at the University 

of Manchester.

17. Fred Keefe and Grover Amen, “Boom,” The New Yorker, May 16, 1962, pp. 33–34.

18. Albion B. Hailey, “AF Expert Dodges Efforts to Detail ‘Sonic Boom’ Loss,” Washington Post, 

Aug. 25, 1960, p. A15.

19. J.P. and E.G.R Taylor, “A Brief Legal History of the Sonic Boom in America,” Aircraft Engine Noise 

and Sonic Boom (Neuilly Sur Seine, France: NATO Advisory Group for Aerospace Research and 

Development [AGARD], 1969), Conference Proceedings (CP) No. 42, Paris, May 1969, pp. 2-1–2-11.
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defending cities and other key targets from aerial attack, sometimes in practice against Strategic Air Command (SAC) bombers. The North American 

Air Defense Command (NORAD) conducted two of the largest such 

exercises, Skyshield I and Skyshield II, in 1960 and 1961. The Federal 

Aviation Agency (FAA) shut down all civilian air traffic while NORAD’s 

interceptors and SAC bombers (augmented by some from the RAF) battled 

overhead—accompanied by a sporadic drumbeat of sonic booms 

reaching the surface.20

Although most fighters and interceptors deployed in the 1960s could 

readily fly faster than sound, they could only do so for a short distance 

because of the rapid fuel consumption of jet engine afterburners. Thus, 

their sonic boom “carpets” were relatively short. However, one supersonic American warplane that became operational in 1960 was designed 

to fly faster than Mach 2 for more than 1,000 miles. 

This innovative but troublesome aircraft was the SAC’s new Convairbuilt B-58 Hustler medium bomber. On March 5, 1962, the Air Force showed 

off the long-range speed of the B-58 by flying one from Los Angles to New 

York in just over 2 hours at an average pace of 1,215 mph (despite having 

to slow down for an aerial refueling over Kansas). After another refueling 

over the Atlantic, the same Hustler “outraced the sun” (i.e., flew faster than 

Earth’s rotation) back to Los Angles with one more refueling, completing 

the record-breaking round trip at an average speed of 1,044 mph.21

Capable of sustained Mach 2+ speeds, the four-engine delta-winged 

Hustler (weighing up to 163,000 pounds) helped demonstrate the feasibility of a supersonic transport. But the B-58’s performance revealed at 

least one troubling omen. Almost wherever it flew supersonic over populated areas, the bomber left sonic boom complaints and claims in its 

wake. Indeed, on its record-shattering flight of March 1962, flown mostly 

at an altitude of 50,000 feet (except when coming down to 30,000 feet for 

refueling), “the jet dragged a sonic boom 20 to 40 miles wide back and 

forth across the country—frightening residents, breaking windows, crack20. “Warplanes Fill Skies Over U.S. and Canada,” Los Angeles Times, Sept. 10, 1960, p. 4; 

Albion B. Halley and Warren Kornberg, “U.S. Tests Air Defenses in 3000-Plane ‘Battle,’’’ Washington Post, Oct. 15, 1961, pp. A1, B1; Richard Witkin, “Civilian Planes Halted 12 Hours in Defense 

Test,” New York Times, Oct. 15, 1961, pp. 1, 46.

21. Marcelle S. Knaack, Post-World War II Bombers, 1945–1973 (Washington, DC: Government 

Printing Office (hereinafter cited as GPO) for Office of Air Force History, 1988), pp. 394–395 (vol. 

2, Encyclopedia of U.S. Air Force Aircraft and Missile Systems).
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Convair B-58 Hustler, the first airplane capable of sustained supersonic flight and a major 

contributor to early sonic boom research. USAF.

ing plaster, and setting dogs to barking.”22 As indicated by Figure 2, the B-58 

became a symbol for sonic boom complaints (despite its small numbers).

Most Americans, especially during times of increased Cold War tensions, tolerated occasional disruptions justified by national defense. But 

how would they react to constantly repeated sonic booms generated by 

civilian jet airliners? Could a practical passenger-carrying supersonic airplane be designed to minimize its sonic signature enough to be acceptable to people below? NASA’s attempts to resolve these two questions 

occupy the remainder of this history.

A Painful Lesson: Sonic Booms and the Supersonic Transport

By the late 1950s, the rapid pace of aeronautical progress—with new 

turbojet-powered airliners flying twice as fast and high as the propellerdriven transports they were replacing—promised even higher speeds in 

coming years. At the same time, the perceived challenge to America’s 

technological superiority implied by the Soviet Union’s early space 

triumphs inspired a willingness to pursue ambitious new aerospace 

ventures. One of these was the Supersonic Commercial Air Transport 

(SCAT). This program was further motivated by competition from 

22. “Jet Breaks 3 Records—and Many Windows,” Los Angeles Times, Mar. 6, 1962, p. 1. In reality, most of the damage was done while accelerating after the refuelings.
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Figure 2. Cover of an Air Force pamphlet for sonic boom claim investigators. USAF.

Britain and France to build an airliner that was expected to dominate the 

future of mid- and long-range commercial aviation.23

23. For the definitive account of political and economic aspects of the SST and subsequent programs (as 

well as many technical details), see Erik M. Conway, High-Speed Dreams: NASA and the Technopolitics 

of Supersonic Transportation, 1945–1999 (Baltimore: Johns Hopkins, 2005), pp. 27–45 cited here. For 

an earlier study by an insider, see F. Edward McLean, “Supersonic Cruise Technology,” NASA Special 

Publication (SP) 472 (Washington, DC: GPO, 1985). For an account focused on its political aspects, 

see Mel Howitch, Clipped Wings: The American SST Conflict (Cambridge: MIT, 1982).
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From SCAT Research to SST Development

The recently established FAA became the major advocate within the U.S. 

Government for a supersonic transport, with key personnel at three of 

the NACA’s former laboratories eager to help with this challenging new 

program. The Langley Research Center in Hampton, VA, (the NACA’s 

oldest and largest lab) and the Ames Research Center at Moffett Field in 

Sunnyvale, CA, both had airframe design expertise and facilities, while 

the Lewis Research Center in Cleveland, OH, specialized in the kind of 

advanced propulsion technologies needed for supersonic cruise.

The strategy for developing the SCAT depended heavily on leveraging 

technologies being developed for another Air Force bomber—one much 

larger, faster, and more advanced than the B-58. This would be the revolutionary B-70, designed to cruise several thousand miles at speeds of 

Mach 3. NACA experts had been helping the Air Force plan this giant 

intercontinental bomber since the mid-1950s (with aerodynamicist Alfred 

Eggers of the Ames Laboratory conceiving the innovative design for it 

to ride partially on compression lift created by its own supersonic shock 

waves). North American Aviation won the B-70 contract in 1958, but 

the projected expense of the program and advances in missile technology led President Dwight Eisenhower to cancel all but one prototype in 

1959. The administration of President John Kennedy eventually approved 

production of two XB-70As. Their main purpose would be to serve as 

Mach 3 testbeds for what had become known simply as the Supersonic 

Transport (SST). NASA continued to refer to design concepts for the SST 

using the older acronym for Supersonic Commercial Air Transport. By 

1962, these concepts had been narrowed down to three Langley designs 

(SCAT-4, SCAT-15, and SCAT-16) and one from Ames (SCAT-17). These 

became the baselines for industry studies and SST proposals.24

Even though Department of Defense resources (especially the Air 

Force’s) would be important in supporting the SST program, the aerospace industry made it clear that direct federal funding and assistance 

would be essential. Thus research and development (R&D) of the SST 

became a split responsibility between the Federal Aviation Agency and 

the National Aeronautics and Space Administration—with NASA conducting and sponsoring the supersonic research and the FAA in charge 

24. McLean, Supersonic Cruise Technology, pp 35–46; Joseph R. Chambers, Innovation in Flight; 

“Research of the NASA Langley Research Center on Revolutionary Concepts for Aeronautics,” NASA 

SP-2005-4539, pp. 25–28.
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of the SST’s overall development. The first two leaders of the FAA, retired 

Lt. Gen. Elwood R. “Pete” Quesada (1958–1961) and Najeeb E. Halaby 

(1961–1965), were both staunch proponents of producing an SST, as to 

a slightly lesser degree was retired Gen. William F. “Bozo” McKee (1965–

1968). As heads of an independent agency that reported directly to the 

president, they were at the same level as NASA Administrators T. Keith 

Glennan (1958–1961) and James Beggs (1961–1968). The FAA and NASA 

administrators, together with Secretary of Defense Robert McNamara 

(somewhat of a skeptic on the SST program), provided interagency oversight and comprised the Presidential Advisory Committee (PAC) for the 

SST established in April 1964. This arrangement lasted until 1967, when 

the Federal Aviation Agency became the Federal Aviation Administration 

under t he new Department of Transportation, whose secretary became 

responsible for the program.25

Much of NASA’s SST-related research involved advancing the stateof-the-art in such technologies as propulsion, fuels, materials, and 

aerodynamics. The latter included designing airframe configurations 

for sustained supersonic cruise at high altitudes, suitable subsonic 

maneuvering in civilian air traffic patterns at lower altitudes, safe takeoffs and landings at commercial airports, and acceptable noise levels—

to include the still-puzzling matter of sonic booms. 

Dealing with the sonic boom entailed a multifaceted approach: (1) 

performing flight tests to better quantify the fluid dynamics and atmospheric physics involved in generating and propagating shock waves, 

as well as their effects on structures and people; (2) conducting community surveys to gather public opinion data on sample populations 

exposed to booms; (3) building and using acoustic simulators to further evaluate human and structural responses in controlled settings; 

(4) performing field studies of possible effects on animals; (5) evaluating various aerodynamic configurations in wind tunnel experiments; 

and (6) analyzing flight test and wind tunnel data to refine theoretical 

constructs and mathematical models for lower-boom aircraft designs. 

Within NASA, the Langley Research Center was a focal point for sonic 

25. FAA Historical Chronology, 1926–1996, http://ww.faa.gov/about/media/b-chron.pdf. 

For Quesada’s role, see Stuart I. Rochester, Takeoff at Mid-Century: Federal Civil Aviation Policy in 

the Eisenhower Years, 1953–1961 (Washington, DC: GPO for FAA, 1976). For the activism of 

Halaby and the demise of the SST after his departure, see Richard J. Kent, Jr., Safe, Separated, and 

Soaring: A History of Civil Aviation Policy, 1961–1972 (Washington, DC: GPO for FAA, 1980). 
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boom studies, with the Flight Research Center (FRC) at Edwards AFB 

conducting many of the supersonic tests.26

Although the NACA, especially at Langley and Ames, had been doing 

research on supersonic flight since World War II, none of its technical 

reports (and only one conference paper) published through 1957 dealt 

directly with sonic booms.27 That situation began to change when Langley’s 

long-time manager and advocate of supersonic programs, John P. Stack, 

formalized the SCAT venture in 1958. During the next year, three Langley 

employees whose names would become well known in the field of sonic 

boom research began publishing NASA’s first scientific papers on the subject. These were Harry W. Carlson, a versatile supersonic aerodynamicist, 

Harvey H. Hubbard, chief of the Acoustics and Noise Control Division, 

and Domenic J. Maglieri, a young engineer who became Hubbard’s top 

sonic boom specialist. Carlson would tend to focus on wind tunnel experiments and sonic boom theory, while the other two men specialized in 

planning and monitoring field tests, then analyzing the data collected.28

These research activities began to expand under the new pro-SST Kennedy 

Administration in 1961. After the president formally approved development of the supersonic transport in June 1963, sonic boom research took 

off. Langley’s experts, augmented by NASA contractors and grantees, published 26 papers on sonic booms just 3 years later.29

Supersonic Flight Tests and Surveys

The systematic sonic boom testing that NASA began in 1958 would 

exponentially expand the heretofore largely theoretical and anecdotal 

26. NASA’s HSFRC became the FRC in 1959. For an overall summary of Langley’s supersonic 

activities, see Chambers, Innovations in Flight, ch. 1, “Supersonic Civil Aircraft: The Need for 

Speed,” pp. 7–70.

27. Based on author’s review of Section 7.4, “Noise, Aircraft” in volumes of the Index of NACA 

Technical Publications (Washington DC: NACA Division of Research Information) covering the years 

1915–1957. 

28. Telephone interview, Domenic Maglieri by Lawrence Benson, Feb. 6, 2009.

29. A.B. Fryer, et al., “Publications in Acoustics and Noise Control from the NASA Langley 

Research Center during 1940–1976,” NASA TM-X-74042, July 1977. The following abbreviations are used for NASA publications cited in the notes: Conference Publication (CP), Contractor 

Report (CR), Reference Publication (RP), Special Publication (SP), Technical Memorandum (TM), 

formerly classified Tech Memo (TM-X), Technical Note (TN), Technical Paper (TP), and Technical 

Report (TR). Bibliographic data and often full text copies can be accessed through the NASA 

Technical Reports Server (NTRS), http://ntrs.nasa.gov/search.jsp.
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knowledge about sonic booms with a vast amount of “real world” 

data. The new information would make possible increasingly sophisticated experiments and provide feedback for checking and refining theories and mathematical models. Because of the priority bestowed on sonic 

boom research by the SST program and the numerous types of aircraft 

then available for creating booms (including some faster than anything 

flying today), the data and findings from the tests conducted in the 1960s 

are still of significant value in the 21st century.30

The Langley Research Center (often referred to as NASA Langley) 

served as the Agency’s “team leader” for supersonic research. Langley’s 

acoustics specialists conducted NASA’s initial sonic boom tests in 1958 

and 1959 at the Wallops Island Station on Virginia’s isolated Delmarva 

Peninsula. During the first year, they used six sorties by NASA F-100 and 

F-101 fighters, flying at speeds between Mach 1.1 and 1.4 and altitudes 

from 25,000 to 45,000 feet, to make the first good ground recordings and 

measurements of sonic booms for steady, level flights (the kind of profile 

a future airliner would fly). Observers judged some of the booms above 

1.0 psf to be objectionable, likening them to nearby thunder, and a sample 

plate glass window was cracked by one plane flying at 25,000 feet. The 1959 

test measured shock waves from 26 flights of a Chance Vought F8U-3 (a 

highly advanced prototype based on the Navy’s Crusader fighter) at speeds 

up to Mach 2 and altitudes up to 60,000 feet. A B-58 from Edwards AFB 

also made two supersonic passes at 41,000 feet. Boom intensities from 

these higher altitudes seemed to be tolerable to observers, with negligible 

increases in measured overpressures between Mach 1.4 and 2.0. These 

results were, however, very preliminary.31

In July 1960, NASA and the Air Force conducted Project Little Boom 

at a bombing range north of Nellis AFB, NV, to measure the effects on 

structures and people of extremely powerful sonic booms. F-104 and F-105 

fighters flew slightly over the speed of sound (Mach 1.09 to 1.2) at altitudes 

30. For a chronological summary of selected projects during first decade, see Johnny M. Sands, 

“Sonic Boom Research (1958–1968),” FAA, Nov. 1968, Defense Technical Information Center 

(DTIC) document AD 684806. 

31. Domenic J. Maglieri, Harvey H. Hubbard, and Donald L. Lansing, “Ground Measurements 

of the Shock-Wave Noise from Airplanes in Level Flight at Mach Numbers to 1.4 and Altitudes to 

45,000 Feet,” NASA TN-D-48, Sept. 1959; Lindsay J. Lina and Domenic J. Maglieri, “Ground 

Measurements of Airplane Shock-Wave Noise at Mach Numbers to 2.0 and at Altitudes to 

60,000 Feet,” NASA TN-D-235, Mar. 1960.
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as low as 50 feet above ground level. There were more than 50 incidents 

of sample windows being broken at 20 to 100 psf, but only a few possible breakages below 20 psf, and no physical or psychological harm to volunteers exposed to overpressures as high as 120 psf.32 At Indian Springs, 

Air Force fighters flew supersonically over an instrumented C-47 transport from Edwards, both in the process of landing and on the ground. 

Despite 120 psf overpressures, there was only very minor damage when 

on the ground and no problems in flight.33 Air Force fighters once again 

would test powerful sonic booms in 1965 in support of Joint Task Force 

2 at Tonopah, NV. The strongest sonic boom ever recorded, 144 psf, was 

generated by an Air Force F-4E Phantom II flying Mach 1.26 at 95 feet.34

In late 1960 and early 1961, NASA and AFFTC followed up on Little 

Boom with Project Big Boom. B-58 bombers made 16 passes flying 

Mach 1.5 at altitudes of 30,000 to 50,000 feet over arrays of sensors, 

which measured a maximum overpressure of 2.1 psf. Varying the bomber’s weight from 82,000 to 120,000 pounds provided the first hard data on 

how an aircraft’s weight and related lift produced higher over-pressures 

than existing theories based on volume alone would indicate.35

Throughout the 1960s, Edwards Air Force Base—with its unequaled 

combination of Air Force and NASA expertise, facilities, instrumentation, airspace, emergency landing space, and types of aircraft—hosted 

the largest number of sonic boom tests. NASA researchers from Langley’s 

Acoustics Division spent much of their time there working with the Flight 

Research Center in a wide variety of flight experiments. The Air Force 

Flight Test Center usually participated as well. 

In an early test in 1961, Gareth Jordan of the FRC led an effort to 

collect measurements from F-104s and B-58s flying at speeds of Mach 

1.2 to 2.0 over sensors located along Edward AFB’s supersonic corridor 

32. Maglieri, Vera Huckel, and Tony L. Parrott, “Ground Measurements of Shock-Wave Pressure 

for Fighter Airplanes Flying at Very Low Altitudes . . .,” NASA TN-D-3443, July 1966 (superseded 

classified TMX-611, 1961).

33. Gareth H. Jordan, “Flight Measurements of Sonic Booms and Effects of Shock Waves on Aircraft,” in Society of Experimental Test Pilots Quarterly Review, vol. 5, No. 1 (1961), pp. 117–131, 

presented at SETP Supersonic Symposium, Sept. 29, 1961.

34. John O. Powers, J.M. Sands, and Maglieri, “Survey of United States Sonic Boom Overflight 

Experimentation,” NASA TM-X-66339, May 1969, p. 5; USAF Fact Sheet, “Sonic Boom,” Oct. 2005, 

http://www.af.mil/factsheets/fsID=184; Telephone interview, Maglieri by Benson, Mar. 19, 2009.

35. Maglieri and Hubbard, “Ground Measurements of the Shock-Wave Noise from Supersonic 

Bomber Airplanes in the Altitude Range from 30,000 to 50,000 Feet,” NASA TN-D-880, July 1961. 
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and at Air Force Plant 42 in Palmdale, about 20 miles south. Most of 

the Palmdale measurements were under 1.0 psf, which the vast majority 

of people surveyed there and in Lancaster (where overpressures tended 

to be somewhat higher) considered no worse than distant thunder. But 

there were some exceptions.36

Other experiments at Edwards in 1961 conducted by Langley personnel with support from the FRC and AFFTC contributed a variety 

of new information. With help from a tethered balloon, they made the 

first good measurements of atmospheric effects, showing that air turbulence in the lower atmosphere (known as the boundary layer) significantly affected wave shape and overpressure. They also gathered the 

first data on booms from very high altitudes. Using an aggressive flight 

profile, AFFTC’s B-58 crew managed to zoom up to 75,000 feet—25,000 

feet higher than the bomber’s normal cruising altitude and 15,000 feet 

over its design limit! The overpressures measured from this high altitude 

proved stronger than predicted (not a promising result for the planned 

SST). Much lower down, fighter aircraft performed accelerating and 

turning maneuvers to generate the kind of acoustical rays that amplified shock waves and produced multiple booms and super booms. The 

various experiments showed that a combination of atmospheric conditions, altitude, speed, flight path, aircraft configuration, and sensor 

location determined the shape of the pressure signatures.37

Of major significance for future boom minimization efforts, NASA 

also began making in-flight shock wave measurements. The first of these, 

at Edwards in 1960, had used an F-100 with a sensor probe to measure 

supersonic shock waves from the sides of an F-100, F-104, and B-58, as 

well as from F-100s speeding past with only 100 feet of separation. The 

data confirmed Whitham’s overall theory, with some discrepancies. In 

early 1963, an F-106 equipped with a sophisticated new sensor probe 

designed at Langley flew seven sorties both above and below a B-58 at 

speeds of Mach 1.42 to 1.69 and altitudes of approximately 40,000 to 

50,000 feet. The data gathered confirmed Walkden’s theory that lift as 

well as volume increases peak shock wave pressures. As indicated by 

36. Jordan, “Flight Measurements of Sonic Booms.”

37. Ibid.; Maglieri and Donald L. Lansing, “Sonic Booms from Aircraft in Maneuvers,” NASA TND-2370, July 1964; Hubbard, et al., “Ground Measurements of Sonic-Boom Measurements for the 

Altitude Range of 10,000 to 75,000 Feet,” NASA TR-R-198, July 1964. (Both reports were based 

on the tests in 1961.) 
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Figure 3, analysis of the readings also found that the bow and tail shock 

waves spread farther apart as they flowed from the B-58 and showed 

how the multiple or “saw tooth” shock waves produced by the rest of an 

airplane’s structure (e.g., fuselage, canopy, wings, engines, nacelles, etc.) 

merged with the stronger bow and tail waves until—at a distance of between 

50 and 90 body lengths—they began to coalesce into the classic N-shaped 

signature.38 This marked a major milestone in sonic boom research.

One of the most publicized and extended flight test programs at 

Edwards had begun in 1959 with the first launch from a B-52 of the fastest aircraft ever flown: the rocket-propelled X-15. Three of these legendary 

aerospace vehicles expanded the envelope and gathered data on supersonic and hypersonic flight for the next 8 years. Although the X-15 was 

not specifically dedicated to sonic boom tests, the Flight Research Center 

did begin placing microphones and tape recorders under the X-15s’ flight 

tracks in the fall of 1961 to gather boom data. FRC researchers much later 

reported on the measurements of these sonic booms, made at speeds of 

Mach 3.5 and Mach 4.8.39

For the first few years, NASA’s sonic boom tests occurred in relative isolation within military airspace in the desert Southwest or over 

Virginia’s rural Eastern Shore. A future SST, however, would have to 

fly over heavily populated areas. Thus, from July 1961 through January 

1962, NASA, the FAA, and the Air Force carried out the Community and 

Structural Response Program at St. Louis, MO. In Operation Bongo, the 

Air Force sent B-58 bombers on 76 supersonic training flights over the city 

at altitudes from 31,000 to 41,000 feet, announcing them as routine SAC 

radar bomb-scoring missions. F-106 interceptors flew 11 additional flights 

at 41,000 feet. Langley personnel installed sensors on the ground, which 

measured overpressures up to 3.1 psf. Investigators from Scott AFB, IL, or 

for a short time, a NASA-contracted engineering firm, responded to dam38. Harriet J. Smith, “Experimental and Calculated Flow Fields Produced by Airplanes Flying at 

Supersonic Speeds,” NASA TN-D-621, Nov. 1960; J.F. Bryant, Maglieri, and V.S. Richie, “In-Flight 

Shock-Wave Measurements Above and Below a Bomber Airplane at Mach Numbers from 1.42 to 

1.69,” NASA TN-D-1968, Oct. 1963.

39. NASA Flight Research Center, “X-15 Program” [monthly report], Sept. 1961, Dryden archive, 

File LI-6-10A-13 (Peter Merlin assisted the author in finding this and other archival documents.); Karen 

S. Green and Terrill W. Putnam, “Measurements of Sonic Booms Generated by an Airplane Flying at 

Mach 3.5 and 4.8,” NASA TM-X-3126, Oct. 1974. (Since hypersonic speeds were not directly relevant for the SST, a formal report was delayed until NASA began planning reentry flights for the Space 

Shuttle.) For a history of the X-15 program, see Hallion and Gorn, On the Frontier, pp. 101–125.
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Figure 3. In-flight sonic boom signatures of B-58 at Mach 1.6. USAF.

age claims, finding some possibly legitimate minor damage in about 20 

percent of the cases. Repeated interviews with more than 1,000 residents 

found 90 percent were at least somewhat affected by the booms and about 

35 percent were annoyed. Scott AFB (a long-distance phone call from St. 

Louis) received about 3,000 complaints during the test and another 2,000 

in response to 74 sonic booms in the following 3 months. The Air Force 

eventually approved 825 claims for $58,648. These results served as a warning that repeated sonic booms could pose an issue for SST operations.40

To obtain more definitive data on structural damage, NASA in 

December 1962 resumed tests at Wallops Island using various sample 

buildings. Air Force F-104s and B-58s and Navy F4H Phantom IIs flew 

at altitudes from 32,000 to 62,000 feet, creating overpressures up to3 

psf. Results indicated that cracks to plaster, tile, and other brittle materials triggered by sonic booms occurred in spots where the materials 

were already under stress (a finding that would be repeated in later more 

comprehensive tests).41

40. Charles W. Nixon and Hubbard, “Results of the USAF–NASA–FAA Flight Program to Study 

Community Response to Sonic Booms in the Greater St. Louis Area,” NASA TN-D-2705, May 

1965; Clark, et al., “Studies of Sonic Boom Damage,” NASA CR-227, May 1965.

41. Sands, “Sonic Boom Research (1958–1968),” p. 3.
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In February 1963, NASA, the FAA, and the USAF conducted Project 

Littleman at Edwards AFB to see what happened when two specially 

instrumented light aircraft were subjected to sonic booms. F-104s made 

23 supersonic passes at distances as near as 560 feet from a little Piper 

Colt and a 2-engine Beech C-45, creating overpressures up to 16 psf. 

Their responses were “so small as to be insignificant”—dismissing one 

possible concern about SST operations.42

The St. Louis survey had left many unanswered questions on public opinion. To learn more, the FAA’s Supersonic Transport Development 

Office, with support from NASA Langley and the USAF (including Tinker 

AFB), next conducted the Oklahoma City Public Reaction Study from 

February through July 1964. This was a much more intensive and systematic test. In an operation named Bongo II, B-58s, F-104s, F-101s, 

and F-106s were called upon to deliver sonic booms between 1.0 and 

2.0 psf, 8 times per day, 7 days a week, for 26 weeks, with another 13 

weeks of followup activities. The aircraft flew a total of 1,253 supersonic 

flights at Mach 1.2 to 2.0 and altitudes between 21,000 and 50,000 feet. 

The FAA (which had a major field organization in Oklahoma City) 

instrumented nine control houses scattered throughout the metropolitan area with various sensors to measure structural effects, while experts 

from Langley instrumented three houses and set up additional sensors 

throughout the area to record overpressures, wave patterns, and meteorological conditions. The National Opinion Research Center at the 

University of Chicago interviewed a sample of 3,000 adults three times 

during the study. By the end of the test, 73 percent of those surveyed felt 

that they could live with the number and strength of the booms experienced, and 27 percent would not accept indefinite booms at the level 

tested. Forty percent believed that they caused some structural damage 

(even though the control houses showed no significant effects). Analysis 

of the shock wave patterns by NASA Langley showed that a small number 

of overpressure measurements were significantly higher than expected, 

indicating probable atmospheric influences, including heat rising from 

urban landscapes. One possible result was the breakage of almost 150 

windows in the city’s two tallest buildings early in the test.43

42. Maglieri and Garland J. Morris, “Measurement of Response of Two Light Airplanes to Sonic 

Booms,” NASA TN-D-1941, Aug. 1963.

43. D.A. Hilton, Maglieri, and R. Steiner, “Sonic-Boom Exposures during FAA Community Response 

Studies over a 6-Month Period in the Oklahoma City Area,” NASA TN-D-2539, Dec. 1964.
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The Oklahoma City study added to the growing knowledge about 

sonic booms and their acceptance by the public at the cost of negative publicity for the FAA. In view of the reactions to the St Louis and 

Oklahoma City tests by much of the public and some politicians, plans 

for another extended sonic boom test over a different city, including 

flights at night, never materialized.44

The FAA and Air Force conducted the next series of tests from 

November 1964 to February 1965 in a much less populated place: the 

remote Oscura camp in the vast White Sands Missile Range of New 

Mexico, where 21 structures of various types and ages with a variety 

of plaster, windows, and furnishings were studied for possible damage. F-104s from nearby Holloman AFB and B-58s from Edwards generated 1,494 booms producing overpressures from 1.6 to 19 psf. The 

680 sonic booms at 5.0 psf caused no real problems, but those above 

7.9 psf revealed varying degrees of damage to glass, plaster, tile, and 

stucco already in vulnerable condition. A parallel study of several 

thousand incubated chicken eggs showed no reduction in hatchability, and audiology tests on 20 personnel subjected daily to the booms 

showed no hearing impairment.45

Before the White Sands test ended, NASA Langley personnel began 

collecting boom data from a highly urbanized setting in winter weather. 

During February and March 1965, they recorded data at five ground 

stations as B-58 bombers flew 22 training missions in a corridor over 

downtown Chicago at speeds of Mach 1.2 to 1.66 and altitudes from 

38,000 to 48,000 feet. The results showed how amplitude and wave 

shape varied widely depending upon atmospheric conditions. These 22 

flights and 27 others resulted in the Air Force approving 1,442 of 2,964 

damage claims for $114,763.46

Also in March 1965, the FAA and NASA, in cooperation with the 

U.S. Forest Service, studied the effect on hazardous mountain snow 

packs in the Colorado Rockies of Air Force fighters creating boom overpressures up to 5.0 psf. Because of stable snow conditions, none of 

these created an avalanche. Interestingly enough, in the early 1960s the 

44. Conway, High-Speed Dreams, pp. 121–122.

45. Thomas H. Higgins, “Sonic Boom Research and Design Considerations in the Development of 

a Commercial Supersonic Transport,” JASA, vol. 39, no. 5, pt. 2 (Nov. 1966), pp. 526–531.

46. David. A. Hilton, Vera Huckel, and Maglieri, “Sonic Boom Measurements during Bomber Training Operations in the Chicago Area,” NASA TN-D-3655, Oct. 1966.
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National Park Service had tried to use newly deployed F-106s at Geiger 

Field, WA, to create controlled avalanches in Glacier National Park 

(Project “Safe Slide”), but presumably found traditional artillery 

fire more suitable.47

From the beginning of the SST program, the aircraft most desired 

for experiments was, of course, the North American XB-70 Valkyrie. 

The first of the giant testbeds (XB-70-1) arrived at Edwards AFB in 

September 1964, and the better performing and better instrumented second aircraft (XB-70-2) arrived in July 1965. With a length of 186 feet, a 

wingspan of 105 feet, and a gross weight of about 500,000 pounds, the 

six-engine giant was less than two-thirds as long as some of the later 

SST concepts, but it was the best real-life surrogate available. 

Even during the initial flight envelope expansion by contractor and 

AFFTC test pilots, the Flight Research Center began gathering sonic 

boom data, including direct comparisons of its shock waves with those 

of a B-58 flying only 800 feet behind.48 Using an array of microphones 

and recording equipment at several ground stations, NASA researchers eventually built a database of boom signatures from 39 flights made 

by the XB-70s (10 with B-58 chase planes), from March 1965 through 

May 1966.49 Because “the XB-70 is capable of duplicating the SST 

flight profiles and environment in almost every respect,” the FRC was 

looking forward to beginning its own experimental research program 

using the second Valkyrie on June 15, 1966, with sonic boom testing 

listed as the first priority.50

On June 8, however, XB-70-2 crashed on its 47th flight as the result 

of an infamous midair collision that killed two pilots and gravely injured 

47. Histories of the 4700 Air Defense Wing, Jan.–Mar. and Apr.–June 1960, IRIS abstracts; History of the 84th Fighter Group, Jan.–Dec. 1961, IRIS abstract; Telephone interview, Maglieri by 

Benson, Mar. 13, 2009.

48. William H. Andrews, “Summary of Preliminary Data Derived from the XB-70 Airplanes,” NASA 

TM-X-1240, June 1966, pp. 11–12. Despite being 3.5 times heavier than the B-58, the XB-70’s 

bow wave proved to be only slightly stronger.

49. Maglieri, et al., “A Summary of XB-70 Sonic Boom Signature Data, Final Report,” NASA 

CR-189630, Apr. 1992. Until this report, the 1965–1966 findings were filed away unpublished. 

The original oscillographs were also scanned and digitized at this time for use in the High-Speed 

Research Program.

50. FRC, “NASA XB-70 Flight Research Program,” Apr. 1966, Dryden archive, File L2-4-4D-3, p. 

10 quoted. See also C.M. Plattner, “XB-70A Flight Research: Phase 2 to Emphasize Operational 

Data,” Aviation Week, June 13, 1966, pp. 60–62.
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An XB-70 taking off at Edwards AFB accompanied by a B-58 chase plane in the mid-1960s, 

when both were used for sonic boom research. North American.

a third. Despite this tragic setback to the test program, the less capable 

XB-70-1 (which underwent modifications until November) eventually 

proved useful for many purposes. After 6 months of joint AFFTC/FRC 

operations, including the boom testing described below, the plane was 

turned over full time to NASA in April 1967 after 60 Air Force flights. 

The FRC, with a more limited budget, then used the Valkyrie for 23 more 

test missions until February 1969, when the unique aircraft was retired 

to the USAF Museum in Dayton, OH.51 All told, NASA acquired sonic 

boom measurements from 51 of the 129 total flights made by the XB-70s, 

using two ground stations on Edwards AFB, one at nearby Boron, and 

two in Nevada.52 This data proved to be of great value in the future.

The loss of one XB-70 and retirement of the other from supersonic 

testing was made somewhat less painful by the availability of another 

smaller but even faster product of advanced aviation technology: the 

Lockheed YF-12 and its cousin, the SR-71—both nicknamed Blackbirds. 

On May 1, 1965, shortly after arriving at Edwards, a YF-12A set nine 

world records, including a closed-course speed of 2,070 mph (Mach 3.14) 

51. NASA Dryden Fact Sheet, “XB-70,” http://www.nasa.gov/centers/dryden/new/FactSheets/

FS-084-DFRC_prt.htm; Hallion and Gorn, On the Frontier, pp. 176–185, 421.

52. Maglieri, “Summary of XB-70 Sonic Boom,” pp. 4–5.
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and a sustained altitude of 80,257 feet. Four of that day’s five flights also 

yielded sonic boom measurements. At speeds of Mach 2.6 to 3.1 and altitudes of 60,000 to 76,500 feet above ground level, overpressures varied 

from 1.2 to 1.7 psf depending on distance from the flight path. During 

another series of flight tests at slower speeds and lower altitudes, overpressures up to 5.0 psf were measured during accelerations after having slowed to refuel. These early results proved consistent with previous 

B-58 data.53 Data gathered over the years from ground arrays measuring the sonic signatures from YF-12s, XB-70s, B-58s, and smaller aircraft flying at various altitudes also showed that the lateral spread of 

a boom carpet (without the influence of atmospheric variables) could 

be roughly equated to 1 mile for every 1,000 feet of altitude, with the 

N-signatures become more rounded with distance until degenerating 

into the approximate shape of a sine wave.54

Although grateful to benefit from the flights of AFFTC’s Blackbirds, 

the FRC wanted its own YF-12 or SR-71 for supersonic research. It 

finally gained the use of two YF-12s through a NASA–USAF memorandum of understanding signed in June 1969, paying for operations with 

funding left over from termination of the X-15 and XB-70 programs.55

In the fall of 1965, with public acceptance of sonic booms becoming a significant public and political issue, the White House Office of 

Science and Technology established the National Sonic Boom Evaluation 

Office (NSBEO) under an interagency Coordinating Committee on Sonic 

Boom Studies. The new organization, which was attached to Air Force 

Headquarters for administrative purposes, planned a comprehensive series 

of tests known as the National Sonic Boom Evaluation Program, to be conducted primarily at Edwards AFB. NASA (in particular the Flight Research 

and Langley Centers) would be responsible for test operations and data 

collection, with the Stanford Research Institute hired to help analyze the 

findings.56 After careful preparations (including specially built structures 

53. R.T. Klinger, “YF-12A Flight Test Sonic Boom Measurements,” Lockheed Advanced Development 

Projects Report SP-815, June 1, 1965, Dryden archive, File LI-4-10A-1.

54. John O. Powers, J.M. Sands, and Maglieri, “Survey of United States Sonic Boom Overflight 

Experimentation,” NASA TM-X-66339, May 1969, pp. 9, 12–13.

55. Peter W. Merlin, From Archangel to Senior Crown: Design and Development of the Blackbird (Reston, VA: AIAA, 2008), pp. 106–107, 116–118, 179; Hallion and Gorn, On the Frontier, p. 187.

56. NSBEO, “Sonic Boom Experiments at Edwards Air Force Base; Interim Report” (prepared under contract by Stanford Research Institute), pp. 1–2, (hereinafter cited as SRI, “Edwards AFB Report”). For political 

and bureaucratic background on the NSBEO, see Conway, High-Speed Dreams, pp. 122–123. 
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and extensive sensor and recording arrays), the National Sonic Boom 

Evaluation began in June 1966. Its main objectives were to address the 

many issues left unresolved from previous tests. Unfortunately, the loss of 

XB-70-2 on June 8 forced a 4-month break in the test schedule, with the 

limited events completed in June designated Phase I. The second phase 

began in November, when XB-70-1 returned to flight status, and lasted 

into January 1967. A total of 367 supersonic missions were flown by 

XB-70s, B-58s, YF-12s, SR-71s, F-104s, and F-106s during the two phases. 

These were supplemented by 256 subsonic flights by KC-135s, WC-135Bs, 

C-131Bs, and Cessna 150s. In addition, the Goodyear blimp Mayflower 

was used in the June phase to measure sonic booms at 2,000 feet.57

By the end of testing, the National Sonic Boom Evaluation had obtained 

new and highly detailed acoustic and seismic signatures from all the different supersonic aircraft in various flight profiles during a variety of atmospheric conditions. The data from 20 XB-70 flights at speeds from Mach 1.38 

to 2.94 were to be of particular long-term interest. For example, Langley’s 

sophisticated nose probe used for the pioneering in-flight flow-field measurements of the B-58 in 1963 was installed on one of the FRC’s F-104s to 

do the same for the XB-70. Comparison of data between blimp and ground 

sensors and variations between the summer and winter tests confirmed the 

significant influence that atmospheric conditions, such as turbulence and 

convective heating near the surface, have on boom propagation. 58 Also, 

the evaluation provided an opportunity to gather data on more than 1,500 

sonic boom signatures created during 35 flights by the recently available 

SR-71s and YF-12s at speeds up to Mach 3.0 and altitudes up to 80,000 feet.59

Some of the findings portended serious problems for planned SST 

operations. The program obtained responses from several hundred volunteers, both outdoors and in houses, to sonic booms of different intensities produced by each of the supersonic aircraft. The time between 

57. SRI, “Edwards AFB Report,” p. 9.

58. Maglieri, et al., “Summary of Variations of Sonic Boom Signatures Resulting from Atmospheric 

Effects,” Feb. 1967, and “Preliminary Results of XB-70 Sonic Boom Field Tests During National 

Sonic Boom Evaluation Program,” Mar. 1967, Annex C-1 and C-2, in SRI, “Edwards AFB Report;” 

H.H. Hubbard and D.J. Maglieri, “Sonic Boom Signature Data from Cruciform Microphone Array 

Experiments during the 1966–1967 EAFB National Sonic Boom Evaluation Program,” NASA TN-D6823, May 1972.

59. SRI, “Edwards AFB Report,” pp. 17–20, Annexes C-F; Maglieri, et al., “Sonic Boom Measurements for SR-71 Aircraft Operating at Mach Numbers to 3.0 and Altitudes to 24834 Meters,” 

NASA TN-D-6823, Sept. 1972.
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the peak overpressure of the bow and tail waves for aircraft at high altitudes ranged from about 0.1 second for the F-104, 0.2 second for the 

B-58, and 0.3 second for the XB-70. The respondents also compared 

sonic booms to the jet engine noise of subsonic aircraft. Although data 

varied for each of the criteria measured, significant minorities tended to 

find the booms either “just acceptable” or unacceptable, with the “sharper” 

N-wave signature from the lower flying F-104 more annoying outdoors 

than the more rounded signatures from the larger aircraft, which had to fly 

at higher altitudes to create the same overpressure. Other factors included 

the frequency, time of day or night, and type of boom signature. Correlating 

how the subjects responded to jet noise (measured in decibels) and sonic 

booms (normally measured in psf), the SRI researchers used the perceived 

noise decibel (PNdB) level to assess how loud booms seem to human ears.60

Employing sophisticated sensors, civil engineers measured the physical effects on houses and a building with a large interior space (the base 

bowling alley) to varying degrees of booms created by F-104s, B-58s, and 

the XB-70. Of special concern for the SST, they found the XB-70’s elongated N-wave created more of the low frequencies that cause indoor vibrations, such as rattling windows (although less bothersome to observers 

outdoors). And although no significant harm was detected to the instrumented structures, 57 complaints of damage were received from residents 

in the surrounding area, and three windows were broken on base. Finally, 

monitoring by the Department of Agriculture detected no ill effects on 

farm animals in the area, although avian species (chickens, turkeys, etc.) 

reacted more than livestock did.61 The National Sonic Boom Evaluation 

remains the most comprehensive such test program yet conducted.

Later, in 1967, the opportunity for collecting additional survey data 

presented itself when the FAA and NASA learned that SAC was starting 

an extensive training program for its growing fleet of SR-71s. TRACOR, 

60. SRI, “Edwards AFB Report,” pp. 11–16, Annex B; K.D. Kryter, “Psychological Experiments on 

Sonic Booms Conducted at Edwards Air Force Base, Final Report,” (Menlo Park: SRI, 1968), summarized by Richard M. Roberds, “Sonic Boom and the Supersonic Transport,” Air University Review, 

vol. 22, No. 7 (July–Aug. 1971), pp. 25–33.

61. SRI, “Edwards AFB Report,” pp. 20–23, Annexes G and H; David Hoffman, “Sonic Boom 

Tests Fail to Win Any Boosters,” Washington Post, Aug. 3, 1967, p. A3; A.J. Bloom, et al. (SRI), 

“Response of Structures to Sonic Booms Produced by XB-70, B-58, and F-104 Aircraft . . . at Edwards 

Air Force Base, Final Report,” NSBEO 2-67, Oct. 1967; D.S. Findley, et al., “Vibration Responses of 

Test Structure No. 1 during the . . . National Sonic Boom Program,” NASA TM-X-72706, June 1975, 

and “Vibration Responses of Test Structure No. 2 . . . ,” NASA TM-X-72704, June 1975.
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Inc., of Austin, TX, which was already under contract to NASA doing 

surveys on airport noise, had its contract expanded in May 1967 to 

include public responses to the SR-71s’ sonic booms in Dallas, Los 

Angeles, Denver, Atlanta, Chicago, and Minneapolis. Between July 3 and 

October 2, Air Force SR-71s made 220 flights over these cities at high 

altitude, ranging from 5 over Atlanta to 60 over Dallas. The minority of 

sonic booms that were measured were almost all N-waves with overpressures from slightly less than 1.0 psf to 2.0 psf. Although the data from 

this impromptu test program were less than definitive, the overall findings (based on 6,375 interviews) were fairly consistent with the previous 

human response surveys. For example, after an initial dropoff, the level 

of annoyance with the booms tended to increase over time, and almost 

all those who complained were worried about damage. Among 15 different adjectives supplied to describe the booms (e.g., disturbing, annoying, 

irritating), the word “startling” was chosen much more than any other.62

Although the FRC and AFFTC continued their missions of supersonic 

flight-testing and experimentation at Edwards, what might be called 

the heroic era of sonic boom testing was drawing to a close. The FAA 

and the Environmental Science Services Administration (a precursor of 

the Environmental Protection Agency) did some sophisticated testing 

of meteorological effects at Pendleton, OR, from September 1968 until 

May 1970, using a dense grid of recently invented unattended recording equipment to measure random booms from SR-71s. On the other 

side of the continent, NASA and the Navy studied sonic booms during 

Apollo missions in 1970 and 1971.63

The most significant NASA testing in 1970 took place from August 

through October at the Atomic Energy Commission’s Jackass Flats test 

site in Nevada. In conjunction with the FAA and the National Oceanic 

and Atmospheric Administration (NOAA), NASA took advantage of the 

1,527-foot-tall BREN Tower (named for its original purpose, the “Bare 

Reactor Experiment Nevada” in 1962) to install a vertical array of 15 

microphones as well as meteorological sensors. (Until then, a 250-foot 

tower at Wallops Island had been the highest used in sonic boom testing.) During the summer and fall of 1970, the FRC’s F-104s from Edwards 

made 121 boom-generating flights to provide measurements of several 

62. TRACOR, Inc., “Public Reactions to Sonic Booms,” NASA CR-1665, Sept. 1970.

63. Hilton and Herbert R. Henderson documented the sonic boom measurements from the Apollo 15, 

16, and 17 missions in NASA TNs D-6950, D-7606, and D-7806, published from 1972 to 1974. 
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still poorly understood aspects of the sonic boom, especially the places, 

known mathematically as caustics, where nonlinear focusing of acoustical rays occurs. Frequently caused by speeds very near Mach 1 or by 

acceleration, they can result in U-shaped signatures with bow and tail 

wave overpressures strong enough to create super booms. The BREN 

Tower allowed such measurements to be made in the vertical dimension 

for the first time. This test resulted in definitive data on the formation 

and nature of caustics, information that would be valuable in helping 

pilots to avoid making focused booms.64

For all intents and purposes, the results of earlier testing and surveys had already helped to seal the fate of the SST before the reports on 

this latest test began coming in. Yet the data gathered from 1958 through 

1970 during the SST program contributed tremendously to the international aeronautical and scientific communities’ understanding of one of 

the most baffling and complicated aspects of supersonic flight. As Harry 

Carlson told the Nation’s top sonic boom scientists and engineers on the 

very same day of the last F-104 mission over Jackass Flats: “The importance 

of flight-test programs cannot be overemphasized. These tests have provided an impressive amount of high-quality data, which has been of great 

value in the verification of theoretical methods for the prediction of nominal overpressures and in the estimation from a statistical standpoint of 

the modifying influence of unpredictable atmospheric nonuniformities.”65

Laboratory Experiments and Sonic Boom Theory 

The rapid progress made in understanding the nature and significance 

of sonic booms during the 1960s resulted from the synergy among flight 

testing, wind tunnel experiments, psychoacoustical studies, theoretical 

refinements, and new computing capabilities. Vital to this process was 

the largely free exchange of information by NASA, the FAA, the USAF, 

the airplane manufacturers, academia, and professional organizations 

such as the American Institute of Aeronautics and Astronautics (AIAA) 

and the Acoustical Society of America (ASA). The sharing of information 

64. George T. Haglund and Edward J. Kane, “Flight Test Measurements and Analysis of Sonic 

Boom Phenomena Near the Shock Wave Extremity,” NASA CR-2167, Feb. 1973; Telephone 

interview, Maglieri by Benson, Mar. 13, 2009.

65. Harry W. Carlson, “Some Notes on the Present Status of Sonic Boom Prediction and Minimization Research,” Third Conference on Sonic Boom Research . . . Washington, DC, Oct. 29–30, 

1970, NASA SP-255, 1971, p. 395.
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even extended to potential rivals in Europe, where the Anglo-French 

Concorde supersonic airliner got off to a headstart on the more ambitious American program. 

Designing commercial aircraft has always required tradeoffs between 

speed, range, capacity, weight, durability, safety, and, of course, costs—

both for manufacturing and operations. Balancing such factors was especially challenging with an aircraft as revolutionary as the SST. Unlike 

with the supersonic military aircraft in the 1950s, NASA’s scientists and 

engineers and their partners in industry also had to increasingly consider the environmental impacts of their designs. At the Agency’s aeronautical Centers, especially Langley, this meant that aerodynamicists 

incorporated the growing knowledge about sonic booms in their equations, models, and wind tunnel experiments.

Harry Carlson of the Langley Center had conducted the first wind tunnel experiment on sonic boom generation in 1959. As reported in December, 

he tested seven models of various geometrical and airplane-like shapes at 

differing angles of attack in Langley’s original 4 by 4 supersonic wind tunnel at a speed of Mach 2.01. The tunnel’s relatively limited interior space 

mandated the use of very small models to obtain sonic boom signatures: 

about 2 inches in length for measuring shock waves at 8 body lengths distance and only about three-quarters inch for trying to measure them at 32 

body lengths (as close as possible to the “far field,” a distance where multiple shock waves coalesce into the typical N-wave signature). Although 

far-field data were problematic, the overall results correlated with existing 

theory, such as Whitham’s formulas on volume-induced overpressures and 

Walkden’s on those caused by lift.66 Carlson’s attempt to design one of the 

models to alleviate the strength of the bow shock was unsuccessful, but this 

might be considered NASA’s first attempt at boom minimization. 

The small size and extreme precision needed for the models, the 

disruptive effects of the assemblies needed to hold them, and the extra 

sensitivity required of pressure-sensing devices all limited a wind tunnel’s ability to measure the type of shock waves that would reach the 

ground from a full-sized aircraft. Even so, substantial progress continued, and the data served as a useful cross-check on flight test data and 

66. Carlson, “An Investigation of Some Aspects of the Sonic Boom by Means of Wind-Tunnel 

Measurements of Pressures about Several Bodies at a Mach Number of 2.01,” NASA TN-D-161, 

Dec. 1959. Carlson used Langley’s 4 by 4 Supersonic Pressure Wind Tunnel, completed in 1948, 

for most of his experiments.
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mathematical formulas.67 For example, in 1962 Carlson used a 1-inch 

model of a B-58 to make the first correlation of flight test data with 

wind tunnel data and sonic boom theory. Results proved that wind tunnel readings, with appropriate extrapolations, could be used with some 

confidence to estimate sonic boom signatures.68

Exactly 5 years after publishing results of the first wind tunnel sonic 

boom experiment, Harry Carlson was able to report, “In recent years, 

intensive research efforts treating all phases of the problem have served 

to provide a basic understanding of this phenomenon. The theoretical 

studies [of Whitham and Walkden] have resulted in the correlations 

with the wind tunnel data...and with the flight data.”69 As for minimization, wind tunnel tests of SCAT models had revealed that some configurations (e.g., the “arrow wing”) produced lower overpressures.70 Such 

possibilities were soon being explored by aerodynamicists in industry, 

academia, and NASA. They included Langley’s long-time supersonic 

specialist, F. Edward McLean, who had discovered extended near-field 

effects that might permit designing airframes for lower overpressures.71

Of major significance (and even more potential in the future), improved 

data reduction methods and numerical evaluations of sonic boom theory were being adapted for high-speed processing with new computer 

codes and hardware, such as Langley’s massive IBM 704. Using these new 

capabilities, Carlson, McLean, and others eventually designed the SCAT15F, an improved SST concept optimized for highly efficient cruise.72

In addition to reports and articles, NASA researchers presented 

findings from the growing knowledge about sonic booms in various 

67. For examples of these wind tunnel experiments, see Runyan, “Sonic Boom Capsule Summaries,” 

as well as the NTRS bibliographical database. 

68. Carlson, “Wind Tunnel Measurements of the Sonic-Boom Characteristics of a Supersonic Bomber 

Model and a Correlation with Flight-Test Ground Measurements,” NASA TM-X-700, July 1962.

69. Carlson, “Correlation of Sonic-Boom Theory with Wind Tunnel and Flight Measurements,” 

NASA TR-R-213, Dec. 1964. p. 1.

70. Evert Clark, “Reduced Sonic Boom Foreseen for New High-Speed Airliner,” New York Times, 

Jan. 1965, pp. 7, 12 (based on visit to NASA Langley).

71. F. Edward McLean, “Some Nonasymptotic Effects of the Sonic Boom of Large Airplanes,” 

NASA TN-D-2877, June 1965.

72. Carlson, “Correlation of Sonic-Boom Theory,” pp. 2–23. For an earlier status report on supersonic work at Langley and some at Ames, see William J. Alford and Cornelius Driver, “Recent 

Supersonic Transport Research,” Astronautics & Aeronautics, vol. 2, No. 9 (Sept. 1964), pp. 

26–37; Chambers, Innovation in Flight, pp. 32–34.
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meetings and professional symposia. One of the earliest took place 

September 17–19, 1963, when NASA Headquarters sponsored an SST 

feasibility studies review at the Langley Center—attended by Government, 

contractor, and airline personnel—that examined every aspect of the 

planned airplane. In a session on noise, Harry Carlson warned that “sonic 

boom considerations alone may dictate allowable minimum altitudes 

along most of the flight path and have indicated that in many cases the 

airframe sizing and engine selection depend directly on sonic boom.”73

On top of that, Harvey Hubbard and Domenic Maglieri discussed how 

atmospheric effects and community response to building vibrations 

might pose problems with the current SST sonic boom objectives (2 psf 

during acceleration and 1.5 psf during cruise).74

The conferees discussed various other technological challenges for 

the planned American SST, some indirectly related to the sonic boom 

issue. For example, because of frictional heating, an airframe covered 

largely with stainless steel (such as the XB-70) or titanium (such as the 

then-top secret A-12/YF-12) would be needed to cruise at Mach 2.7+ 

and over 60,000 feet, an altitude that many hoped would allow the sonic 

boom to weaken by the time it reached the surface. Manufacturing such 

a plane, however, would be much more expensive than building a Mach 

2.2 SST with aluminum skin, such as the Concorde.

Despite such concerns, the FAA had already released the SST request 

for proposals (RFP) on August 15, 1963. Thereafter, as explained by Ed 

McLean, “NASA’s role changed from one of having its own concepts evaluated by the airplane industry to one of evaluating the SST concepts 

of the airplane industry.”75 By January 1964, Boeing, Lockheed, North 

American, and their jet engine partners had submitted initial proposals. 

In retrospect, advocates of the SST were obviously hoping that technology would catch up with requirements before it went into production.

Although the SST program was now well underway, a growing awareness of the public response to booms became one factor in many that triagency (FAA–NASA–DOD) groups in the mid-1960s, including the PAC 

73. Carlson, “Configuration Effects on Sonic Boom,” Proceedings of NASA Conference on Supersonic-Transport Feasibility Studies and Supporting Research, Sept. 17–19, 1963 . . . Hampton, VA, 

NASA TM-X-905, Dec. 1963, p. 381.

74. Hubbard and Maglieri, “Factors Affecting Community Acceptance of the Sonic Boom,” ibid., 

pp. 399–412. 

75. McLean, Supersonic Transport Technology, p. 46.
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chaired by Robert McNamara, considered in evaluating the proposed 

SST designs. The sonic boom issue also became the focus of a special 

committee of the National Academy of Sciences and attracted increasing attention from the academic and scientific community at large.

The Acoustical Society of America, made up of professionals of all fields 

involving sound (ranging from music to noise to vibrations), sponsored the 

first Sonic Boom Symposium on November 3, 1965, during its 70th meeting 

in—appropriately enough—St. Louis. McLean, Hubbard, Carlson, Maglieri, 

and other Langley experts presented papers on the background of sonic 

boom research and their latest findings.76 The paper by McLean and Barrett 

L. Shrout included details on a breakthrough in using near-field shock waves 

to evaluate wind tunnel models for boom minimization, in this case a reduction in maximum overpressure in a climb profile from 2.2 to 1.1 psf. This 

technique also allowed the use of 4-inch models, which were easier to fabricate to the close tolerances required for accurate measurements.77

In addition to the scientists and engineers employed by the aircraft 

manufactures, eminent researchers in academia took on the challenge 

of discovering ways to minimize the sonic boom, usually with support 

from NASA. These included the team of Albert George and A. Richard 

Seebass of Cornell University, which had one of the Nation’s premier 

aeronautical laboratories. Seebass edited the proceedings of NASA’s first 

sonic boom research conference, held on April 12, 1967. The meeting 

was chaired by another pioneer of minimization, Wallace D. Hayes of 

Princeton University, and attended by more than 60 other Government, 

industry, and university experts. Boeing had been selected as the SST 

contractor less than 4 months earlier, but the sonic boom was becoming recognized far and wide as a possibly fatal flaw for its future production, or at least for allowing it to fly supersonically over land.78 The 

two most obvious theoretical ways to reduce sonic booms during supersonic cruise—flying much higher with no increase in weight or building 

76. JASA, vol. 39, no. 5, pt. 2 (Nov. 1966), pp. 519–572.

77. F. Edward McLean and Barrett L. Shrout, “Design Methods for Minimization of Sonic Boom 

Pressure-Field Disturbances,” ibid., 519–525. For an updated report, see Carlson, McLean, and 

Shrout, “A Wind Tunnel Study of Sonic-Boom Characteristics for Basic and Modified Models of a 

Supersonic Transport Configuration,” NASA TM-X-1236, May 1966.

78. Evert Clark, “Sonic Boom to Limit Speed of Superjets Across U.S.,” New York Times, Oct. 31, 

1966, pp. 1, 71; “George Gardner, “Overland Flights by SST Still in Doubt,” Washington Post, 

July 10, 1967, p. A7.
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an airframe 50 percent longer at half the weight—were not considered 

practical.79 Furthermore, as apparent from a presentation by Domenic 

Maglieri on flight test findings, such an airplane would still have to deal 

with the problem of booms caused by maneuvering and accelerating, 

and from atmospheric conditions.80

The stated purpose of this conference was “to determine whether 

or not all possible aerodynamic means of reducing sonic boom overpressure were being explored.”81 In that regard, Harry Carlson showed 

how various computer programs then being used at Langley for aerodynamic analyses (e.g., lift and drag) were also proving to be a useful 

tool for bow wave predictions, complementing improved wind tunnel 

experiments for examining boom minimization concepts.82 After presentations by representatives from NASA, Boeing, and Princeton, and 

follow-on discussions by other experts, some of the attendees thought 

more avenues of research could be explored. But many were still concerned whether low enough sonic booms were possible using contemporary technologies. Accordingly, NASA’s Office of Advanced Research 

and Technology, which hosted the conference, established specialized 

research programs on seven aspects of sonic boom theory and applications at five American universities and the Aeronautical Research 

Institute of Sweden.83 This mobilization of aeronautical brainpower 

almost immediately began to pay dividends.

Seebass and Hayes cochaired NASA’s second sonic boom conference 

on May 9–10, 1968. It included 19 papers on the latest boom-related testing, research, experimentation, and theory by specialists from NASA 

and the universities. The advances made in one year were impressive. 

In the area of theory, for example, the straightforward linear technique 

for predicting the propagation of sonic booms from slender airplanes 

such as the SST had proven reliable, even for calculating some nonlinear 

(mathematically complex and highly erratic) aspects of their signatures. 

79. A.R. Seebass, ed., Sonic Boom Research: Proceedings of a Conference . . . Washington, DC, 

Apr. 12, 1967, NASA SP-147, 1967.

80. Maglieri, “Sonic Boom Flight Research—Some Effects of Airplane Operations and the Atmosphere on Sonic Boom Signatures,” ibid., pp. 25–48.

81. A.R. Seebass, “Preface,” ibid., p. iii.

82. Carlson, “Experimental and Analytical Research on Sonic Boom Generation at NASA,” ibid., 

pp. 9–23. 

83. Ira R. Schwartz, ed., Sonic Boom Research, Second Conference, Washington, DC, May 

9–10, 1968, NASA SP-180, 1968, pp. iv–v.
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Additional field testing had improved understanding of the geometrical acoustics caused by atmospheric conditions. Computational capabilities needed to deal with such complexities continued to accelerate. 

Aeronautical Research Associates of Princeton (ARAP), under a NASA 

contract, had developed a computer program to calculate overpressure 

signatures for supersonic aircraft in a horizontally stratified atmosphere. 

Offering another preview of the digital future, researchers at Ames had 

begun using a computer with graphic displays to perform flow-field 

analyses and to experiment with a dozen diverse aircraft configurations 

for lower boom signatures. Several other papers by academic experts, 

such as Antonio Ferri of New York University (a notable prewar Italian 

aerodynamicist who had worked at the NACA’s Langley Laboratory after 

escaping to the United States in 1944), dealt with progress in the aerodynamic techniques to reduce sonic booms.84

Nevertheless, several important theoretical problems remained, such 

as the prediction of sonic boom signatures near a caustic (an objective of 

the previously described Jackass Flats testing in 1970), the diffraction of 

shock waves into “shadow zones” beyond the primary sonic boom carpet, nonlinear shock wave behavior near an aircraft, and the still mystifying effects of turbulence. Ira R. Schwartz of NASA’s Office of Advanced 

Research and Technology summed up the state of sonic boom minimization as follows: “It is yet too early to predict whether any of these 

design techniques will lead the way to development of a domestic SST 

that will be allowed to fly supersonic over land as well as over water.”85

Rather than conduct another meeting the following year, NASA 

deferred to a conference by NATO’s Advisory Group for Aerospace 

Research & Development (AGARD) on aircraft engine noise and sonic 

boom, held in Paris during May 1969. Experts from the United States 

and five other nations attended this forum, which consisted of seven sessions. Three of the sessions, plus a roundtable, dealt with the status of 

boom research and the challenges ahead.86 As reflected by these conferences, the three-way partnership between NASA, Boeing, and the academic aeronautical community during the late 1960s continued to yield 

new knowledge about sonic booms as well as technological advance in 

84. Ibid., 1–193; For more on the ARAP computer program, see Wallace D. Hayes, et al., “Sonic 

Boom Propagation in a Stratified Atmosphere with Computer Program,” NASA CR-1299, Apr. 1969. 

85. Second Conference on Sonic Boom Research, p. vii.

86. AGARD, Aircraft Engine Noise and Sonic Boom (see note 19 for bibliographical data).
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exploring ways to deal with them. In addition to more flight test data and 

improved theoretical constructs, much of this progress was the result 

of various experimental apparatuses. 

The use of wind tunnels (especially Langley’s 4 by 4 supersonic wind 

tunnels and the 9 by 7 and 8 by 7 supersonic sections of Ames’s Unitary 

Wind Tunnel complex) continued to advance the understanding of shock 

wave generation and aircraft configurations that could minimize the 

sonic boom.87 As two of Langley’s sonic boom experts reported in 1970, 

the many challenges caused by nonuniform tunnel conditions, model 

and probe vibrations, boundary layer effects, and the precision needed 

for small models “have been met with general success.”88

Also during the latter half of the 1960s, NASA and its contractors developed several new types of simulators that proved useful in 

studying the physical and psychoacoustic effects of sonic booms. The 

smallest (and least expensive) was a spark discharge system. The Langley 

Center and other laboratories used these “bench-type” devices for basic 

research into the physics of pressure waves. Langley’s system created 

miniature sonic booms by using parabolic or two-dimensional mirrors to focus the shock waves caused by discharging high voltage bolts 

of electricity between tungsten electrodes toward precisely placed 

microphones. Such experiments were used to verify laws of geometrical acoustics. The system’s ability to produce shock waves that spread 

out spherically proved useful for investigating how the cone-shaped 

waves generated by aircraft interact with buildings.89

For studying the effect of temperature gradients on boom propagation, Langley used a ballistic range consisting of a helium gas launcher 

that shot miniature projectiles at constant Mach numbers through a 

partially enclosed chamber. The inside could be heated to ensure a stable atmosphere for accuracy in boom measurements. Innovative NASAsponsored simulators included Ling-Temco-Vought’s shock-expansion 

tube, basically a mobile 13-foot-diameter conical horn mounted on a 

87. For a survey, see Daniel D. Baals and William R. Corliss, “Wind Tunnels of NASA,” SP-440 

(Washington, DC: NASA, 1981).

88. Phillip M. Edge and Harvey H. Hubbard, “Review of Sonic-Boom Simulation Devices and 

Techniques,” Dec. 1970, JASA Journal, vol. 51, No. 2, pt. 2 (Feb. 1972), p. 723.

89. W.D. Beasly, J.D. Brooks, and R.L. Barger, “A Laboratory Investigation of N-Wave Focusing,” 

NASA TN-D-5306, July 1969; J.D. Brooks, et al., “Laboratory Investigation of Diffraction and 

Reflection of Sonic Booms by Buildings,” NASA TN-D-5830, June 1970.
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trailer, and General American Research Division’s explosive gas-filled 

envelopes suspended above sensors at Langley’s sonic boom simulation 

range.90 NASA also contracted with Stanford Research Institute for simulator experiments that showed how sonic booms could interfere with 

sleep, especially for older people.91

Other simulators were devised to handle both human and structural response to sonic booms. (The need to better understand effects on 

people was called for in a report released in June 1968 by the National 

Academy of Sciences.)92 Unlike the previously described studies using 

actual sonic booms created by aircraft, these devices had the advantages of a controlled laboratory environment. They allowed researchers to produce multiple boom signatures of varying shapes, pressures, 

and durations as often as needed at a relatively low cost.93 The Langley 

Center’s Low-Frequency Noise Facility—built earlier in the 1960s to generate the intense chest-pounding sounds of giant Saturn boosters during 

Apollo launches—also performed informative sonic boom simulation 

experiments. Consisting of a cylindrical test chamber 24 feet in diameter and 21 feet long, it could accommodate people, small structures, 

and materials for testing. Its electrohydraulically operated 14-foot piston was capable of producing sound waves from 1–50 hertz (sort of a 

super subwoofer) and sonic boom N-waves from 0.5 to 20 psf at durations from 100 to 500 milliseconds.94

To provide an even more versatile system designed specifically for 

sonic boom research, NASA contracted with General Applied Science 

Laboratories (GASL) of Long Island, NY, to develop an ideal simulator 

using a quick action valve and shock tube design. (Antonio Ferri was 

90. Edge and Hubbard, “Review of Sonic Boom Simulation,” pp. 724–728; Hugo E. Dahlke, et 

al., “The Shock-Expansion Tube and Its Application as a Sonic Boom Simulator,” NASA CR-1055, 

June 1968; R.T Sturgielski, et al., “The Development of a Sonic Boom Simulator with Detonable 

Gases,” NASA CR-1844, Nov. 1971.

91. Jerome Lukas and Karl D. Kryler, “A Preliminary Study of the Awakening and Startle Effects of 

Simulated Sonic Booms,” NASA CR-1193, Sept. 1968, “Awakening Effects of Simulated Sonic 

Booms and Subsonic Aircraft Noise . . . ,” NASA CR-1599, May 1970.

92. David Hoffman, “Report Sees Need for Study on Sonic Boom Tolerance,” Washington Post, 

June 26, 1968, p. A3.

93. Ira R. Schwartz, Sonic Boom Simulation Facilities,” AGARD, Aircraft Engine Noise and Sonic 

Boom, p. 29-1.

94. Philip M. Edge and William H. Mayes, “Description of Langley Low-Frequency Noise Facility and 

Study of Human Response to Noise Frequencies below 50 cps,” NASA TN-D-3204, Jan. 1966.
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the president of GASL, which he had cofounded with the illustrious 

aeronautical scientist Theodore von Kármán in 1956). Completed in 

1969, this new simulator consisted of a high-speed flow valve that sent 

pressure wave bursts through a heavily reinforced 100-foot-long conical duct that expanded into an 8 by 8 test section with an instrumentation and model room. It could generate overpressures up to 10 psf 

with durations from 50 to 500 milliseconds. Able to operate at less than 

a 1-minute interval between bursts, its sonic boom signatures proved 

very accurate and easy to control.95 In the opinion of Ira Schwartz, “the 

GASL/NASA facility represents the most advanced state of the art in 

sonic boom simulation.”96

While NASA and its partners were learning more about the nature 

of sonic booms, the SST was becoming mired in controversy. Many in 

the public, the press, and the political arena were concerned about the 

noise SSTs would create, with a growing number expressing hostility 

to the entire SST program. As one of the more reputable critics wrote 

in 1966, with a map showing a dense network of future boom carpets 

crossing the United States, “the introduction of supersonic flight, as it 

is at present conceived, would mean that hundreds of millions of people would not only be seriously disturbed by the sonic booms . . . they 

would also have to pay out of their own pockets (through subsidies) to 

keep the noise-creating activity alive.”97

Opposition to the SST grew rapidly in the late 1960s, becoming a 

cause celebre for the burgeoning environmental movement as well as 

target for small-Government conservatives opposed to Federal subsidies.98 Typical of the growing trend among opinion makers, the New 

York Times published its first strongly anti–sonic-boom editorial in June 

1968, linking the SST’s potential sounds with an embarrassing incident 

the week before when an F-105 flyover shattered 200 windows at the Air 

Force Academy, injuring a dozen people.99 The next 2 years brought a 

95. Roger Tomboulian, Research and Development of a Sonic Boom Simulation Device, NASA CR1378, July 1969; Stacy V. Jones, “Sonic Boom Researchers Use Simulator,” New York Times, May 

10, 1969, pp. 37, 41.

96. Ira R. Schwartz, “Sonic Boom Simulation Facilities,” p. 29-6.

97. B.K. O. Lundberg, “Aviation Safety and the SST,” Astronautics & Aeronautics, vol. 3, No. 1 

(Jan. 1966), p. 28. Lundberg was a Swedish scientist very critical of SSTs.

98. See Conway, High-Speed Dreams, pp. 118–156.

99. “The Shattering Boom,” New York Times, June 8, 1968, p. 30.
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growing crescendo of complaints about the supersonic transport, both 

for its expense and the problems it could cause—even as research on 

controlling sonic booms began to bear some fruit. 

By the time 150 scientists and engineers gathered in Washington, 

DC, for NASA’s third sonic boom research conference on October 29–30, 

1970, the American supersonic transport program was less than 6 months 

away from cancellation. Thus the 29 papers presented at the conference 

and others at the ASA’s second sonic boom symposium in Houston the 

following month might be considered, in their entirety, as a final status 

report on sonic boom research during the SST decade.100 Of future if not 

near-term significance, considerable progress was being made in understanding how to design airplanes that could fly faster than the speed of 

sound while leaving behind a gentler sonic footprint.

As summarized by Ira Schwartz: “In the area of boom minimization, the NASA program has utilized the combined talents of Messrs. E. 

McLean, H.L. Runyan, and H.R. Henderson at NASA Langley Research 

Center, Dr. W.D. Hayes at Princeton University, Drs. R. Seebass and A.R. 

George at Cornell University, and Dr. A. Ferri at New York University 

to determine the optimum equivalent bodies of rotation [a technique 

for relating airframe shapes to standard aerodynamic rules governing 

simple projectiles with round cross sections] that minimize the overpressure, shock pressure rise, and impulse for given aircraft weight, 

length, Mach number, and altitude of operation. Simultaneously, 

research efforts of NASA and those of Dr. A. Ferri at New York University 

have provided indications of how real aircraft can be designed to provide 

values approaching these optimums. . . . This research must be continued or even expanded if practical supersonic transports with minimum 

and acceptable sonic boom characteristics are to be built.”101

Any consensus among the attendees about the progress they were 

making was no doubt tempered by their awareness of the financial 

problems now plaguing the Boeing Company and the political difficulties facing the administration of President Richard Nixon in continuing to subsidize the American SST. From a technological standpoint, 

many of them also seemed resigned that Boeing’s final 2707-300 

100. Ira R. Schwartz, ed., Third Conference on Sonic Boom Research . . . Washington, DC, 

Oct. 29–30, 1970, NASA SP-255, 1971. The papers from the ASA’s Houston symposium were 

published in JASA, vol. 51, No. 2 (Feb. 1972), pt. 2.

101. Third Conference on Sonic Boom Research, Preface by Ira Schwartz, p. iv.
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design (despite its 306-foot length and 64,000-foot cruising altitude) 

would not pass the overland sonic boom test. Richard Seebass, who 

was in the vanguard of minimization research, admitted that “the first 

few generations of supersonic transport (SST) aircraft, if they are built 

at all, will be limited to supersonic flight over oceanic and polar regions.”102

In view of such concerns, some of the attendees were even looking 

toward hypersonic aerospace vehicles, in case they might cruise high 

enough to leave an acceptable boom carpet. 

As for the more immediate prospects of a domestic supersonic transport, Lynn Hunton of the Ames Research Center warned that “with 

regard to experimental problems in sonic boom research, it is essential 

that the techniques and assumptions used be continuously questioned 

as a requisite for assuring the maximum in reliability.”103 Harry Carlson 

probably expressed the general opinion of Langley’s aerodynamicists 

when he cautioned that “the problem of sonic boom minimization 

through airplane shaping is inseparable from the problems of 

optimization of aerodynamic efficiency, propulsion efficiency, and structural weight. . . . In fact, if great care is not taken in the application of 

sonic boom design principles, the whole purpose can be defeated by performance degradation, weight penalties, and a myriad of other practical considerations.”104

After both the House and Senate voted in March 1971 to eliminate SST funding, a joint conference committee confirmed its termination in May.105 This and related cuts in supersonic research inevitably 

slowed momentum in dealing with sonic booms. Even so, researchers in NASA, as well as in academia and the aerospace industry, would 

keep alive the possibility of civilian supersonic flight in a more constrained 

and less technologically ambitious era. Fortunately for them, the illfated SST program left behind a wealth of data and discoveries about sonic 

booms. As evidence, the Langley Research Center produced or sponsored 

more than 200 technical publications on the subject over 19 years, 

most related to the SST program. (Many of those published in the 

early 1970s were based on previous research and testing.) This 

102. R. Seebass, “Comments on Sonic Boom Research,” ibid., p. 411.

103. Lynn W. Hunton, “Comments on Low Sonic Boom Configuration Research, ibid., p. 417.

104. Carlson, “Sonic Boom Prediction and Minimization Research,” ibid., p. 397.

105. For a detailed postmortem, see Edward Wenk, “SST—Implications of a Political Decision,” 

Astronautics & Aeronautics, vol. 9, No. 10 (Oct. 1971), pp. 40–49.
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literature, depicted in Figure 4, would be a legacy of enduring value in 

the future.106

Keeping Hopes Alive: Supersonic Cruise Research

“The number one technological tragedy of our time.” That was how 

President Nixon characterized the votes by the Congress to stop funding an American supersonic transport.107 Despite its cancellation, the 

White House, the Department of Transportation (DOT), and NASA—as 

well as some in Congress—did not allow the progress in supersonic technologies the SST had engendered to completely dissipate. During 1971 

and 1972, the DOT and NASA allocated funds for completing some of 

the tests and experiments that were underway when the program was 

terminated. The administration then added line-item funding to NASA’s 

fiscal year (FY) 1973 budget for scaled-down supersonic research, especially as related to environmental problems. In response, NASA established the Advanced Supersonic Technology (AST) program in July 1972. 

To more clearly indicate the exploratory nature of this effort and 

allay fears that it might be a potential follow-on to the SST, the AST 

program was renamed Supersonic Cruise Aircraft Research (SCAR) in 

1974. When the term aircraft in its title continued to raise suspicion 

in some quarters that the goal might be some sort of prototype, NASA 

shortened the program’s name to Supersonic Cruise Research (SCR) in 

1979.108 For the sake of simplicity, the latter name is often applied to all 

9 years of the program’s existence. For NASA, the principal purpose of 

AST, SCAR, and SCR was to conduct and support focused research into 

the problems of supersonic flight while advancing related technologies. 

NASA’s aeronautical Centers, most of the major airframe manufactures, 

and many research organizations and universities participated. From 

Washington, NASA’s Office of Aeronautics and Space Technology (OAST) 

106. Compiled by screening B.A. Fryer, et al., “Publications in Acoustics and Noise Control from the 

NASA Langley Research Center During 1940–1976,” NASA TM-X-7402, July 1977. Five reports for 

1967 that Maglieri (in reviewing the draft of this chapter) found missing from Fryer’s compilation have 

been added to that column.

107. Stephen D. Ambrose, Nixon: Triumph of a Politician, vol. 2 (New York: Simon and Schuster, 

1989), p. 433, cited by Conway, High-Speed Dreams, p. 153. For the political and bureaucratic 

background of the AST program, see Conway, pp. 153–158.

108. F. Edward McLean, “SCAR Program Overview,” Proceedings of the SCAR Conference . . . 

Hampton, VA, Nov. 9–12, 1976, pt. 1, NASA CP-001, 1976, pp. 1–3; McLean, Supersonic 

Cruise Technology, pp. 101–102. 
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provided overall supervision but delegated day-to-day management to the 

Langley Research Center, which established an AST Project Office in its 

Directorate of Aeronautics, soon placed under a new Aeronautics System 

Division. The AST program was organized into four major elements—

propulsion, structure and materials, stability and control, and aerodynamic performance—plus airframe-propulsion integration. (NASA spun 

off the propulsion work on a variable cycle engine [VCE] as a separate 

program in 1976.) Sonic boom research was one of 16 subelements.109

At the Aeronautical Systems Division, Cornelius “Neil” Driver, who 

headed the Vehicle Integration Branch, and Ed McLean, as chief of the 

AST Project Office, were key officials in planning and managing the 

AST/SCAR effort. After McLean retired in 1978, the AST Project Office 

passed to a fellow aerodynamicist, Vincent R. Mascitti, while Driver 

took over the Aeronautical Systems Division. One year later, Domenic 

Maglieri replaced Mascitti in the AST Project Office.110 Despite Maglieri’s 

sonic boom expertise, the goal of minimizing the AST’s sonic boom for 

overland cruise had long since ceased being an SCR objective. As later 

explained by McLean: “The basic approach of the SCR program . . . was 

to search for the solution of supersonic problems through disciplinary research. Most of these problems were well known, but no satisfactory solution had been found. When the new SCR research suggested 

a potential solution . . . the applicability of the suggested solution was 

assessed by determining if it could be integrated into a practical commercial supersonic airplane and mission. . . . If the potential solution 

could not be integrated, it was discarded.”111

To meet the practicality standard for integration into a supersonic 

airplane, solving the sonic boom problem had to clear a new and almost 

insurmountable hurdle. In April 1973, responding to years of political pressure, the FAA announced a new rule that banned commercial or civil aircraft from supersonic flight over the land mass or territorial waters of the 

United States if measurable overpressure would reach the surface.112 One 

of the initial objectives of the AST’s sonic boom research had been to 

109. McLean, Supersonic Cruise Technology, pp. 104–108; Sherwood Hoffman, “Bibliography of 

Supersonic Cruise Aircraft Research (SCAR),” NASA RP-1003, Nov. 1977, pp. 1–5.

110. Chambers, Innovation in Flight, pp. 39–40.

111. McLean, Supersonic Cruise Technology, p. 103.

112. FAA Chronology, Apr. 27, 1973. The rule was included as Federal Aviation Regulation (FAR) 

Section 91.817, Civil Aircraft Sonic Boom, effective Sept. 30, 1963. 
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Figure 4. Reports produced or sponsored by NASA Langley, 1958–1976. NASA.

establish a metric on public acceptability of sonic boom signatures for 

use in the aerodynamic design process. The FAA’s stringent new regulation seemed to rule out any such flexibility. 

As a result, when Congress cut FY 1974 funding for the AST program from $40 million to about $10 million, the subelement for sonic 

boom research went on NASA’s chopping block. The design criteria for 

the SCAR/SCR program became a 300-foot-long, 270-passenger airplane 

that could fly as effectively as possible over land at subsonic speeds yet 

still cruise efficiently at 60,000 feet and Mach 2.2 over water. To meet 

these criteria, Langley aerodynamicists modified their SCAT-15F design 

from the late 1960s into a notional concept with better low-speed performance (but higher sonic boom potential) called the ATF-100. This 

served as a baseline for three industry teams in coming up with their 

own designs.113

When the AST program began, however, prospects for a significant 

quieting of its sonic footprint appeared possible. Sonic boom theory had 

advanced significantly during the 1960s, and some promising if not yet 

practical ideas for reducing boom signatures had begun to emerge. As 

113. Marvin Miles, “Hopes for SST Are Dim but R&D Continues—Just in Case,” Los Angeles Times, 

Nov. 25, 1973, pp. G1, 11; McLean, Supersonic Cruise Technology, pp. 117-118; Conway, 

High-Speed Dreams, pp. 176–180.
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indicated by Figure 4, some findings based on that research continued 

to come out in print during the early 1970s. 

As far back as 1965, NASA’s Ed McLean had discovered that the sonic 

boom signature from a very long supersonic aircraft flying at the proper 

altitude could be nonasymptotic (i.e., not reach the ground in the form 

of an N-wave). This confirmed the possibility of tailoring an airplane’s 

shape into something more acceptable.114 Some subsequent theoretical 

suggestions, such as various ways of projecting heat fields to create a 

longer “phantom” fuselage, are still decidedly futuristic, while others, 

such as adding a long spike to the nose of an SST to slow the rise of the 

bow shock wave, would (as described later) eventually prove more realistic.115 Meanwhile, researchers under contract to NASA kept advancing 

the state of the art in more conventional directions. For example, Antonio 

Ferri of New York University in partnership with Hans Sorensen of the 

Aeronautical Research Institute of Sweden used new 3-D measuring 

techniques in Sweden’s trisonic wind tunnel to more accurately correlate 

near-field effects with linear theory. Testing NYU’s model of a 300-footlong SST cruising at Mach 2.7 at 60,000 feet, it showed the opportunity 

for sonic booms of less than 1.0 psf.116 Ferri’s early death in 1975 left a 

void in supersonic aerodynamics, not least in sonic boom research.117

By the end of the SST program, Albert George and Richard Seebass 

had formulated a mathematical foundation for many of the previous 

theories. They also devised a near-field boom-minimization theory, applicable in an isothermal atmosphere, for reducing the overpressures of 

flattop and ramp-type signatures. It was applicable to both front and 

rear shock waves along with their parametric correlation to airframe lift 

and area. In a number of seminal papers and articles in the early 1970s, 

they explained the theory along with some ideas on possible aerodynamic 

114. F. Edward McLean, “Some Non-asymptotic Effects on the Sonic Boom of Large Airplanes,” 

NASA TN-D-2877, June 1965, as interpreted by Plotkin and Maglieri, “Sonic Boom Research,” p. 5.

115. Miles, “Sonic Boom Not Insoluble, Scientist Says,” Los Angeles Times, Dec. 10, 1970, pp. E4–

E5; David S. Miller and Carlson, “Sonic Boom Minimization by Application of Heat or Force Fields to 

Airplane Airflow,” NASA TN-D-5582, Dec. 1969; Rudolph J. Swigart, “An Experimental Study in the 

Validity of the Heat-Field Concept for Sonic Boom Alleviation,” NASA CR-2381, Mar. 1974. 

116. Antonio Ferri, Huai-Chu Wang, and Hans Sorensen, “Experimental Verification of Low Sonic 

Boom Configuration,” NASA CR-2070, June 1973.

117. For a retrospective, see Percy J. Bobbit and Maglieri, “Dr. Antonio Ferri’s Contribution to 

Supersonic Transport Sonic-Boom Technology,” Journal of Spacecraft and Rockets, vol. 40, no. 4 

(July–Aug. 2003), pp.459–466.
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shaping (e.g., slightly blunting an aircraft’s nose) and the optimum cruise 

altitude (lower than previously thought) for reducing boom signatures.118

Theoretical refinements and new computer modeling techniques continued to appear in the early 1970s. For example, in June 1972, Charles 

Thomas of the Ames Research Center explained a mathematical procedure using new algorithms for waveform parameters to extrapolate the 

formation of far-field N-waves. This was an alternative to using F-function 

effects (the pattern of near-field shock waves emanating from an airframe), which were the basis of the previously discussed program developed by Wallace Hayes and colleagues at ARAP. Although both methods 

accounted for acoustical ray tracing and could arrive at similar results, 

Thomas’s program allowed easier input of flight information (speed, 

altitude, atmospheric conditions, etc.) for automated data processing.119

In June 1973, at the end of the AST program’s first year, NASA Langley’s 

Harry Carlson, Raymond Barger, and Robert Mack published a study on 

the applicability of sonic boom minimization concepts for overland supersonic transport designs. They examined four reduced-boom concepts for 

a commercially viable Mach 2.7 SST with a range of 2,500 nautical miles 

(i.e., coast to coast in the United States). Using experimentally verified 

minimization concepts of George, Seebass, Hayes, Ferri, Barger, and the 

English researcher L.B. Jones, along with computational techniques developed at Langley during the SST program, Carlson’s team examined ways to 

manipulate the F-function to project a flatter far-field sonic boom signature. 

In doing this, the team was handicapped by the continuing lack of established signature characteristics (the combinations of initial peak overpressure, maximum shock strength, rise time, and duration) that people would 

best tolerate, both outdoors and especially indoors. Also, the complexity 

of aft aircraft geometry made measuring effects on tail shocks difficult.120

Even so, their study confirmed the advantages of designs with highly 

swept wings toward the rear of the fuselage with twist and camber for 

118. A.R. George and R. Seebass, “Sonic-Boom Minimization,” Nov. 1970, in JASA, vol. 51, no. 2, 

pt. 3 (Feb. 1972), pp. 686–694; A.R. Seebass and A.R George, “Sonic Boom Minimization through 

Aircraft Design and Operation,” AIAA Paper 73-241, Jan. 1973; A.R. Seebass and A.R. George, 

“The Design and Operation of Aircraft to Minimize Their Sonic Boom,” Journal of Aircraft, vol. 11, no. 

9 (Sept. 1974), pp. 509–517. (Quote is from p. 516.)

119. Charles L. Thomas, “Extrapolation of Sonic Boom Pressure Signatures by the Waveform 

Parameter Method,” NASA TN-D-6823, June 1972.

120. Carlson, Raymond L. Barger, and Robert J. Mack, “Application of Sonic-Boom Minimization 

Concepts in Supersonic Transport Design,” NASA TN-D-7218, June 1973.
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sonic boom shaping. It also found the use of canards (small airfoils used as 

horizontal stabilizers near the nose of rear-winged aircraft) could optimize 

lift distribution for sonic boom benefits. Although two designs showed 

bow shocks of less than 1.0 psf, their report noted “that there can be no 

assurance at this time that [their] shock-strength values . . . if attainable, 

would permit unrestricted overland operations of supersonic transports.”121

Ironically, these words were written just before the new FAA rule rendered 

them largely irrelevant.

In October 1973, Edward J. Kane of Boeing, who had been a key 

sonic boom expert during the SST program, released the results of a 

similar NASA-sponsored study on the feasibility of a commercially viable low-boom transport using technologies projected to be available in 

1985. Based on the latest theories, Boeing explored two longer-range concepts: a high-speed (Mach 2.7) design that would produce a sonic boom 

of 1.0 psf or less, and a medium-speed (Mach 1.5) design with a signature 

of 0.5 psf or less.122 In retrospect, this study, which reported mixed results, 

represented industry’s perspective on the prospects for boom minimization just as the AST program dropped plans for supersonic cruise over land.

Obviously, the virtual ban on civilian supersonic flight in the United 

States dampened any enthusiasm by private industry to continue investing very much capital in sonic boom research. Within NASA, some 

of those with experience in sonic boom research also redirected their 

efforts into other areas of expertise. Of the approximately 1,000 technical reports, conference papers, and articles by NASA and its contractors 

listed in bibliographies of the SCR program from 1972 to 1980, only 8 

dealt directly with the sonic boom.123

Even so, progress in understanding sonic booms did not come to a 

complete standstill. In 1972, Christine M. Darden, a Langley mathematician in an engineering position, had developed a computer code to adapt 

Seebass and George’s minimization theory, which was based on an isothermal (uniform) atmosphere, into a program that applied to a standard (stratified) atmosphere. It also allowed more design flexibility than 

121. Ibid., p. 28.

122. Edward J. Kane, “A Study to Determine the Feasibility of a Low Sonic Boom Supersonic Transport,” AIAA Paper 73-1035, Oct. 1973. See also NASA CR-2332, Dec. 1973.

123. Sherwood Hoffman, “Bibliography of Supersonic Cruise Aircraft Research (SCAR)” [1972–

1977], NASA RP-1003, Nov. 1977, and “Bibliography of Supersonic Cruise Research (SCR) 

Program from 1977 to Mid-1980,” NASA RP-1063, Dec. 1980.
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previous low-boom configuration theory did, such as better aerodynamics in the nose area.124

Using this new computer program, Darden and Robert Mack followed up on the previously described study by Carlson’s team by designing wing-body models with low-boom characteristics: one for cruise at 

Mach 1.5 and two for cruise at Mach 2.7. At 6 inches in length, these 

were the largest yet tested for sonic boom propagation in a4 by 4 supersonic wind tunnel—an improvement made possible by continued progress in measuring and extrapolating near-field effects to signatures in the 

far field. The specially shaped models (all arrow-wing configurations, 

which distributed lift effects to the rear) showed significantly lower overpressures and flatter signatures than standard designs did, especially at 

Mach 1.5, at which both the bow and tail shocks were softened. Because 

of funding limitations, this promising research could not be sustained 

long enough to develop definitive boom minimization techniques.125 It 

was apparently the last significant experimentation on sonic boom minimization for more than a decade.

While this work was underway, Darden and Mack presented a paper 

on current sonic boom research at the first SCAR conference, held at 

Langley on November 9–12, 1976 (the only paper on that subject among the 

47 presentations). “Contrary to earlier beliefs,” they explained, “it has been 

found that improved efficiency and lower sonic boom characteristics do 

not always go hand in hand.” As for the acceptability of sonic booms, they 

reported that the only research in North America was being done at the 

University of Toronto.126 Another NASA contribution to understanding 

sonic booms came in early 1978 with the publication by Harry Carlson 

124. Christine M. Darden, “Minimization of Sonic-Boom Parameters in Real and Isothermal Atmospheres,” 

NASA TN D-7842, Mar. 1975; Darden, “Sonic Boom Minimization with Nose-Bluntness Relaxation,” 

NASA TP-1348, Jan. 1979; Darden, “Affordable/Acceptable Supersonic Flight: Is It Near?” 40th Aircraft 

Symposium, Japan Society for Aeronautical and Space Sciences (JSASS), Yokohama, Oct. 9–11, 2002.

125. Robert J. Mack and Darden, “Wind-Tunnel Investigation of the Validity of a Sonic-BoomMinimization Concept,” NASA TP-1421, Oct. 1979. They had previously presented their findings 

at an AIAA conference in Seattle on Mar. 12–14, 1979 as “Some Effects of Applying Sonic Boom 

Minimization to Supersonic Cruise Aircraft Design,” AIAA Paper 79-0652, also published in Journal 

of Aircraft, vol. 17, no. 3 (Mar. 1980), pp. 182–186.

126. Darden and Mack, “Current Research in Sonic-Boom Minimization,” Proceedings of the SCAR 

Conference [1976], pt. 1, pp. 525–541 (quote from p. 526). Darden had discussed some of these 

topics in “Sonic Boom Theory – Its Status in Prediction and Minimization,” AIAA Paper 76-1, presented 

at the AIAA Aerospace Sciences Meeting, Washington, DC, Jan. 26–28, 1976. 
NASA’s Contributions to Aeronautics

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of “Simplified Sonic-Boom Prediction,” a how-to guide on a relatively 

quick and easy method to determine sonic boom characteristics. It could 

be applied to a wide variety of supersonic aircraft configurations as well 

as spacecraft at altitudes up to 76 kilometers. Although his clever series of 

graphs and equations would not provide the accuracy needed to predict 

booms from maneuvering aircraft or in designing airframe configurations, Carlson explained that “for many purposes (including the conduct 

of preliminary engineering studies or environmental impact statements), 

sonic-boom predictions of sufficient accuracy can be obtained by using 

a simplified method that does not require a wind tunnel or elaborate 

computing equipment. Computational requirements can in fact be met 

by hand-held scientific calculators, or even slide rules.”127

The month after publication of this study, NASA released its final 

environmental impact statement (EIS) for the Space Shuttle program, 

which benefited greatly from the Agency’s previous research on sonic 

booms, including that with the X-15 and Apollo missions, and adaptations of Charles Thomas’s waveform-based computer program.128 While 

ascending, the EIS estimated maximum overpressures of 6 psf (possibly up to 30 psf with focusing effects) about 40 miles downrange over 

open water, caused by both its long exhaust plume and its curving flight 

profile while accelerating toward orbit. During reentry of the manned 

vehicle, the sonic boom was estimated at a more modest 2.1 psf, which 

would affect about 500,000 people as it crossed the Florida peninsula 

or 50,000 when landing at Edwards.129 In following decades, as populations in those areas boomed, millions more would be hearing the 

sonic signatures of returning Shuttles, more than 120 of which would be 

monitored for their sonic booms.130

Some other limited experimental and theoretical work on sonic 

booms continued in the late 1970s. Richard Seebass at Cornell and 

Kenneth Plotkin of Wyle Research, for example, delved deeper into the 

127. Carlson, “Simplified Sonic-Boom Prediction,” NASA TP-1122, Mar. 1978, p. 1.

128. Paul Holloway of Langley and colleagues from the Ames, Marshall, and Johnson centers 

presented an early analysis, “Shuttle Sonic Boom—Technology and Predictions,” in AIAA Paper 

73-1039, Oct. 1973.

129. NASA HQ (Myron S. Malkin), Environmental Impact Statement: Space Shuttle Program (Final), 

Apr. 1978, pp. 106–116.

130. Including measurements in Hawaii, with the Shuttle at 253,000 feet and moving at Mach 

23. Telephone interview, Maglieri by Benson, Mar. 18, 2009.
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challenging phenomena of caustics and focused booms.131 At the end of 

the decade, Langley’s Raymond Barger published a study on the 

relationship of caustics to the shape and curvature of acoustical wave 

fronts caused by actual aircraft maneuvers. To graphically display these 

effects, he programmed a computer to draw simulated three-dimensional 

line plots of the acoustical rays in the wave fronts. Figure 5 shows how 

even a simple decelerating turn, in this case from Mach 2.4 to Mach 1.5 

in a radius of 23 kilometers (14.3 miles), can focus the kind of caustic 

that might result in a super boom.132

Unlike in the 1960s, there was little if any NASA sonic boom flight 

testing during the 1970s. As a case in point, NASA’s YF-12 Blackbirds at 

Edwards (where the Flight Research Center was renamed the Dryden 

Flight Research Center in 1976) flew numerous supersonic missions in 

support of the AST/SCAR/SCR program, but none of them were dedicated 

to sonic boom issues.133 On the other hand, operations of the Concorde 

began providing a good deal of empirical data on sonic booms. 

One discovery about secondary booms came after British Airways 

and Air France began Concorde service to the United Sates in May 1976. 

Although the Concordes slowed to subsonic speeds while well offshore, 

residents along the Atlantic seaboard began hearing what were called 

the “East Coast Mystery Booms.” These were detected all the way from 

Nova Scotia to South Carolina, some measurable on seismographs.134

Although a significant number of the sounds defied explanation, studies 

by the Naval Research Laboratory, the Federation of American Scientists, 

a committee of the Jason DOD scientific advisory group, and the FAA 

eventually determined that most of the low rumbles heard in Nova Scotia 

and New England were secondary booms from the Concorde. They 

were reaching land after being bent or reflected by temperature varia131. Plotkin and Maglieri, “Sonic Boom Research,” pp. 5–6, 10.

132. Raymond L. Barger, “Sonic-Boom Wave-Front Shapes and Curvatures Associated with Maneuvering Flight,” NASA TP-1611, Dec. 1979. Fig. 5 is from p. 23.

133. James and Associates, ed., YF-12 Experiments Symposium: A conference held at Dryden Flight 

Research Center . . . Sept. 13–15, 1978, NASA CP-2054, 1978; Hallion and Gorn, On the Frontier, Appendix P [YF-12 Flight Chronology, 1969–1978], pp. 423–429. The Dryden Center tested 

an oblique wing aircraft, the AD-1, from 1979 to 1982. Although this configuration might have sonic 

boom benefits at mid-Mach speeds, it was not a consideration in this experimental program.

134. “Second Concorde Noise Report for Dulles Shows Consistency,” Aviation Week, July 19, 

1976, p. 235; William Claiborne, “Those Mystery Booms Defy Expert Explanation,” Washington 

Post, Dec. 24, 1977, p. A1.
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Figure 5. Acoustic wave front above a maneuvering aircraft. NASA.

tions high up in the thermosphere from Concordes still about 75 to 150 

miles offshore. In July 1978, the FAA issued new rules prohibiting the 

Concorde from creating sonic booms that could be heard in the United 

States. The new FAA rules did not address the issue of secondary booms 

because of their low intensity; nevertheless, after Concorde secondary 

booms were heard by coastal communities, the Agency became even 

more sensitive to the sonic boom potential inherent in AST designs.135

The second conference on Supersonic Cruise Research, held at NASA 

Langley in November 1979, was the first and last under its new name. 

More than 140 people from NASA, other Government agencies, and the 

aerospace industry attended. This time there were no presentations on 

the sonic boom, but a representative from North American Rockwell did 

describe the concept of a Mach 2.7 business jet for 8–10 passengers that 

would generate a sonic boom of only 0.5 psf.136 It would take another 

135. Deborah Shapely, “East Coast Mystery Booms: A Scientific Suspense Tale,” Science, vol. 199, 

no. 4336 (Mar. 31, 1978), pp. 1416–1417; “Concordes Exempted from Noise Rules,” Aviation 

Week, July 3, 1978, p. 33; G.J. MacDonald, et al., “Jason 1978 Sonic Boom Report,” JSR-78-09 (Arlington, VA: SRI International, Nov. 1978); Richard Kerr, “East Coast Mystery Booms: Mystery Gone but 

Booms Linger On,” Science, vol. 203, no. 4337 (Jan. 19, 1979), p. 256; John H. Gardner and Peter 

H. Rogers, “Thermospheric Propagation of Sonic Booms from the Concorde Supersonic Transport,” 

Naval Research Laboratory Memo Report 3904, Feb. 14, 1979 (DTIC AD A067201).

136. Robert Kelly, “Supersonic Cruise Vehicle Research/Business Jet,” Supersonic Cruise Research ’79: 

Proceedings of a Conference . . . Hampton, VA, Nov. 13–16, 1979, NASA CP-2108, pt. 2, pp. 

935–944.
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20 years for ideas about low-boom supersonic business jets to result in 

more than just paper studies.

Despite SCR’s relatively modest cost versus its significant technological accomplishments, the program suffered a premature death in 

1981. Reasons for this included the Concorde’s economic woes, opposition to civilian R&D spending by key officials in the new administration 

of President Ronald Reagan, and a growing federal deficit. These factors, combined with cost overruns for the Space Shuttle, forced NASA 

to abruptly cancel Supersonic Cruise Research without even funding 

completion of many final reports.137 As regards sonic boom research, an 

exception to this was a compilation of charts for estimating minimum 

sonic boom levels published by Christine Darden in May 1981. She and 

Robert Mack also published results of their previous experimentation that 

would be influential when efforts to soften the sonic boom resumed.138

SST Reincarnated: Birth of the High-Speed Civil Transport 

For much of the next decade, the most active sonic boom research took 

place as part of the Air Force’s Noise and Sonic Boom Impact Technology 

(NSBIT) program. This was a comprehensive effort started in 1981 to 

study the noises resulting from military training and operations, especially those involving environmental impact statements and similar assessments. Although NASA was not intimately involved with NSBIT, Domenic 

Maglieri (just before his retirement from the Langley Center) and the 

recently retired Harvey Hubbard compiled a comprehensive annotated bibliography of sonic boom research, organized into 10 major areas, to help 

inform NSBIT participants of the most relevant sources of information.139

One of the noteworthy achievements of the NSBIT program was to 

continue building a detailed sonic boom database (known as Boomfile) 

on all U.S. supersonic aircraft by flying them over a large array of newly 

developed sensors at Edwards AFB in the summer of 1987. Called the 

Boom Event Analyzer Recorder (BEAR), these unmanned devices 

137. Conway, High-Speed Dreams, pp. 180–188; Chambers, Innovations in Flight, p. 48.

138. Darden, “Charts for Determining Potential Minimum Sonic-Boom Overpressures for Supersonic Cruise 

Aircraft,” NASA TP-1820, May 1981; Darden and R.J. Mack, “Some Effects of Applying Sonic Boom 

Minimization to Supersonic Aircraft,” Journal of Aircraft, vol. 17, no. 3 (Mar. 1980), pp. 182–186.

139. Hubbard, Maglieri, and David G. Stephens, “Sonic-Boom Research—Selected Bibliography 

with Annotation,” NASA TM-87685, Sept. 1986.
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recorded the full sonic boom waveform in digital format.140 Other contributions of NSBIT were long-term sonic boom monitoring of combat 

training areas, continued assessment of structures exposed to sonic booms, 

studies of the effects of sonic booms on livestock and wildlife, and intensified research on focused booms (long an issue with maneuvering fighter 

aircraft). The latter included a specialized computer program (derived from 

that originated by NASA’s Thomas) called PCBoom to predict these events.141

In a separate project, fighter pilots were successfully trained to lay down super 

booms at specified locations (an idea first broached in the early 1950s).142

By the mid-1980s, the growing economic importance of nations in Asia 

was drawing attention to the long flight times required to cross the Pacific 

Ocean or the ability to reach most of Asia from Europe. The White House 

Office of Science and Technology (OST), reversing the administration’s 

initial opposition to civilian aeronautical research, took various steps to 

gain support for such activities. In March 1985, the OST released a report, 

“National Aeronautical R&D Goals: Technology for America’s Future,” 

which included a long-range supersonic transport.143 Then, in his State of 

the Union Address in January 1986, President Reagan ignited interest in 

the possibility of a hypersonic transport—the National Aero-Space Plane 

(NASP)—dubbed the “Orient Express.” The Battelle Memorial Institute, 

which established the Center for High-Speed Commercial Flight in April 

1986, became a focal point and influential advocate for these proposals.144

NASA had been working with the Defense Advanced Research 

Projects Agency (DARPA) on hypersonic technology for what became the 

NASP since the early 1980s. In February 1987, the OST issued an updated 

National Aeronautical R&D Goals, subtitled “Agenda for Achievement.” 

140. J. Micah Downing, “Lateral Spread of Sonic Boom Measurements from US Air Force Boomfile 

Flight Tests,” High Speed Research: Sonic Boom; Proceedings of a Conference . . . Hampton, VA, 

Feb. 25–27, 1992, vol. 1, NASA CR-3172, pp. 117–129. For a description, see Robert E. Lee 

and Downing, “Boom Event Analyzer Recorder: the USAF Unmanned Sonic Boom Monitor,” AIAA 

Paper 93-4431, Oct. 1993.

141. Plotkin and Maglieri, “Sonic Boom Research,” p. 6.

142. John G. Norris, “AF Says ‘Sonic Boom’ Can Peril Civilians—Might be Used as Weapon,” 

Washington Post, Nov. 9, 1954, pp. 1, 12; Downing, et al., “Measurement of Controlled Focused Sonic Booms from Maneuvering Aircraft,” JASA, vol. 104, no. 1 (July 1998), pp. 112–121.

143. Judy A. Rumerman, NASA Historical Data Book, vol. 6, NASA Space Applications, Aeronautics . . . and Resources, 1979–1988, NASA SP-2000-4012, 2000, pp. 177–178.

144. Conway, High-Speed Flight, pp. 201–215; Paul Proctor, “Conference [sponsored by Battelle] Cites 

Potential Demand for Mach 5 Transports by Year 2000,” Aviation Week, Nov. 10, 1986, pp. 42–46.
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It called for both aggressively pursuing the NASP and developing the 

“fundamental technology, design, and business foundation for a longrange supersonic transport.”145 In response, NASA accelerated its hypersonic research and began a new quest to develop commercially viable 

supersonic technology. This started with contracts to Boeing and Douglas 

aircraft companies in October 1986 for market and feasibility studies on 

what was now named the High-Speed Civil Transport (HSCT), accompanied by several internal NASA assessments. These studies soon ruled 

out hypersonic speeds (above Mach 5) as being impractical for passenger service. Eventually, NASA and its industry partners settled on a 

cruise speed of Mach 2.4.146 Although only marginally faster than the 

Concorde, the HSCT was expected to double its range and carry three 

times as many passengers. Meanwhile, the NASP survived as a NASA–

DOD program (the X-30) until 1994, with its sonic boom potential studied by current and former NASA specialists.147

The contractual studies on the HSCT emphasized the need to resolve 

environmental issues, including the restrictions on cruising over land 

because of sonic booms, before it could meet the goal of efficient longdistance supersonic flight. On January 19–20, 1988, the Langley Center 

hosted a workshop on the status of sonic boom methodology and understanding. Sixty representatives from Government, academia, and industry 

attended—including many of those involved in the SST and SCR efforts 

and several from the Air Force’s NSBIT program. Working groups on sonic 

boom theory, minimization, atmospheric effects, and human response determined that the following areas most needed more research: boom carpets, 

focused booms, high-Mach predictions, atmospheric effects, acceptability 

metrics, signature prediction, and low-boom airframe designs. 

The report from this workshop served as a baseline on the latest 

knowledge about sonic booms and some of the challenges that lay ahead. 

One of these was the disconnect between aerodynamic efficiency and 

lowering shock strength that had long plagued efforts at boom minimization. Simply stated, near-field shockwaves from a streamlined 

145. Rumerman, NASA Historical Data Book, vol. 6, p. 178.

146. Conway, High-Speed Dreams, pp. 218–228; Chambers, Innovations in Flight, p. 50.

147. Maglieri, Victor E. Sothcroft, and John Hicks, “Influence of Vehicle Configurations and Flight 

Profile on X-30 Sonic Booms,” AIAA Paper 90-5224, Oct. 29, 1990; Maglieri, “A Brief Review 

of the National Aero-Space Plane Sonic Booms Final Report,” USAF Aeronautical Systems Center 

TR-94-9344, Dec. 1992.
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airframe coalesce more readily into strong front and tail shocks, while 

the near-field shock waves from a higher-drag airframe are less likely 

to join together, thus allowing a more relaxed N-wave signature. This 

paradox (illustrated by Figure 6) would have to be solved before a lowboom supersonic transport would be both permissible and practical.148

Trying Once More: The High-Speed Research Program

While Boeing and Douglas were reporting on early phases of their HSCT 

studies, the U.S. Congress approved an ambitious new program for HighSpeed Research (HSR) in NASA’s budget for FY 1990. This effort envisioned 

Government and industry sharing the cost, with NASA taking the lead for 

the first several years and industry expanding its role as research progressed. 

(Because of the intermingling of sensitive and proprietary information, much 

of the work done during the HSR program was protected by a limited distribution system, and some has yet to enter the public domain.) Although 

the aircraft companies made some early progress on lower-boom concepts 

for the HSCT, they identified the need for more sonic boom research by 

NASA, especially on public acceptability and minimization techniques, 

before they could design a practical HSCT able to cruise over land.149

Because solving environmental issues would be a prerequisite to 

developing the HSCT, NASA structured the HSR program into two 

phases. Phase I—focusing on engine emissions, noise around airports, 

and sonic booms, as well as preliminary design work—was scheduled 

for 1990–1995. Among the objectives of Phase I were predicting HCST 

sonic boom signatures, determining feasible reduction levels, and finding a scientific basis on which to set acceptability criteria. After hopefully making sufficient progress on the environmental problems, Phase 

II would begin in 1994. With more industry participation and greater 

funding, it would focus on economically realistic airframe and propulsion technologies and was hoped to have extended until 2001.150

148. Darden, et al., Status of Sonic Boom Methodology and Understanding; Proceedings . . . Langley 

Research Center . . . Jan. 19–20, 1988, NASA CP-3027, June 1989. Fig. 6 is copied from p. 6.

149. Boeing Commercial Airplanes, “High-Speed Civil Transport Study; Final Report,” NASA CR-4234, 

Sept. 1989; Douglas Aircraft Company, “1989 High-Speed Civil Transport Studies,” NASA CR-4375, 

May 1991 (published late with an extension). For a summary of Boeing’s early design process, see 

George T. Haglund, “HSCT Designs for Reduced Sonic Boom,” AIAA Paper 91-3103, Sept. 1991. 

150. Allen H. Whitehead, ed., First Annual High-Speed Research Workshop; Proceedings . . . Williamsburg, VA, May 14–16, 1991, NASA CP-10087, Apr. 1992, pt. 1, pp. 5–22, 202.
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When NASA convened its first workshop for the entire High-Speed 

Research program in Williamsburg, VA, from May 14–16, 1991, the headquarters status report on sonic boom technology warned that “the importance of reducing sonic boom cannot be overstated.” One of the Douglas 

studies had projected that even by 2010, overwater-only routes would 

account for only 28 percent of long-range air traffic, but with overland 

cruise, the proposed HSCT could capture up to 70 percent of all such 

traffic. Based on past experience, the study admitted that research on 

low boom designs “is viewed with some skepticism as to its practical 

application. Therefore an early assessment is warranted.”151

NASA, its contractors, academic grantees, and the manufactures 

were already busy conducting a wide range of sonic boom research 

projects. The main goals were to demonstrate a waveform shape that 

could be acceptable to the public, to prove that a viable airplane could 

be built to generate such a waveform, to determine that such a shape 

would not be too badly disrupted during its propagation through the 

atmosphere, and to estimate that the economic benefit of overland supersonic flight would make up for any performance penalties imposed by 

a low-boom design.152

During the next 3 years, NASA and its partners went into a full-court 

press against the sonic boom. They began several dozen major experiments and studies, the results of which were published in reports and 

presented at conferences and workshops dealing solely with the sonic 

boom. These were held at the Langley Research Center in February 

1992,153 the Ames Research Center in May 1993,154 the Langley Center 

in June 1994,155 and again at Langley in September 1995.156 The workshops, like the sonic boom effort itself, were organized into three major 

151. Ibid., pp. 272 and 275 quoted.

152. Ibid., Table of Contents, pp. iv-v.

153. Darden, ed., High-Speed Research: Sonic Boom; Proceeding . . . Langley Research Center . 

. . Feb. 25–27, 1992, NASA CR-3172, Oct. 1992, vols. 1, 2. 

154. Thomas A. Edwards, ed., High-Speed Research: Sonic Boom; Proceedings . . . Ames Research 

Center . . . May 12–14, 1993, NASA CP-10132, vol. 1.

155. David A. McCurdy, ed., High-Speed Research: 1994 Sonic Boom Workshop, Atmospheric 

Propagation and Acceptability Studies; Proceedings . . . Hampton, VA, June 1–3, 1994, NASA 

CP-3209; High-Speed Research: 1994 Sonic Boom Workshop: Configuration, Design, Analysis, 

and Testing . . . Hampton, VA, June 1–3, 1994, NASA CP-209669, Dec. 1999.

156. Daniel G. Baize, 1995 NASA High-Speed Research Program Sonic Boom Workshop: Proceedings . . . Langley Research Center . . . Sept. 12–13, 1995, NASA CP-3335, vol. 1, July 1966.
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Figure 6. Low-boom/high-drag paradox. NASA.

areas of research: (1) configuration design and analysis (managed by 

Langley’s Advanced Vehicles Division), (2) atmospheric propagation, 

and (3) human acceptability (both managed by Langley’s Acoustics 

Division). The reports from these workshops were each well over 500 

pages long and included dozens of papers on the progress or completion of 

various projects.157

The HSR program precipitated major advances in the design of supersonic configurations for reduced sonic boom signatures. Many of these 

were made possible by the new field of computational fluid dynamics 

(CFD). Researchers were now able to use complex computational algorithms processed by supercomputers to calculate the nonlinear aspects of 

near-field shock waves, even at high Mach numbers and angles of attack. 

Results could be graphically displayed in mesh and grid formats that emulated three dimensions. (In simple terms: before CFD, the nonlinear characteristics of shock waves generated by a realistic airframe had involved 

too many variables and permutations to calculate by conventional means.) 

The Ames Research Center, with its location in the rapidly growing 

Silicon Valley area, was a pioneer in applying CFD capabilities to aerodynamics. At the 1991 HSR workshop, an Ames team led by Thomas 

Edwards and including modeling expert Samsun Cheung predicted 

that “in many ways, CFD paves the way to much more rapid progress 

157. For help in deciding which of the many research projects to cover, the author referred to 

Darden, “Progress in Sonic-Boom Understanding: Lessons Learned and Next Steps,” 1994 Sonic 

Boom Workshop, pp. 269–292, for guidance. 
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in boom minimization. . . . Furthermore, CFD offers fast turnaround 

and low cost, so high-risk concepts and perturbations to existing geometries can be investigated quickly.”158

At the same time, Christine Darden and a team that included Robert 

Mack and Peter G. Coen, who had recently devised a computer program 

for predicting sonic booms, used very realistic 12-inch wind tunnel models (the largest yet to measure for sonic boom). Although the model was 

more realistic than previous ones and validating much about the designs, 

including such details as engine nacelles, signature measurements in 

Langley’s 4 by 4 Unitary Wind Tunnel and even Ames 9 by 7 Unitary Wind 

Tunnel still left much to be desired.159 During subsequent workshops and 

at other venues, experts from Ames, Langley and their local contractors 

reported optimistically on the potential of new CFD computer codes—

with names like UPS3D, OVERFLOW, AIRPLANE, and TEAM—to help 

design configurations optimized for both constrained sonic booms and 

aerodynamic efficiency. In addition to promoting the use of CFD, former Langley employee Percy “Bud” Bobbitt of Eagle Aeronautics pointed 

out the potential of hybrid laminar flow control (HLFC) for both aerodynamic and low-boom purposes.160 At the 1992 workshop, Darden and Mack 

admitted how recent experiments at Langley had revealed limitations in 

using near-field wind tunnel data for extrapolating sonic boom signatures.161

Even the numbers-crunching capability of supercomputers was not 

yet powerful enough for CFD codes and the grids they produced to accurately depict effects beyond the near field, but the use of parallel computing 

held the promise of eventually doing so. It was becoming apparent that, 

for most aerodynamic purposes, CFD was the design tool of the future, 

with wind tunnel models becoming more a means of verification. As just 

one example of the value of CFD methods, Ames researchers were able 

to design an airframe that generated a type of multishock signature that 

158. Thomas A. Edwards, et al., “Sonic Boom Prediction and Minimization using Computational 

Fluid Dynamics,” First Annual High Speed Workshop [1991], pt. 2, p. 732. 

159. Darden, et al., “Design and Analysis of Low Boom Concepts at Langley Research Center,” ibid., 

pp. 675–699); Peter G. Coen, “Development of a Computer Technique for Prediction of Transport 

Aircraft Flight Profile Sonic Boom Signatures,” NASA CR-188117, Mar. 1991 (M.S. Thesis, George 

Washington University).

160. Percy J. Bobbit, “Application of Computational Fluid Dynamics and Laminar Flow Technology for 

Improved Performance and Sonic Boom Reduction,” 1992 Sonic Boom Workshop, vol. 2, pp. 137–144.

161. Mack and Darden, “Limitations on Wind-Tunnel Pressure Signature Extrapolation,” 1992 

Sonic Boom Workshop, vol. 2, pp. 201–220.
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might reach the ground with a quieter sonic boom than either the ramp 

or flattop wave forms that were a goal of traditional minimization theories.162 Although not part of the HSCT effort, Ames and its contractors 

also used CFD to continue exploring the possible advantages of obliquewing aircraft, including sonic boom minimization.163

Since neither wind tunnels nor CFD could as yet prove the persistence of waveforms for more than a small fraction of the 200 to 300 body 

lengths needed to represent the distance from an HSCT to the surface, 

Domenic Maglieri of Eagle Aeronautics led a feasibility study in 1992 on 

the most cost effective ways to verify design concepts with realistic testing. After exploring a wide range of alternatives, the team selected the 

Teledyne-Ryan BQM-34 Firebee II remotely piloted vehicle (RPV), which 

the Air Force and Navy had used as a supersonic target drone. Four of 

these 28-feet-long RPVs, which could sustain a speed of Mach 1.3 at 9,000 

feet (300 body lengths from the surface) were still available as surplus. 

Modifying them with low-boom design features such as specially configured 40-inch nose extensions (shown in Figure 7 with projected waveforms from 20,000 feet), could provide far-field measurements needed 

to verify the waveform shaping projected by CFD and wind tunnel models.164 Meanwhile, a complementary plan at the Dryden Flight Research 

Center led to NASA’s first significant sonic boom testing there since the 

1960s. SR-71 program manager David Lux, atmospheric specialist L.J. 

Ehernberger, aerodynamicist Timothy R. Moes, and principal investigator Edward A. Haering came up with a proposal to demonstrate CFD 

design concepts by having one of Dryden’s SR-71s modified with a lowboom configuration. As well as being much larger, faster, and higher-flying 

than the little Firebee, an SR-71 would also allow easier acquisition 

of near-field measurements for direct comparison with CFD predic162. Susan E. Cliff, et al., “Design and Computational/Experimental Analysis of Low Sonic Boom 

Configurations,” 1994 Sonic Boom Workshop, vol. 2, pp. 33–57. For a review of CFD work 

at Ames from 1989–1994, see Samsun Cheung, “Supersonic Civil Airplane Study and Design: 

Performance and Sonic Boom,” NASA CR-197745, Jan. 1995.

163. Christopher A. Lee, “Design and Testing of Low Sonic Boom Configurations and an Oblique 

All-Wing Supersonic Transport,” NASA CR-197744, Feb. 1995.

164. Maglieri, et al., “Feasibility Study on Conducting Overflight Measurements of Shaped Sonic 

Boom Signatures Using the Firebee BQM-34E RPV,” NASA CR-189715, Feb. 1993. Fig. 7 is copied 

from p. 52, with waveforms based on a speed of Mach 1.3 at 20,000 feet rather than the 9,000 feet 

of planned flight tests.
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tions.165 To lay the groundwork for this modification, the Dryden Center 

obtained baseline data from a standard SR-71 using one of its distinctive F-16XL aircraft (built by General Dynamics in the early 1980s for 

evaluation by the Air Force as a long-range strike version of the F-16 

fighter). In tests at Edwards during July 1993, the F-16XL flew as close as 

40 feet below and behind an SR-71 cruising at Mach 1.8 to collect nearfield pressure measurements. Both the Langley Center and McDonnellDouglas analyzed this data, which had been gathered by a standard flighttest nose boom. Both reached generally favorable conclusions about the 

ability of CFD and McDonnell-Douglas’s proprietary MDBOOM program (derived from PCBoom) to serve as design tools. Based on these 

results, McDonnell-Douglas Aerospace West designed modifications to 

reduce the bow and middle shock waves of the SR-71 by reshaping the 

front of the airframe with a “nose glove” and adding to the midfuselage cross-section. An assessment of these modifications by Lockheed 

Engineering & Sciences found them feasible.166 The next step would be 

to obtain the considerable funding that would be needed for the modifications and testing.

In May 1994, the Dryden Center used two of its fleet of F-18 Hornets 

to measure how near-field shockwaves merged to assess the feasibility of a similar low-cost experiment in waveform shaping using two 

SR-71s. Flying at Mach 1.2 with one aircraft below and slightly behind 

the other, the first experiment positioned the canopy of the lower F-18 

in the tail shock extending down from the upper F-18 (called a tailcanopy formation). The second experiment had the lower F-18 fly with 

its canopy in the inlet shock of the upper F-18 (inlet-canopy). Ground 

sensor recordings revealed that the tail-canopy formation caused two 

separate N-wave signatures, but the inlet-canopy formation yielded a 

single modified signature, which two of the recorders measured as a flattop waveform. Even with the excellent visibility from the F-18’s bubble 

canopy (one pilot used the inlet shock wave as a visual cue for positioning 

165. David Lux, et al., “Low-Boom SR-71 Modified Signature Demonstration Program,” 1994 Sonic 

Boom Workshop: Configuration, Design, Analysis and Testing, pp. 237–248. 

166. Edward H. Haering, et al., “Measurement of the Basic SR-71 Airplane Near-Field Signature,” 

1994 Sonic Boom Workshop: Configuration, Design, Analysis, and Testing, pp. 171–197; John 

M. Morgenstern, et al., “SR-71A Reduced Sonic Boom Modification Design,” ibid., pp. 199–217; 

Kamran Fouladi, “CFD Predictions of Sonic-Boom Characteristics for Unmodified and Modified SR71 Configurations,” ibid., pp. 219–235. Fig. 8 is copied from p. 222.
NASA’s Contributions to Aeronautics

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Figure 7. Proposed modifications to BQM-34 Firebee II. NASA.

his aircraft) and its responsive flight controls, maintaining such precise positions was still not easy, and the pilots recommended against 

trying to do the same with the SR-71, considering its larger size, slower 

response, and limited visibility.167

Atmospheric effects had long posed many uncertainties in understanding sonic booms, but advances in acoustics and atmospheric science since the SCR program promised better results. Scientists needed 

a better understanding not only of the way air molecules absorb sound 

waves, but also old issue of turbulence. In addition to using the Air 

Force’s Boomfile and other available material, Langley’s Acoustic Division 

had Eagle Aeronautics, in a project led by Domenic Maglieri, restore 

and digitize data from the irreplaceable XB-70 records.168

The division also took advantage of the NATO Joint Acoustic 

Propagation Experiment (JAPE) at the White Sands Missile Range in 

August 1991 to do some new testing. The researchers arranged for F-15, 

F-111, T-38 aircraft, and one of Dryden’s SR-71s to make 59 supersonic 

passes over an extensive array of BEAR, other recording systems, and 

meteorological sensors—both early in the morning (when the air was 

still) and during the afternoon (when there was usually more turbulence). 

Although meteorological data were incomplete, results later showed 

167. Catherine M. Bahm and Edward A. Haering, “Ground-Recorded Sonic Boom Signatures of 

F-18 Aircraft in Formation Flight,” 1995 Sonic Boom Workshop, vol. 1, pp. 220–243.

168. J. Micah Downing, “Lateral Spread of Sonic Boom Measurements from US Air Force Boomfile 

Flight Tests,” 1992 Sonic Boom Workshop, vol. 1, pp. 117–136; Maglieri, et al., “A Summary of 

XB-70 Sonic Boom Signature Data, Final Report,” NASA CR-189630, Apr. 1992.
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Figure 8. Proposed SR-71 low-boom modification. NASA.

the effects of molecular relaxation and turbulence on both the rise time 

and overpressure of bow shocks.169 Additional atmospheric information 

came from experiments on waveform freezing (persistence), measuring diffraction and distortion of sound waves, and trying to discover the 

actual relationship among molecular relaxation, turbulence, humidity, 

and other weather conditions.170

Leonard Weinstein of the Langley Center even developed a way 

to capture images of shock waves in the real atmosphere. He did this 

using a ground-based schlieren system (a specially masked and filtered tracking camera with the Sun providing backlighting). As shown 

in the accompanying photo, this was first demonstrated in December 

1993 with a T-38 flying just over Mach 1 at Wallops Island.171 All of the 

research into the theoretical, aerodynamic, and atmospheric aspects of 

sonic boom—no matter how successful—would not protect the Achilles’ 

heel of previous programs: the subjective response of human beings. 

169. William L. Willshire and David W. DeVilbiss, “Preliminary Results from the White Sands Missile 

Range Sonic Boom Propagation Experiment,” 1992 Sonic Boom Workshop, vol. 1, pp. 137–144.

170. Gerry L. McAnich, “Atmospheric Effects on Sonic Boom—A Program Review,” First Annual HSR 

Workshop, pp. 1201–1207; Allan D. Pierce and Victor W. Sparrow, “Relaxation and Turbulence 

Effects on Sonic Boom Signatures,” ibid., pp. 1211–1234; Kenneth J. Plotkin, “The Effect of Turbulence 

and Molecular Relaxation on Sonic Boom Signatures,” ibid., pp. 1241–1261; Lixin Yao, et al., “Statistical and Numerical Study of the Relation Between Weather and Sonic Boom,” ibid., pp. 1263–1284.

171. Leonard M. Weinstein, “An Optical Technique for Examining Aircraft Shock Wave Structures 

in Flight,” 1994 Sonic Boom Workshop, Atmospheric Propagation, pp. 1–18. The following year 

Weinstein demonstrated improved results using a digital camera: “An Electronic Schlieren Camera 

for Aircraft Shock Wave Visualization,” 1995 Sonic Boom Workshop, vol. 1, pp. 244–258.
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As a result, the Langley Center, led by Kevin Shepherd of the Acoustics 

Division, had begun a systematic effort to measure human responses 

to different strengths and shapes of sonic booms and hopefully determine a tolerable level for community acceptance. As an early step, 

the division built an airtight foam-lined sonic boom simulator booth 

(known as the “boom box”) based on one at the University of Toronto. 

Using the latest in computer-generated digital amplification and loudspeaker technology, it was capable of generating shaped waveforms up 

to 4 psf (140 decibels). Based on responses from subjects, researchers 

selected the perceived-level decibel (PLdB) as the preferred metric. For 

responses outside a laboratory setting, Langley planned several additional 

acceptance studies.172

By 1994, early results had become available from two of these projects. The Langley Center and Wyle Laboratories had developed mobile 

boom simulator equipment called the In-Home Noise Generation/

Response System (IHONORS). It consisted of computerized sound systems installed in 33 houses for 8 weeks at a time in a network connected 

by modems to a monitor at Langley. From February to December 1993, 

these households were subjected to almost 58,500 randomly timed sonic 

booms of various signatures for 14 hours a day. Although definitive analyses were not available until the following year, the initial results confirmed how the level of annoyance increased whenever subjects were 

startled or trying to rest.173

Preliminary results were also in from the first phase of the Western 

USA Sonic Boom Survey of civilians who had been exposed to such 

sounds for many years. This part of the survey took place in remote desert towns around the Air Force’s vast Nellis combat training range complex in Nevada. Unlike previous community surveys, it correlated citizen 

responses to accurately measured sonic boom signatures (using BEAR 

devices) in places where booms were a regular occurrence, yet where 

the subjects did not live on or near a military installation (i.e., where 

172. Kevin P. Shepherd, “Overview of NASA Human Response to Sonic Boom Program,” First 

Annual HSR Workshop, pt. 3, pp. 1287–1291; Shepherd, et al., “Sonic Boom Acceptability Studies,” ibid., pp. 1295–1311.

173. David A. McCurdy, et al., “An In-Home Study of Subjective Response to Simulated Sonic 

Booms,” 1994 Sonic Boom Workshop: Atmospheric Propagation and Acceptability, pp. 193–

207; McCurdy and Sherilyn A. Brown, “Subjective Response to Simulated Sonic Boom in Homes,” 

1995 Sonic Boom Workshop, pp. 278–297.
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Leonard Weinstein’s innovative schlieren photograph showing shock waves emanating from a 

T-38 flying Mach 1.1 at 13,000 feet, December 1993. NASA.

the economic benefits of the base for the local economy might influence 

their opinions). Although findings were not yet definitive, these 1,042 

interviews proved more decisive than any of the many other research 

projects in determining the future direction of the HSCT effort. Based 

on a metric called day-night average noise level, the respondents found 

the booms much more annoying than previous studies on other types 

of aircraft noise had, even at the levels projected for low-boom designs. 

Their negative responses, in effect, dashed hopes that the HSR program 

might lead to an overland supersonic transport.174

Well before the paper on this survey was presented at the 1994 Sonic 

Boom Workshop, its early findings had prompted NASA Headquarters to 

reorient High-Speed Research toward an HSCT design that would only 

fly supersonic over water. Just as with the AST program 20 years earlier, this became the goal of Phase II of the HSR program (which began 

using FY 1994 funding left over from the canceled NASP).175

At the end of the 1994 workshop, Christine Darden discussed lessons 

learned so far and future directions. While the design efforts had shown 

outstanding progress, dispersal of the work among two NASA Centers 

174. James M. Fields, et al., “Residents’ Reactions to Long-Term Sonic Boom Exposure: Preliminary Results,” 

1994 Sonic Boom Workshop: Atmospheric Propagation and Acceptability, vol. 1, pp. 193–217.

175. Conway, High-Speed Dreams, p. 253.
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and two major aircraft manufacturers had resulted in communication 

problems as well as a certain amount of unhelpful competition. The milestone-driven HSR effort required concurrent progress in various different areas, which is inherently difficult to coordinate and manage. And 

even if low-boom airplane designs had been perfected to meet acoustic 

criteria, they would have been heavier and suffer from less acceptable 

low-speed performance than unconstrained designs. Under the new HSR 

strategy, any continued minimization research would now aim at lowering the sonic boom of the “baseline” overwater design, while propagation 

studies would concentrate on predicting boom carpets, focused booms, 

secondary booms, and ground disturbances. In view of the HSCT’s overwater mission, new environmental studies would devote more attention to the penetration of shock waves into water and the effect of sonic 

booms on marine mammals and birds.176

Although the preliminary results of the first phase of the Western 

USA Survey had already had a decisive impact, Wyle Laboratories completed Phase II with a similar polling of civilians in Mojave Desert communities exposed regularly to sonic booms, mostly from Edwards AFB. 

Surprisingly, this phase of the survey found the people there much more 

amenable to sonic booms than the desert dwellers in Nevada were, but 

they were still more annoyed by booms than by other aircraft noise of 

comparable perceived loudness.177

With the decision to end work on a low-boom HSCT, the proposed 

modifications of the Firebee RPVs and SR-71 had of course been canceled (postponing for another decade any full-scale demonstrations of 

boom shaping). Nevertheless, some testing continued that would prove 

of future value. From February through April 1995, the Dryden Center 

conducted more SR-71 and F-16XL sonic boom flight tests. Led by Ed 

Haering, this experiment included an instrumented YO-3A light aircraft 

from the Ames Center, an extensive array of various ground sensors, a 

network of new differential Global Positioning System (GPS) receivers 

accurate to within 12 inches, and installation of a sophisticated new nose 

boom with four pressure sensors on the F-16XL. On eight long missions, 

one of Dryden’s SR-71s flew at speeds between Mach 1.25 and Mach 

176. Darden, “Progress in Sonic-Boom Understanding: Lessons Learned and Next Steps,” 1994 

Sonic Boom Workshop, Atmospheric Propagation and Acceptability, pp. 269–290.

177. James M. Fields, “Reactions of Residents to Long-Term Sonic Boom Noise Environments,” 

NASA CR-201704, June 1997.
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1.6 at 31,000–48,000 feet, while the F-16XL (kept aloft by in-flight refuelings) made numerous near- and mid-field measurements at distances from 

80 to 8,000 feet. Some of these showed that the canopy shock waves were 

still distinct from the bow shock after 4,000–6,000 feet. Comparisons of 

far-field measurements obtained by the YO-3A flying at 10,000 feet above 

ground level and the recording devices on the surface revealed effects of 

atmospheric turbulence. Analysis of the data validated two existing sonic 

boom propagation codes and clearly showed how variations in the SR-71’s 

gross weight, speed, and altitude caused differences in shock wave patterns and their coalescence into N-shaped waveforms.178

This successful experiment marked the end of dedicated sonic boom 

flight-testing during the HSR program.

By late 1998, a combination of economic, technological, political, and budgetary problems (including NASA’s cost overruns for 

the International Space Station) convinced Boeing to cut its support 

and the Administration of President Bill Clinton to terminate the HSR 

program at the end of FY 1999. Ironically, NASA’s success in helping 

the aircraft industry develop quieter subsonic aircraft, which had the 

effect of moving the goalpost for acceptable airport noise, was one of 

the factors convincing Boeing to drop plans for the HSCT. Nevertheless, 

the HSR program was responsible for significant advances in 

technologies, techniques, and scientific knowledge, including a better 

understand of the sonic boom and ways to diminish it.179

Proof at Last: The Shaped Sonic Boom Demonstration

After the HSR program dropped plans for an overland supersonic airliner, Domenic Maglieri compiled a NASA study of all known proposals 

178. Edward A. Haering, L.J. Ehernberger, and Stephen A. Whitmore, “Preliminary Airborne 

Experiments for the SR-71 Sonic Boom Propagation Experiment,” 1995 Sonic Boom Workshop, vol. 

1, pp. 176–198; Stephen R. Norris, Haering, and James E. Murray, “Ground-Bases Sensors for the 

SR-71 Sonic Boom Propagation Experiment,” ibid., pp. 199–218; Hugh W. Poling, “Sonic Boom 

Propagation Codes Validated by Flight Test,” NASA CR-201634, Oct. 1996. 

179. Conway, High-Speed Dreams, pp. 286–300; James Schultz, “HSR Leaves Legacy of Spinoffs,” 

Aerospace America, vol. 37, no. 9 (Sept. 1999), pp. 28–32. The Acoustical Society held its third sonic 

boom symposium in Norfolk from Oct. 15–16, 1998. Because of HSR distribution limitations, many of 

the presentations could be oral only, but a few years later, the ASA was able to publish some of them in 

a special edition of its journal. For a status report as of the end of the HSR, see Kenneth J. Plotkin, “State 

of the Art of Sonic Boom Modeling,” JASA, vol. 111, No. 1, pt. 3 (Jan. 2002), pp. 530–536.
NASA’s Contributions to Aeronautics

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for smaller supersonic aircraft intended for business customers.180 In 1998, 

one year after the drafting of this report, Richard Seebass (now with the 

University of Colorado) gave some lectures at NATO’s von Kármán Institute 

in Belgium. He reflected on NASA’s conclusion that a practical, commercialsized supersonic transport would have a sonic boom that was not acceptable to enough people. On the other hand, he believed the recent high-speed 

research “leads us to conclude that a small, appropriately designed supersonic business jet’s sonic boom may be nearly inaudible outdoors and 

hardly discernible indoors.” Such an airplane, he stated, “appears to have 

a significant market . . . if . . . certifiable over most land areas.”181

At the start of the new century, the prospects for a small supersonic 

aircraft received a shot in the arm from the Defense Advanced Research 

Projects Agency, well known for encouraging innovative technologies. 

DARPA received $7 million in funding starting in FY 2001 to explore design 

concepts for a Quiet Supersonic Platform (QSP)—an airplane that could 

have both military and civilian potential. Richard W. Wlezien, a NASA 

official on loan to DARPA as QSP program manager, wanted ideas that 

might lead to a Mach 2.4, 100,000-pound aircraft that “won’t rattle your 

windows or shake the china in your cabinet.” It was hoped that a shaped 

sonic boom signature of no more than 0.3 psf would allow unrestricted 

operations over land. By the end of 2000, 16 companies and laboratories 

had been selected to participate in the QSP project, with the University 

of Colorado and Stanford University to work on sonic boom propagation 

and minimization.182 Support from NASA would include modeling expertise, wind tunnel facilities, and flight-test operations.

Although the later phase of the QSP program emphasized military 

requirements, its most publicized achievement was the Shaped Sonic 

Boom Demonstration (SSBD). This was not one of its original components. 

180. The study, originally completed in 1997, was in the process of being formally published by 

NASA as “A Compilation and Review of Supersonic Business Jet Studies from 1960–1995” as this 

history was being written.

181. Richard Seebass, “History and Economics of, and Prospects for, Commercial Supersonic 

Transport,” (Paper 1) and “Sonic Boom Minimization” (Paper 6), NATO Research and Technology 

Organization, Fluid Dynamics Research on Supersonic Aircraft [proceedings . . . Rhode SaintGenèse, Belgium, May 25–29, 1998], RTO-EN-4, Nov. 1998 (pp. 1–6 of Paper 1 and abstract 

of Paper 2 quoted). Sadly, Seebass would not live to see a low boom airplane configuration finally 

demonstrated in 2003. 

182. Robert Wall, “Darpa Envisions New Supersonic Designs,” Aviation Week, Aug. 28, 2000, 

p. 47; and “Novel Technologies in Quest for Quiet Flight,” Aviation Week, Jan. 8, 2001, p. 61.
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In 1995, the Dryden Flight Research Center used an F-16XL to make detailed in-flight supersonic 

shock wave measurements as near as 80 feet from an SR-71. NASA.

Resurrecting an idea from the HSR program, Domenic Maglieri and colleagues at Eagle Aeronautics recommended that DARPA include a flighttest program using the BQM-34E Firebee II as a proof-of-concept for 

the QSP’s sonic boom objectives. Liking this idea, Northrop Grumman 

Corporation (NGC) wasted no time in acquiring the last remaining 

Firebee IIs from the Naval Air Weapons Station at Point Mugu, CA, but 

later determined that they were now too old for test purposes. As an 

alternative, NGC aerodynamicist David Graham recommended using 

different versions of the Northrop F-5 (which had been modified into 

larger training and reconnaissance models) for sonic boom comparisons. Maglieri then suggested modifications to an F-5E that could flatten its sonic boom signature. Based largely on NGC’s proposal for an 

F-5E Shaped Sonic Boom Demonstration, DARPA in July 2001 selected 

it over QSP proposals from the other two system integrators, Boeing 

Phantom Works and Lockheed Martin’s Skunk Works.183

In designing the modifications, a Northrop Grumman team in El 

Segundo, CA, led by David Graham, benefited from its partnership with 

183. Joseph W. Pawlowski, David H. Graham, Charles H. Boccadoro (NGC), Peter G. Coen (LaRC), 

and Domenic J. Maglieri (Eagle Aero.), “Origins and Overview of the Shaped Sonic Boom Demonstration Program,” AIAA Paper 2005-5, presented at the 43rd Aerospace Sciences Meeting, Reno, NV, 

Jan. 10–13, 2005, pp. 3–7 (also published with briefing slides by the Air Force Research Laboratory 

as AFRL-VA-WP-2005-300, Jan. 2005).
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a multitalented working group. This team included Kenneth Plotkin 

of Wyle Laboratories, Domenic Maglieri and Percy Bobbitt of Eagle 

Aeronautics, Peter G. Coen and colleagues at the Langley Center, John 

Morgenstern of Lockheed Martin, and other experts from Boeing, 

Gulfstream, and Raytheon. They applied knowledge gained from the 

HSR with the latest in CFD technology to begin design of a nose extension and other modifications to reshape the F-5E’s sonic boom. The moderate size and flexibility of the basic F-5E design, which had allowed 

different configurations in the past, made it the perfect choice for the 

SSBD. The shaped-signature modifications (which harked back to the 

stillborn SR-71 proposal of the HSR program) were tested in a supersonic 

wind tunnel at NASA’s Glenn Research Center with favorable results.184

In further preparation for the SSBD, the Dryden Center conducted the 

Inlet Spillage Shock Measurement (ISSM) experiment in February 2002. 

One of its F-15Bs equipped with an instrumented nose boom gathered 

pressure data from a standard F-5E flying at about Mach 1.4 and 32,000 

feet. The F-15 did these probes at separation distances ranging from 60 

to 1,355 feet. In addition to serving as a helpful “dry run” for the planned 

demonstration, the ISSM experiment proved to be of great value in validating and refining Grumman’s proprietary GCNSfv CFD code (based 

on the Ames Center’s ARC3D code), which was being used to design the 

SSBD configuration. Application of the flight test measurements nearly 

doubled the size of the CFD grid, to approximately 14 million points.185

For use in the Shaped Sonic Boom Demonstration, the Navy loaned 

Northrop Grumman one of its standard F-5Es, which the company began 

to modify at its depot facility in St. Augustine, FL, in January 2003. 

Under supervision of the company’s QSP program manager, Charles 

Boccadoro, NGC technicians installed a nose glove and 35-foot fairing 

under the fuselage (resulting in a “pelican-shaped” profile). The modifications, which extended the plane’s length from 46 to approximately 

50 feet, were designed to strengthen the bow shock but weaken and 

stretch out the shock waves from the cockpit, inlets, and wings—keep184. Ibid., p. 8; Edward D. Flinn, “Lowering the Boom on Supersonic Flight Noise,” Aerospace 

America, vol. 40, No. 2 (Feb. 2002), pp. 20–21; Wyle Laboratories, “Wyle Engineers Play Significant Role in Northrop Grumman Sonic Boom Test Program,” News Release 09-11, Sept. 11, 2003.

185. Keith H. Meredith, et al., “Computational Fluid Dynamics Comparison and Flight Test Measurement of F-5E Off-Body Pressures,” AIAA Paper 2005-6, presented at 43rd Aerospace Sciences 

Meeting, Reno, NV, Jan. 10–13, 2005.
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ing them from coalescing to form the sharp initial peak of the N-wave 

signature.186 After checkout flights in Florida starting on July 25, 2003, 

the modified F-5E, now called the SSBD F-5E, arrived in early August 

at Palmdale, CA, for more functional check flights. 

On August 27, 2003, on repeated runs through an Edwards supersonic corridor, the SSBD F-5E, piloted by NGC’s Roy Martin, proved 

for the first time that—as theorized since the 1960s—a shaped sonic 

boom signature from a supersonic aircraft could persist through the 

real atmosphere to the ground. Flying at Mach 1.36 and 32,000 feet on 

an early-morning run, the SSBD F-5E was followed 45 seconds later 

by an unmodified F-5E from the Navy’s aggressor training squadron at 

Fallon, NV. They flew over a high-tech ground array of various sensors 

manned by personnel from Dryden, Langley, and almost all the organizations in the SSBD working group. Figure 9 shows the subtle but significant difference between the flattened waveform from the SSBD F-5E 

(blue) and the peaked N-wave from its unmodified counterpart (red) as 

recorded by a Base Amplitude and Direction Sensor (BADS) on this historic occasion. As a bonus, the initial rise in pressure of the shaped signature was only about 0.83 psf as compared with the 1.2 psf from the 

standard F-5E—resulting in a much quieter sonic boom.187

During the last week of August, the two F-5Es flew three missions to 

provide many more comparative sonic boom recordings. On two other missions, using the technique developed for the SR-71 during HSR, a Dryden 

F-15B with a specially instrumented nose boom followed the SSBD-modified 

F-5E to gather near-field measurements. The data from the F-15B probing 

missions showed how the F-5E’s modifications changed its normal shock 

wave signature, which data from the ground sensors revealed as persisting down through the atmosphere to consistently produce the quieter flattopped sonic boom signatures. The SSBD met expectations, but unusually 

high temperatures (even for the Antelope Valley in August) limited the 

top speed and endurance of the F-5Es. Because of this and a desire to 

gather more data on maneuvers and different atmospheric conditions, 

186. Graham Warwick, “F-5E Shapes Up to Change Sonic Boom,” Flight International, Aug. 

5, 2003, p. 30; T.A. Heppenheimer, “The Boom Stops Here,” Air and Space Magazine, Nov. 

2005, http://www.airspacemag.com/fight-today/boom.html. 

187. Pawlowski, et al., “Origins and Overview of the SSBD,” pp. 10–12; Peter G. Coen and Roy 

Martin, “Fixing the Sound Barrier: The DARPA/NASA/Northrop-Grumman Shaped Sonic Boom Flight 

Demonstration,” Briefing at EAA AirVenture, Oshkosh, WI, July 2004. (Fig. 9 is taken from slide 21.)
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Peter Coen, Langley’s manager for supersonic vehicles technology, and 

researchers at Dryden led by SSBD project manager David Richwine 

and principal investigator Ed Haering, began planning a NASA-funded 

Shaped Super Boom Experiment (SSBE) to follow up on the SSBD.188

NASA successfully conducted the SSBE with 21 more flights during 11 

days in January 2004. These met or exceeded all test objectives. Eight of 

these flights were again accompanied by an unmodified Navy F-5E from 

Fallon, while Dryden’s F-15B flew four more probing flights to acquire 

additional near-field measurements. An instrumented L-23 sailplane from 

the USAF Test Pilot School obtained boom measurements from 8,000 feet 

(well above the ground turbulence layer) on 13 flights. All events were precisely tracked by differential GPS receivers and Edwards AFB’s extensive 

telemetry system. In all, the SSBE yielded over 1,300 sonic boom signature 

recordings and 45 probe datasets—obtaining more information about the 

effects of turbulence, helping to confirm CFD predictions and wind tunnel validations, and bequeathing a wealth of data for future engineers and 

designers.189 In addition to a series of scientific papers, the SSBD–SSBE 

accomplishments were the subject of numerous articles in the trade and 

popular press, and participants presented well-received briefings at various aeronautics and aviation venues. 

Breaking Up Shock Waves with “Quiet Spike”

In June 2003, the FAA—citing a finding by the National Research Council 

that there were no insurmountable obstacles to building a quiet supersonic aircraft—began seeking comments on its noise standards in 

advance of a technical workshop on the issue. In response, the Aerospace 

Industries Association, the General Aviation Manufactures Association, 

and most aircraft companies felt that the FAA’s sonic boom restriction 

188. Pawlowski, et al., “Origins and Overview of the SSBD,” pp. 11–12; NASA Press Release 

03-50, “NASA Opens New Chapter in Supersonic Flight,” Sept. 4, 2003; Gary Creech, “NASA, 

Northrop Study Sonic Boom Reduction,” Dryden X-Press, vol. 46, issue 2 (Mar. 2004). p. 1.

189. Pawlowski, et al., “Origins and Overview of the SSBD,” pp 12–13; David H. Graham, et 

al., “Wind Tunnel Validation of Shaped Sonic Boom Demonstration Aircraft Design,” AIAA Paper 

2005-7; Haering, et al., “Airborne Shaped Sonic Boom Demonstration Pressure Measurements with 

Computational Fluid Dynamics Comparisons,” AIAA Paper 2005-9; Plotkin, et al., “Ground Data 

Collection of Shaped Sonic Boom Experiment Aircraft Pressure Signatures,” AIAA Paper 2005-10; 

John M. Morgenstern, et al., “F-5 Shaped Sonic Boom Demonstrator’s Persistence of Boom Shaping 

Reduction through Turbulence,” AIAA 2005-12; all papers presented at 43rd Aerospace Sciences 

Meeting, Reno, NV, Jan. 10–13, 2005. 
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was the still the most serious impediment to creating the market for a 

supersonic business jet (SSBJ), which would be severely handicapped 

if unable to fly faster than sound over land.190

By the time the FAA workshop was held in mid-November, Peter 

Coen of the Langley Center and a Gulfstream vice president were able 

to report on the success of the SSBD. Coen also outlined future initiatives in NASA’s Supersonic Vehicles Technology program. In addition to 

leveraging the results of DARPA’s QSP research, NASA hoped to engage 

industry partners for follow-on projects on the sonic boom, and was also 

working with Eagle Aeronautics on new three-dimensional CFD boom 

propagation models. For additional psychoacoustical studies, Langley 

had reconditioned its boom simulator booth. And as a possible followup 

to the SSBD, NASA was considering a shaped low-boom demonstrator 

that could fly over populated areas, allowing definitive surveys on public acceptance of minimized boom signatures.191

The Concorde made its final transatlantic flights just a week after 

the FAA’s workshop. Its demise marked the first time in modern history that a mode of transportation had retreated back to slower speeds. 

This did, however, leave the future supersonic market entirely open to 

business jets. Although the success of the SSBD hinted at the feasibility of such an aircraft, designing one—as explained in a new study by 

Langley’s Robert Mack—would still not be at all easy.192

During the next several years, a few individual investors and a number 

of American and European aircraft companies—including Gulfstream, 

Boeing, Lockheed, Cessna, Raytheon, Dassault, Sukhoi, and the privately 

held Aerion Corporation—pursued assorted SSBJ concepts with varying 

degrees of cooperation, competition, and commitment. Some of these and 

other aviation-related companies also worked together on supersonic 

strategies through three consortiums: Supersonic Aerospace International 

190. James R. Asker, “FAA Seeks Information on Sonic Boom Research,” Aviation Week, June 2, 

2003, p. 21; David Bond, “The Time is Right,” Aviation Week, Oct. 20, 2003, pp. 57–58.

191. Peter G. Coen, “Supersonic Vehicles Technology: Sonic Boom Technology Development and 

Demonstration” and Preston A. Henne, “A Gulfstream Perspective on the DARPA QSP Program and Future 

Supersonic Initiatives,” Briefing Slides, FAA Civil Supersonic Aircraft Workshop, Washington, DC, Nov. 

13, 2003; Aimee Cunningham, “Sonic Booms and Human Ears: How Much Can the Public Tolerate,” 

Popular Science, July 30, 2004, http://www.popsci.com/military-aviation-space/article/2004-07/

sonic-booms-and-human-ears.

192. Robert J. Mack, “A Supersonic Business Jet Concept Designed for Low Sonic Boom,” NASA 

TM-2003-212435, Oct. 2003.
NASA’s Contributions to Aeronautics

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Figure 9. Normal and flattened F-5E sonic boom signatures. NASA.

(SAI), which had support from Lockheed-Martin; the 10-member 

Supersonic Cruise Industry Alliance (SCIA); and Europe’s High-Speed 

Aircraft Industrial Project (HISAC), comprising more than 30 companies, 

universities, and other members. Meanwhile, the FAA began the lengthy 

process for considering a new metric on acceptable sonic booms and, in 

the interest of global consistency, prompted the International Civil Aviation 

Organization (ICAO) to also put the issue on its agenda. It was in this 

environment of both renewed enthusiasm and ongoing uncertainty about 

commercial supersonic flight that NASA continued to study and experiment on ways to make the sonicboom more acceptable to the public.193

Richard Wlezien (back from DARPA as NASA’s vehicle systems manager) hoped to follow up on the SSBD with a truly low-boom super193. For examples, see Graham Warwick, “Quiet Progress: Aircraft Designers Believe They Can Take 

the Loud Boom out of Supersonic Travel,” Flight International, Oct. 20, 2004, pp. 32–33; Edward H. 

Phillips, “Boom Could Doom: Debate over Hybrid SSBJ Versus Pure Supersonic Is Heating Up,” Aviation 

Week, June 13, 2005, pp. 84–85; Francis Fiorino, “Lowering the Boom,” Aviation Week, Nov. 7, 

2005, p. 72; “Supersonic Private Jets in Development,” Business Travel News Online, Oct. 23, 2006; 

John Wiley, “The Super-Slow Emergence of Supersonic,” Business and Commercial Aviation, Sept. 1, 

2007, pp. 48–50; Edward H. Phillips, “Shock Wave: Flying Faster than Sound Is the Holy Grail of Business Aviation,” Aviation Week, Oct. 8, 2007, pp. 50–51; Mark Huber, “Mach 1 for Millionaires,” Air 

and Space Magazine, Mar.–Apr. 2006, http://www.airspacemag.com/flight-today/millionaire.html.
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sonic demonstrator, possibly by 2010. In July 2005, NASA announced the 

Sonic Boom Mitigation Project, which began with concept explorations by 

major aerospace companies on the feasibility of either modifying another 

existing aircraft or designing a new demonstrator.194 As explained by Peter 

Coen, “these studies will determine whether a low sonic boom demonstrator can be built at an affordable cost in a reasonable amount of time.”195

Although numerous options for using existing aircraft were under investigation, most of the studies were leaning toward the need to build a new 

experimental airplane as the most effective solution. On August 30, 2005, 

however, NASA Headquarters announced the end of the short-lived Sonic 

Boom Mitigation Project because of changing priorities.196

Despite this setback, there was still one significant boom lowering 

experiment in the making. Gulfstream Aerospace Corporation, which 

had been teamed with Northrop Grumman in one of the canceled studies, 

had already patented a new sonic boom mitigation technique.197 Testing 

this invention—a retractable lance-shaped device to extend the length of 

an aircraft—would become the next major sonic boom flight experiment. 

In the meantime, NASA continued some relatively modest sonic boom 

testing at the Dryden Center, mainly to help improve simulation capabilities. In a joint project with the FAA and Transport Canada in the summer 

of 2005, researchers from Pennsylvania State University strung an array of 

advanced microphones at Edwards AFB to record sonic booms created by 

Dryden F-18s passing overhead. Eighteen volunteers, who sat on lawn chairs 

alongside the row of microphones during the flyovers to experience the real 

thing, later gauged the fidelity of the played-back recordings. These were then 

used to help improve the accuracy of the booms replicated in simulators.198

“Quiet Spike” was the name that Gulfstream gave to its nose boom 

concept. Based on CFD models and results from Langley’s 4 by 4 super194. Graham Warwick, “NASA Narrows R&D Agenda,” Flight International, Feb. 15, 2006, p. 

28; Ellen H. Thompson, et al., “NASA Funds Studies for Quieter Supersonic Boom,” NASA News 

Release 05-176, July 8, 2005. 

195. David Collogan, “Manufacturers, NASA working on Bizjet Sonic Boom Project,” The Weekly 

of Business Aviation, July 18, 2005, p. 21.

196. Michael A. Dornheim, “Will Low Boom Fly? NASA Cutbacks Delay Flight Test of Shaped 

Demonstrator . . . ,” Aviation Week, Nov. 7, 2005, pp. 68–69.

197. U.S. Patent No. 6,698,684, “Supersonic Aircraft with Spike for Controlling and Reducing 

Sonic Boom,” Mar. 2, 2004, http://www.patentstorm.us/patents/6698684/description.html.

198. Jay Levine, “Lowering the Boom,” July 29, 2005, http://www.nasa.gov/centers/dryden/

news/X-Press/stories/2005/072905. 
NASA’s Contributions to Aeronautics

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Close-up view of the SSBD F-5E, showing its enlarged “pelican” nose and lower fuselage designed 

to shape the shock waves from the front of the airframe. NASA.

sonic wind tunnel, Gulfstream was convinced that the Quiet Spike device 

could greatly mitigate a sonic boom by breaking up the typical nose shock 

into three less-powerful waves that would propagate in parallel to the 

ground.199 However, the company needed to test the structural and aerodynamic suitability of the device and also obtain supersonic in-flight data 

on its shock scattering ability. NASA’s Dryden Flight Research Center had 

the capabilities needed to accomplish these tasks. Under this latest publicprivate partnership, Gulfstream fabricated a telescoping 30-foot-long nose 

boom (made of molded graphite epoxy over an aluminum frame) to attach 

to the radar bulkhead of Dryden’s frequently modified F-15B No. 836. A 

motorized cable and pulley system could extend the spike up to 24 feet and 

retract it back to 14 feet. After extensive static testing at its Savannah, GA, 

facility, Gulfstream and NASA technicians at Dryden attached the specially 

instrumented spike to the F-15’s radar bulkhead in April 2006 and began 

conducting further ground tests, such as for vibration.200

199. Donald C. Howe, et al., “Development of the Gulfstream Quiet Spike for Sonic Boom Minimization,” AIAA Paper 2008-124, presented at 46th Aerospace Sciences Meeting, Reno, NV, Jan. 7–10, 

2008; Natalie D. Spivey, et al., “Quiet Spike Build-up Ground Vibration Testing Approach,” NASA 

TN-2007-214625, Nov. 2007.

200. James W. Smolka, et al., “Flight Testing of the Gulfstream Quiet Spike on a NASA F-15B,” 

paper presented to the Society of Experimental Test Pilots, Anaheim, CA, Sept. 27, 2007, 1-24; 

Stephen B. Cumming, et al., “Aerodynamic Effects of a 24-foot Multi-segmented Telescoping Nose 

Boom on an F-15B Airplane,” NASA TM-2008-214634, Apr. 2008.
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After various safety checks, aerodynamic assessments and checkout 

flights, Dryden conducted Quiet Spike flight tests from August 10, 2006 

until February 14, 2007. Key engineers on the project included Dryden’s 

Leslie Molzahn and Thomas Grindle, and Gulfstream’s Robbie Cowart. 

Veteran NASA test pilot Jim Smolka gradually expanded the F-15B’s flight 

envelope up to Mach 1.8 and performed sonic boom experiments with 

the telescoping nose boom at speeds up to Mach 1.4 at 40,000 feet. Aerial 

refueling by AFFTC’s KC-135 allowed extended missions with multiple 

test points. Because it was known that the weak shock waves from the 

spike would rather quickly coalesce with the more powerful shock waves 

generated by the rest of the F-15’s unmodified high-boom airframe, data 

were collected from distances of no more than 1,000 feet. These measurements, made by a chase plane using similar probing techniques to 

those of the SR-71 and SSBD tests, confirmed CFD models on the spike’s 

ability to generate a sawtooth wave pattern that, if reaching the surface, 

would cause only a muffled sonic boom. Analysis of the data appeared 

to confirm that shocks of equal strength would not coalesce into a single strong shock. In February 2007, with all major test objectives having been accomplished, the Quiet Spike F-15B was flown to Savannah 

for Gulfstream to restore to its normal configuration.201

For this successful test of an innovative design concept for a future 

SSBJ, James Smolka, Leslie Molzahn, and three Gulfstream employees 

subsequently received Aviation Week and Space Technology’s Laureate 

Award in Aeronautics and Propulsion. One month later, however, both 

the Gulfstream Corporation and the Dryden Center were saddened by 

the death in an airshow accident of Gerard Schkolnik, Gulfstream’s 

Director of Supersonic Technology Programs, who had been a Dryden 

employee for 15 years.202

Focusing on Fundamentals: The Supersonics Project

In January 2006, NASA Headquarters announced its restructured 

aeronautics mission. As explained by Associate Administrator for 

Aeronautics Lisa J. Porter, “NASA is returning to long-term investments 

201. Smolka, et al., “Flight Testing of the Gulfstream Quiet Spike,” pp. 28–38; Haering, et al., 

“Preliminary Results from the Quiet Spike Flight Test,” briefing presented at the Fundamental Aeronautics Program meeting, New Orleans, Oct. 30–Nov. 1, 2007. 

202. “Aeronautics/Propulsion Laureate,” Aviation Week, Mar. 17, 2008, p. 40; “Obituary: 

Gerard Schkolnik,” ibid., Apr. 21, 2008, p. 22.
NASA’s Contributions to Aeronautics

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in cutting-edge fundamental research in traditional aeronautical 

disciplines . . . appropriate to NASA’s unique capabilities.” One of the 

four new program areas announced was Fundamental Aeronautics 

(which included supersonic research), with Rich Wlezien as acting 

director.203

During May, NASA released more details on Fundamental 

Aeronautics, including plans for what was called the Supersonics 

Project, managed by Mary Jo Long-Davis with Peter Coen as its principal investigator. One of the project’s major technical challenges was to 

accurately model the propagation of sonic booms from aircraft to the 

ground incorporating all relevant physical phenomena. These included 

realistic atmospheric conditions and the effects of vibrations on structures and the people inside (for which most existing research involved 

military firing ranges and explosives). “The research goal is to model 

sonic boom impact as perceived both indoors and outdoors.”Developing 

the propagation models would involve exploitation of existing databases 

and additional flight tests as necessary to validate the effects of 

molecular relaxation, rise time, and turbulence on the loudness 

of sonic booms.204

As the Supersonics Project evolved, it added aircraft concepts more 

challenging than an SSBJ to serve as longer-range targets on which to 

focus advanced research and technologies. These were a medium-sized 

(100–200 passenger) Mach 1.6–1.8 supersonic airliner that could have 

an acceptable sonic boom by about 2020 and an efficient multi-Mach 

aircraft that might have an acceptably low boom when flying at a 

speed somewhat below Mach 2 by the years 2030–2035. NASA 

awarded advanced concept studies for these in October 2008.205 NASA 

203. Michael Braukus/Doc Mirleson, “NASA Restructures Aeronautics Research,” NASA News Release 

06-008, Jan. 12, 2006; Lisa Porter, “Reshaping NASA’s Aeronautics Program, Briefing,” Jan. 12, 

2006, http://www.nasa.gov/home/hqnews/2006/jan/ HQ_06008_ARMD_ Restructuring.html.

204. Peter Coen, Mary Jo Long-Davis, and Louis Povinelli, “Fundamental Aeronautics Program 

Supersonics Project, Reference Document,” May 26, 2006, pp. 36–37, http://www.aeronautics.

nasa.gov/fap/documents.html (quote from p. 36). 

205. Jefferson Morris, “Quiet, Please: With More Emphasis on Partnering, NASA Continues Pursuit of 

Quieter Aircraft,” Aviation Week, June 25, 2007, p. 57; Lisa J. Porter, “NASA’s Aeronautics Program,” 

Fundamental Aeronautics Annual Meeting, New Orleans, Oct. 30, 2007, slide on “Supersonics 

System Level Metrics,” http://www.aeronautics.nasa.gov/fap/PowerPoints/ARMD&FA_Intro.pdf; 

Beth Hickey, “NASA Awards Future Aircraft Research Contracts,” Contract Release C08-060, Oct. 6, 

2008; Graham Warwick, “Forward Pitch,” Aviation Week, Oct. 20, 2008, p. 22.
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also began working with Japan’s Aerospace Exploration Agency (JAXA) 

on supersonic research, including sonic boom modeling.206 Although 

NASA was not ready as yet to develop a new low-boom supersonic 

research airplane, it supported an application by Gulfstream to the Air 

Force that reserved the designation X-54A just in case this would be 

done in the future.207

Meanwhile, existing aircraft had continued to prove their value 

for sonic boom research. During 2005, the Dryden Center began 

applying a creative new flight technique called low-boom/no-boom to 

produce controlled booms. Ed Haering used PCBoom4 modeling in 

developing this concept, which Jim Smolka then refined into a flyable maneuver with flight tests over an extensive array of pressure 

sensors and microphones. The new technique allowed F-18s to generate shaped (“low boom”) signatures as well as the evanescent sound 

waves (“no-boom”) that remain after the refraction and absorption of 

shock waves generated allow Mach speeds (known as the Mach cutoff) 

before they reach the surface. 

The basic low-boom/no-boom technique requires cruising just 

below Mach 1 at about 50,000 feet, rolling into an inverted position, 

diving at a 53-degree angle, keeping the aircraft’s speed at Mach 1.1 

during a portion of the dive, and pulling out to recover at about 

32,000 feet. This flight profile took advantage of four attributes 

that contribute to reduced overpressures: a long propagation distance 

(the relatively high altitude of the dive), the weaker shock waves 

generated from the top of an aircraft (by diving while upside down), 

low airframe weight and volume (the relatively small size of an F-18), and a 

low Mach number. This technique allowed Dryden’s F-18s, which normally 

generate overpressures of 1.5 psf in level flight, to produce overpressures 

under 0.1 psf. Using these maneuvers, Dryden’s skilled test pilots could 

precisely place these focused quiet booms on specific locations, such as 

those with observers and sensors. Not only were the overpressures 

low, they had a slower rise time than the typical N-shaped sonic 

206. Hans Greimel, “Japan to Talk with NASA on Supersonic Jet,” Washington Post, May 8, 2006, 

http://www.washingtonpost.com/wp-dyn/content/article/2006/05/09/AR2226050800267; 

Beth Dickey, “NASA and JAXA to Conduct Joint Research on Sonic Boom Modeling,” NASA News 

Release 09-117, May 8, 2008. 

207. “X-54A Designation Issued as Placeholder for Future Boom Research Aircraft,” Aerospace 

Daily, July 21, 2008, p. 1.
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NASA F-15B No. 836 in flight with Quiet Spike, September 2006. NASA.

signature. The technique also resulted in systematic recordings of 

evanescent waves—the kind that sound merely like distant thunder.208

Dryden researchers used this technique in July 2007 during a test 

called House Variable Intensity Boom Effect on Structures (House 

VIBES). Following up on a similar test from the year before with an 

old (early 1960s) Edwards AFB house slated for demolition,209 Langley 

engineers installed 112 sensors (a mix of accelerometers and microphones) inside the unoccupied half of a modern (late 1990s) duplex 

house. Other sensors were placed outside the house and on a nearby 

35-foot tower. These measured pressures and vibrations from 12 normal intensity N-shaped booms (up to 2.2 psf) created by F-18s in steady 

and level flight at Mach 1.25 and 32,000 feet as well as 31 shaped booms 

(registering only 0.1 to 0.7 psf) from F-18s using the Low Boom/No 

Boom flight profile. The latter booms were similar to those that would 

208. Haering, Smolka, James E. Murray, and Plotkin, “Flight Demonstration of Low Overpressure NWave Sonic Booms and Evanescent Waves,” Innovations in Non-Linear Acoustics: 17th International 

Symposium on Nonlinear Acoustics, International Sonic Boom Forum, State College, PA, July 21–22, 

2005, American Institute of Physics Conference Proceedings, vol. 838 (May 2006), pp. 647–650. 

209. Jacob Klos and R.D. Bruel, “Vibro-Acoustical Response of Buildings Due to Sonic Boom Exposure: June 2006 Field Test,” NASA TM-2007-214900, Sept. 2007.
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be expected from an acceptable supersonic business jet. The specially 

instrumented F-15B No. 852 performed six flights, and an F-18A did 

one flight. Above the surface boundary layer, an instrumented L-23 

sailplane from the Air Force Test Pilot School recorded shock waves 

at precise locations in the path of the focused booms to account for 

atmospheric effects. The data from the house sensors confirmed fewer 

vibrations and noise levels in the modern house than had been the case 

with the older house. At the same time, data gathered by the outdoor 

sensors added greatly to NASA’s variable intensity sonic boom database, which was expected to help program and validate sonic boom 

propagation codes for years to come, including more advanced threedimensional versions of PCBoom.210

With the awakening of interest in an SSBJ, NASA Langley acoustics 

specialists including Brenda Sullivan and Kevin Shepherd had resumed 

an active program of studies and experiments on human and structural 

response to sonic booms. They upgraded the HSR-era simulator booth 

with an improved computer-controlled playback system, new loudspeakers, and other equipment to more accurately replicate the sound 

of various boom signatures, such as those recorded at Edwards. In 2005, 

they also added predicted boom shapes from several low-boom aircraft 

designs.211 At the same time, Gulfstream created a new mobile sonic 

boom simulator to help demonstrate the difference between traditional 

and shaped sonic booms to a wider audience. Although Gulfstream’s 

folded horn design could not reproduce the very low frequencies of 

Langley’s simulator booth, it created a “traveling” pressure wave that 

moved past the listener and resonated with postboom noises, features 

that were judged more realistic than other simulators. 

Under the aegis of the Supersonics Project, plans for additional simulation capabilities accelerated. Based on multiple studies that had long 

cited the more bothersome effects of booms experienced indoors, the 

210. Creech, “Sonic Boom Tests Scheduled,” Dryden News Release 07-38, July 5, 2007; Guy 

Norris, “Sonic Spike,” Aviation Week, Oct. 8, 2007, p. 52; Haering, et al., “Initial Results from the 

Variable Intensity Sonic Boom Propagation Database,” AIAA Paper 2008-3034, presented at the 

14th AIAA/CEAS Aeroacoustics Conference, Vancouver, BC, Canada, May 5–7, 2008; Jacob 

Klos, “Vibro-Acoustic Response of Buildings Due to Sonic Boom Exposure: July 2007 Field Test,” 

NASA TM-2008-215349, Sept. 2008.

211. Brenda M. Sullivan, “Research on Subjective Response to Simulated Sonic Booms at NASA 

Langley Research Center,” paper presented at International Sonic Boom Forum, State College, PA, 

July 21–22, 2005.
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Langley Center began in the summer of 2008 to build one of the most 

sophisticated sonic boom simulation systems yet. Scheduled for completion in early 2009, it would consist of a carefully constructed 12- by 

14-foot room with sound and pressure systems that would replicate all 

the noises and vibrations caused by various levels and types of sonic 

booms.212 Such studies would be vital if most concepts for supersonic 

business jets were ever to be realized. When the FAA updated its policy 

on supersonic noise certification in October 2008, it acknowledged the 

promising results of recent experiments but cautioned that any future 

changes in the rules against supersonic flight would still depend on 

public acceptance.213

NASA’s Supersonics Project also put a new flight test on its agenda: 

the Lift and Nozzle Change Effects on Tail Shocks (LaNCETS). Both the 

SSBD and Quiet Spike experiments had only involved shock waves from 

the front of an aircraft. Yet shocks from the rear of an aircraft as well as 

jet engine exhaust plumes also contribute to sonic booms—especially 

the recompression phase of the typical N-wave signature—but have long 

been more difficult to control. NASA initiated the LaNCETS experiment 

to address this issue. As described in the Supersonic Project’s original 

planning document, one of the metrics for LaNCETS was to “investigate 

control of aft shock structure using nozzle and/or lift tailoring with the 

goal of a 20% reduction in near-field tail shock strength.”214

NASA Dryden had just the airplane with which to do this: F-15B 

No. 837. Originally built in 1973 as the Air Force’s first preproduction 

TF-15A two-seat trainer (soon redesignated as the F-15B), it had been 

extensively modified for various experiments over its long lifespan. 

These included the Short Takeoff and Landing Maneuvering Technology 

Demonstration, the High-Stability Engine Control project, the Advanced 

Control Technology for Integrated Vehicles Experiment (ACTIVE), 

212. Brenda M. Sullivan, “Design of an Indoor Sonic Boom Simulator at NASA Langley Research 

Center,” July 28, 2008, paper presented at Noise-Con 2008, Baltimore, July 12–14, 2008, and 

“Research at NASA on Human Response to Sonic Booms,” Nov. 17, 2008, at 5th International 

Conference on Flow Dynamics, Sendai, Japan, Nov. 17–20, 2008; Coen, Lou Povinelli, and Kaz 

Civinskas, “Supersonics Project Overview,” Fundamental Aeronautics Annual Meeting, Atlanta, Oct. 

7, 2008, http://www.aeronautics.nasa.gov/fap/ PowerPoints/SUP_ATL_Overview.pdf.

213. “FAA Updates Policy on SST Noise Certification,” The Weekly of Business Aviation, Oct. 27, 

2008, p. 195.

214. “Supersonics Project Reference Document,” p. 43. The acronym “LaNCETS” was devised by 

Haering: Discussion with Benson, at the Dryden Flight Research Center, Dec. 12, 2008.
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and Intelligent Flight Control Systems (IFCS). F-15B No. 837 had the 

following special features: digital fly-by-wire controls, canards ahead of 

the wings for changing longitudinal lift distribution, and thrust-vectoring 

variable area ratio nozzles on its twin jet engines that could (1) constrict 

and expand to change the shape the exhaust plumes and (2) change the 

pitch and yaw of the exhaust flow.215 It was planned to use these capabilities for validating computational tools developed at Langley, Ames, 

and Dryden to predict the interactions between shocks from the tail and 

exhaust under various lift and plume conditions. 

Tim Moes, one of the Supersonics Project’s associate managers, was 

the LaNCETS project manager at the Dryden Center. Jim Smolka, who 

had flown most of F-15B No. 837’s previous missions at Dryden, was 

its test pilot. He and Nils Larson in F-15B No. 836 conducted Phase I 

of the test program with three missions from June 17–19, 2008. They 

gathered baseline measurements with 29 probes, all at 40,000 feet and 

speeds of Mach 1.2, 1.4, and 1.6.216

Several months before Phase II of LaNCETS, NASA specialists and 

affiliated researchers in the Supersonics Project announced significant 

progress in near-field simulation tools using the latest in computational 

fluid dynamics. They even reported having success as far out as 10 body 

lengths (a mid-field distance). As seven of these researchers claimed 

in August 2008, “[It] is reasonable to expect the expeditious development of an efficient sonic boom prediction methodology that will eventually become compatible with an optimization environment.”217 Of 

course, more data from flight-testing would increase the likelihood of 

this prediction. 

LaNCETS Phase II began on November 24, 2008, with nine missions flown by December 11. After being interrupted by a freak snowstorm during the third week of December and then having to break for 

the holiday season, the LaNCETS team completed the project with flight 

215. Dryden Flight Research Center, “F-15B #837,” http://www.nasa.gov/centers/dryden/ 

aircraft/F-15B-837/index.html, accessed Feb. 11, 2009.

216. Larry Cliatt, et al., “Overview of the LaNCETS Flight Experiment and CFD Analysis,” Briefing, 

Fundamental Aeronautics Annual Meeting, Atlanta, Oct. 2, 2008.

217. J.H. Casper, et al., “Assessment of Near-Field Sonic Boom Simulation Tools,” AIAA Paper 

2008-6592 (p. 8 quoted) and Richard L. Campbell, et al., “Efficient Unstructured Grid Adaptation 

Methods for Sonic Boom Prediction,” AIAA Paper 2008-7327, both presented at 26th Applied 

Aerodynamics Conference, Honolulu, Aug. 18–21, 2008. 
NASA’s Contributions to Aeronautics

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tests on January 12, 15, and 30, 2009. In all, Jim Smolka flew 13 missions in F-15B No. 837, 11 of which included in-flight shock wave measurements by No. 836 from distances of 100 to 500 feet. Nils Larson 

piloted the probing flights, with Jason Cudnik or Carrie Rhoades in the 

back seat. The aircrews tested the effects of both positive and negative 

canard trim at Mach 1.2, 1.4, and 1.6 as well as thrust vectoring at Mach 

1.2 and 1.4. They also gathered supersonic data on plume effects with 

different nozzle areas and exit pressure ratios. Once again, GPS equipment recorded the exact locations of the two aircraft for each of the 

datasets. On January 30, 2009, with Jim Smolka at the controls for the 

last time, No. 837 made a final flight before its well-earned retirement.218

The large amount of data collected will be made available to industry and academia, in addition to NASA researchers at Langley, Ames, 

and Dryden. For the first time, analysts and engineers will be able to 

use actual flight test results to validate and improve CFD models on tail 

shocks and exhaust plumes—taking another step toward the design of 

a truly low-boom supersonic airplane.219

Perspectives on the Past, Prospects for the Future

Unfortunately for the immediate future of civilian supersonic flight, the 

successful LaNCETS project coincided almost exactly with the spread 

of the global financial crisis and the start of a severe recession. These 

negative economic developments hit almost all major industries, not 

the least being air carriers and aircraft manufacturers. The impact on 

those recently thriving companies making business jets was aggravated 

even more by populist and political backlash at executives of troubled 

corporations, some now being subsidized by the Federal Government, 

for continuing to fly in corporate jets. Lamenting this unsought negative publicity, Aviation Week and Space Technology examined the plight 

218. Norris, “Sonic Solutions: NASA Uses Unique F-15B to Complete Design Tools for Quiet 

Supersonic Aircraft,” Aviation Week, Jan. 5, 2009, p. 53; Creech and Beth Dickey, “Lancets Flights 

Probe Supersonic Shockwaves,” Dryden News Release 09-04, Jan. 22, 2009; Tim Moes, “Sonic 

Boom Research at NASA Dryden: Objectives and Flight Results of the Lift and Nozzle Change 

Effects on Tail Shock (LaNCETS) Project,” Partial Briefing Slides, International Test & Evaluation 

Association, Antelope Valley Chapter, Feb. 24, 2009; E-mail, Tim Moes to Benson, “Re: More 

Details on LaNCETS,” Mar. 11, 2009.

219. For an early analysis, see Trong T. Bui, “CFD Analysis of Nozzle Jet Plume Effects on Sonic 

Boom Signature,” AIAA Paper 2009-1054, presented at 47th Aerospace Sciences Meeting, 

Orlando, FL, Jan. 5–8, 2009.
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of the small-jet manufacturers in a story with following subheading: “As 

if the economy were not enough, business aviation becomes a scapegoat for executive excess.”220 Nevertheless, NASA was continuing to 

invest in supersonic technologies and sonic boom research, and the aircraft industry was not ready to abandon the ultimate goal of supersonic 

civilian flight. For example, Boeing—under a Supersonics Project contract—was studying low-boom modifications for one of NASA’s F-16XL 

aircraft as one way to seek the holy grail for practical supersonic commercial flight: acceptance by the public. This relatively low-cost idea for 

a shaped sonic boom demonstrator had been one of the options being 

considered during NASA’s short-lived Sonic Boom Mitigation Project in 

2005. Since then, findings from the Quiet Spike and LaNCETS experiments, along with continued progress in computational fluid dynamics, were helping to confirm and refine the aerodynamic and propulsion 

attributes needed to mitigate the strength of sonic booms. 

In the case of the F-16XL, the modifications proposed by Boeing 

included an extended nose glove (reminiscent of the SSBD), lateral chines 

that blend into the wings (as with the SR-71), a sharpened V-shaped front 

canopy (like those of the F-106 and SR-71), an expanded nozzle for its jet 

engine (similar to those of F-15B No. 837), and a dorsal extension (called 

a “stinger”) to lengthen the rear of the airplane. Although such add-ons 

would preclude the low-drag characteristics also desired in a demonstrator, Boeing felt that its “initial design studies have been encouraging with 

respect to shock mitigation of the forebody, canopy, inlet, wing leading 

edge, and aft lift/volume distribution features.” Positive results from more 

detailed designs and successful wind tunnel testing would be the next 

requirements for continuing consideration of the proposed modifications.221

It was clear that NASA’s discoveries about sonic booms and how 

to control them were beginning to pay dividends. Whatever the fate of 

Boeing’s idea or any other proposals yet to come, NASA was committed to finding the best way to demonstrate fully shaped sonic booms. As 

another encouraging sign, the FAA was working with NASA on a roadmap 

for studying community reactions to sonic booms, one that would soon 

be presented to the ICAO.222

220. Graham Warwick, et al., “Open Season,” Aviation Week, Mar. 2, 2009, pp. 20–21.

221. Warwick, “Beyond the N-Wave: Modifying NASA’s Arrow-Wing F-16XL Could Help Pave 

the Way for Low-Boom Supersonic Transports,” Aviation Week, Mar. 23, 2009, p. 52.

222. E-mail, Coen, Langley Research Center, to Benson, Apr. 10, 2009.
NASA’s Contributions to Aeronautics

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As shown in this study, past expectations for a quiet civilian supersonic transport had repeatedly run up against scientific, technical, economic, and political hurdles too high to overcome. That is why such an 

airplane has yet to fly. Yet the knowledge gained and lessons learned from 

each attempt attest to the value of persistence in pursuing both basic and 

applied research. Recent progress in shaping sonic booms builds upon the 

work of dedicated NASA civil servants over more than half a century, the 

data and documentation preserved through NASA’s scientific and technical information program, the special facilities and test resources maintained and operated by NASA’s research Centers, and NASA’s support of 

and partnership with contractors and universities. 

Since the dawn of civilization, conquering the twin tyrannies of time 

and distance has been a powerful human aspiration, one that has served 

as a catalyst for many technological innovations. It seems reasonable to 

assume that this need for speed will eventually break down the barriers in 

the way of practical supersonic transportation, to include solving the problem of the sonic boom. When that time finally does come, it will have been 

made possible by NASA’s many years of meticulous research, careful testing, and inventive experimentation on ways to soften the sonic footprint. 
Case 4 | Softening the Sonic Boom: 50 Years of NASA Research

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Recommended Additional Readings

Reports, Papers, Articles, and Presentations: 

William J. Alford and Cornelius Driver, “Recent Supersonic Transport 

Research,” Astronautics & Aeronautics, vol. 2, no. 9 (Sept. 1964), 

pp. 26–37.

William H. Andrews, “Summary of Preliminary Data Derived from the 

XB-70 Airplanes,” NASA TM-X-1240 (1966). 

Daniel G. Baize, “1995 NASA High-Speed Research Program Sonic Boom 

Workshop,” NASA CP-3335, vol. 1 (1996). 

Raymond L. Barger, “Sonic-Boom Wave-Front Shapes and Curvatures 

Associated with Maneuvering Flight,” NASA TP-1611 (1979).

W.D. Beasly, J.D. Brooks, and R.L. Barger, “A Laboratory Investigation 

of N-Wave Focusing,” NASA TN-D-5306 (1969). 

Percy J. Bobbit and Domenic J. Maglieri, “Dr. Antonio Ferri’s Contribution 

to Supersonic Transport Sonic-Boom Technology,” Journal of 

Spacecraft and Rockets, vol. 40, no. 4 (July–Aug. 2003), pp. 459–466.

J.D. Brooks, et al., “Laboratory Investigation of Diffraction and Reflection 

of Sonic Booms by Buildings,” NASA TN-D-5830 (1970).

J.F. Bryant, D.J. Maglieri, and V.S. Richie, “In-Flight Shock-Wave 

Measurements Above and Below a Bomber Airplane at Mach 

Numbers from 1.42 to 1.69,” NASA TN-D-1968 (1963).

Trong T. Bui, “CFD Analysis of Nozzle Jet Plume Effects on Sonic Boom 

Signature,” AIAA Paper 2009-1054 (2009).

Richard L. Campbell, et al., “Efficient Unstructured Grid Adaptation 

Methods for Sonic Boom Prediction,” AIAA Paper 2008-7327 (2008). 

Harry W. Carlson, “An Investigation of Some Aspects of the Sonic Boom 

by Means of Wind-Tunnel Measurements of Pressures about Several 

Bodies at a Mach Number of 2.01,” NASA TN-D-161 (1959).
NASA’s Contributions to Aeronautics

264

4

Harry W. Carlson, “Configuration Effects on Sonic Boom,” Proceedings 

of NASA Conference on Supersonic-Transport Feasibility Studies and 

Supporting Research, Sept. 17–19, 1963 . . . Hampton, VA, NASA 

TM-X-905 (1963).

Harry W. Carlson, “Correlation of Sonic-Boom Theory with Wind Tunnel 

and Flight Measurements,” NASA TR-R-213 (1964).

Harry W. Carlson, “Some Notes on the Present Status of Sonic Boom 

Prediction and Minimization Research,” Third Conference on Sonic 

Boom Research . . . Washington, DC, Oct. 29–30, 1970, NASA SP-255 

(1971).

Harry W. Carlson, “Simplified Sonic-Boom Prediction,” NASA TP-1122 

(1978).

Harry W. Carlson, et al., “A Wind Tunnel Study of Sonic-Boom 

Characteristics for Basic and Modified Models of a Supersonic 

Transport Configuration,” NASA TM-X-1236 (1966).

Harry W. Carlson, et al., “Application of Sonic-Boom Minimization 

Concepts in Supersonic Transport Design,” NASA TN-D-7218 (1973).

J.H. Casper, et al., “Assessment of Near-Field Sonic Boom Simulation 

Tools,” AIAA Paper 2008-6592 (2008).

Samsun Cheung, “Supersonic Civil Airplane Study and Design: 

Performance and Sonic Boom,” NASA CR-197745 (1995).

Clark, Buhr, & Nexen [firm]. “Studies of Sonic Boom Damage,” NASA 

CR-227 (1965).

Peter G. Coen, “Development of a Computer Technique for Prediction 

of Transport Aircraft Flight Profile Sonic Boom Signatures,” NASA 

CR-188117 (1991).

Hugo E. Dahlke, et al., “The Shock-Expansion Tube and Its Application 

as a Sonic Boom Simulator,” NASA CR-1055 (1968). 
Case 4 | Softening the Sonic Boom: 50 Years of NASA Research

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Christine M. Darden, “Minimization of Sonic-Boom Parameters in Real 

and Isothermal Atmospheres,” NASA TN-D-7842 (1975). 

Christine M. Darden, “Sonic Boom Theory—Its Status in Prediction and 

Minimization,” AIAA Paper 76-1 (1976). 

Christine M. Darden, “Sonic Boom Minimization with Nose-Bluntness 

Relaxation,” NASA TP-1348 (1979). 

Christine M. Darden, “Charts for Determining Potential Minimum SonicBoom Overpressures for Supersonic Cruise Aircraft,” NASA TP-1820 

(1981).

Christine M. Darden, ed., High Speed Research: Sonic Boom; Proceedings 

of a Conference at the Langley Research Center, Feb. 25–27, 1992, 

NASA CR-3172 (1992), vols. 1, 2. 

Christine M. Darden and R.J. Mack, “Some Effects of Applying Sonic 

Boom Minimization to Supersonic Aircraft,” Journal of Aircraft, vol. 

17, no. 3 (Mar. 1980), pp. 182–186.

Christine M. Darden, et al., “Status of Sonic Boom Methodology and 

Understanding,” NASA CP-3027 (1989). 

J.W.M. Dumond, et al., “A Determination of the Wave Forms and Laws of 

Propagation and Dissipation of Ballistic Shock Waves,” Journal of the 

Acoustical Society of America, vol. 18, no. 1 (Jan. 1946), pp. 97–118.

Philip M. Edge and William H. Mayes, “Description of Langley LowFrequency Noise Facility and Study of Human Response to Noise 

Frequencies below 50 cps,” NASA TN-D-3204 (1966).

Phillip M. Edge and Harvey H. Hubbard, “Review of Sonic-Boom 

Simulation Devices and Techniques,” Dec. 1970, Journal of the 

Acoustical Society of America, vol. 51, no. 2, pt. 2 (Feb. 1972), p. 723.

Thomas A. Edwards, ed., High-Speed Research: Sonic Boom; Proceedings 

of a Conference at the Ames Research Center, May 12–14, 1993, NASA 

CP-10132 (1993), vol. 1.
NASA’s Contributions to Aeronautics

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4

Antonio Ferri, Huai-Chu Wang, and Hans Sorensen, “Experimental 

Verification of Low Sonic Boom Configuration,” NASA CR-2070 

(1973).

James M. Fields, “Reactions of Residents to Long-Term Sonic Boom 

Noise Environments,” NASA CR-201704 (1997).

A.B. Fryer, et al., “Publications in Acoustics and Noise Control from the 

NASA Langley Research Center during 1940–1976,” NASA TM-X74042 (1977). 

John H. Gardner and Peter H. Rogers, “Thermospheric Propagation 

of Sonic Booms from the Concorde Supersonic Transport,” Naval 

Research Laboratory Memo Report 3904, Feb. 14, 1979 (DTIC AD 

A067201).

A.R. George and A.R. Seebass, “Sonic-Boom Minimization,” Nov. 1970, 

Journal of the Acoustical Society of America, vol. 51, no. 2, pt. 3 (Feb. 

1972), pp. 686–694.

David H. Graham, et al., “Wind Tunnel Validation of Shaped Sonic Boom 

Demonstration Aircraft Design,” AIAA Paper 2005-7 (2005).

Karen S. Green and Terrill W. Putnam, “Measurements of Sonic Booms 

Generated by an Airplane Flying at Mach 3.5 and 4.8,” NASA TM-X3126 (1974). 

Edward A. Haering, et al., “Airborne Shaped Sonic Boom Demonstration 

Pressure Measurements with Computational Fluid Dynamics 

Comparisons,” AIAA Paper 2005-9 (2005).

Edward A. Haering, James W. Smolka, James E. Murray, and Kenneth J. 

Plotkin, “Flight Demonstration of Low Overpressure N-Wave Sonic 

Booms and Evanescent Waves,” Innovations in Non-Linear Acoustics: 

17th International Symposium on Nonlinear Acoustics, International 

Sonic Boom Forum, State College, PA, July 21–22, 2005, American 

Institute of Physics Conference Proceedings, vol. 838 (May 2006), 

pp. 647–650. 
Case 4 | Softening the Sonic Boom: 50 Years of NASA Research

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Edward A. Haering, et al., “Initial Results from the Variable Intensity 

Sonic Boom Propagation Database,” AIAA Paper 2008-3034 (2008).

George T. Haglund and Edward J. Kane, “Flight Test Measurements 

and Analysis of Sonic Boom Phenomena Near the Shock Wave 

Extremity,” NASA CR-2167 (1973). 

Thomas H. Higgins, “Sonic Boom Research and Design Considerations 

in the Development of a Commercial Supersonic Transport,” Journal 

of the Acoustical Society of America, vol. 39, no. 5, pt. 2 (Nov. 1966), 

pp. 526–531.

David. A. Hilton, Vera Huckel, and Domenic J. Maglieri, “Sonic Boom 

Measurements during Bomber Training Operations in the Chicago 

Area,” NASA TN-D-3655 (1966).

Harvey H. Hubbard, et al., “Ground Measurements of Sonic-Boom 

Measurements for the Altitude Range of 10,000 to 75,000 Feet,” 

NASA TR-R-198 (1964). 

D.A. Hilton, D.J. Maglieri, and R. Steiner, “Sonic-Boom Exposures during FAA Community Response Studies over a 6-Month Period in the 

Oklahoma City Area,” NASA TN-D-2539 (1964). 

Sherwood Hoffman, “Bibliography of Supersonic Cruise Aircraft 

Research (SCAR)” [1972–1977], NASA RP-1003 (1977).

Sherwood Hoffman, “Bibliography of Supersonic Cruise Research (SCR) 

Program from 1977 to Mid-1980,” NASA RP-1063 (1980).

Donald C. Howe, et al., “Development of the Gulfstream Quiet Spike for 

Sonic Boom Minimization,” AIAA Paper 2008-124 (2008).

Harvey H. Hubbard and Domenic J. Maglieri, “Sonic Boom Signature 

Data from Cruciform Microphone Array Experiments during the 

1966–67 EAFB National Sonic Boom Evaluation Program,” NASA 

TN-D-6823 (1972).
NASA’s Contributions to Aeronautics

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4

Harvey H. Hubbard, Domenic J. Maglieri, and David G. Stephens, “SonicBoom Research—Selected Bibliography with Annotation,” NASA 

TM-87685 (1986).

James and Associates, ed., YF-12 Experiments Symposium: A conference 

held at Dryden Flight Research Center . . . Sept. 13–15, 1978, NASA 

CP-2054 (1978).

Gareth H. Jordan, “Flight Measurements of Sonic Booms and Effects 

of Shock Waves on Aircraft,” in Society of Experimental Test Pilots 

Quarterly Review, vol. 5, no. 1 (1961), pp. 117–131.

Jacob Klos and R.D. Bruel, “Vibro-Acoustical Response of Buildings Due 

to Sonic Boom Exposure: June 2006 Field Test,” NASA TM-2007-

214900 (2007).

Jacob Klos, “Vibro-Acoustic Response of Buildings Due to Sonic Boom 

Exposure: July 2007 Field Test,” NASA TM-2008-215349 (2008).

Christopher A. Lee, “Design and Testing of Low Sonic Boom Configurations 

and an Oblique All-Wing Supersonic Transport,” NASA CR-197744 

(1995).

G.M. Lilley, et al., “Some Aspects of Noise from Supersonic Aircraft,” 

Journal of the Royal Aeronautical Society, vol. 57 (June 1953), pp. 

396–414.

Lindsay J. Lina and Domenic J. Maglieri, “Ground Measurements of 

Airplane Shock-Wave Noise at Mach Numbers to 2.0 and at Altitudes 

to 60,000 Feet,” NASA TN-D-235, Mar. 1960.

Jerome Lukas and Karl D. Kryler, “A Preliminary Study of the Awakening 

and Startle Effects of Simulated Sonic Booms,” NASA CR-1193 

(1968). 

Jerome Lukas and K.D. Kryler, “Awakening Effects of Simulated Sonic 

Booms and Subsonic Aircraft Noise,” NASA CR-1599 (1970).
Case 4 | Softening the Sonic Boom: 50 Years of NASA Research

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B.K.O. Lundberg, “Aviation Safety and the SST,” Astronautics & 

Aeronautics, vol. 3, no. 1 (Jan. 1966), p. 28. 

G.J. MacDonald, et al., “Jason 1978 Sonic Boom Report,” JSR-78-09 

(Arlington, VA: SRI International, Nov. 1978).

Robert J. Mack, “A Supersonic Business Jet Concept Designed for Low 

Sonic Boom,” NASA TM-2003-212435 (2003).

Robert J. Mack and Christine M. Darden, “Wind-Tunnel Investigation of 

the Validity of a Sonic-Boom-Minimization Concept,” NASA TP-1421 

(1979). 

Robert J. Mack and Christine M. Darden, “Some Effects of Applying 

Sonic Boom Minimization to Supersonic Cruise Aircraft Design,” 

Journal of Aircraft, vol. 17, no. 3 (Mar. 1980), pp. 182–186.

Domenic J. Maglieri, “A Brief Review of the National Aero-Space Plane 

Sonic Booms Final Report,” USAF Aeronautical Systems Center 

Report TR-94-9344 (1992).

Domenic J. Maglieri, Vera Huckel, and Tony L. Parrott, “Ground 

Measurements of Shock-Wave Pressure for Fighter Airplanes Flying 

at Very Low Altitude,” NASA TN-D-3443 (1966) (superseded classified TM-X-611, 1961).

Domenic J. Maglieri, Harvey H. Hubbard, and Donald L. Lansing, 

“Ground Measurements of the Shock-Wave Noise from Airplanes 

in Level Flight at Mach Numbers to 1.4 and Altitudes to 45,000 Feet,” 

NASA TN-D-48 (1959).

Domenic J. Maglieri and Harvey H. Hubbard, “Ground Measurements 

of the Shock-Wave Noise from Supersonic Bomber Airplanes in the 

Altitude Range from 30,000 to 50,000 Feet,” NASA TN-D-880 (1961).

Domenic J. Maglieri and Donald L. Lansing, “Sonic Booms from Aircraft 

in Maneuvers,” NASA TN-D-2370 (1964). 
NASA’s Contributions to Aeronautics

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Domenic J. Maglieri and Garland J. Morris, “Measurement of Response 

of Two Light Airplanes to Sonic Booms,” NASA TN-D-1941 (1963).

Domenic J. Maglieri, et al., “Sonic Boom Measurements for SR-71 

Aircraft Operating at Mach Numbers to 3.0 and Altitudes to 24834 

Meters,” NASA TN-D-6823 (1972).

Domenic J. Maglieri, Victor E. Sothcroft, and John Hicks, “Influence 

of Vehicle Configurations and Flight Profile on X-30 Sonic Booms,” 

AIAA Paper 90-5224 (1990).

Domenic J. Maglieri, et al., “A Summary of XB-70 Sonic Boom Signature 

Data, Final Report,” NASA CR-189630 (1992). 

Domenic J. Maglieri, et al., “Feasibility Study on Conducting Overflight 

Measurements of Shaped Sonic Boom Signatures Using the Firebee 

BQM-34E RPV,” NASA CR-189715 (1993).

David A. McCurdy, ed., “High-Speed Research: 1994 Sonic Boom 

Workshop, Atmospheric Propagation and Acceptability Studies,” 

NASA CP-3209, and “ Configuration, Design, Analysis, and Testing,” 

NASA CP-209669 (1999).

F. Edward McLean, “Some Nonasymptotic Effects of the Sonic Boom 

of Large Airplanes,” NASA TN-D-2877 (1965).

Keith H. Meredith, et al., “Computational Fluid Dynamics Comparison 

and Flight Test Measurement of F-5E Off-Body Pressures,” AIAA 

Paper 2005-6 (2005).

David S. Miller and Harry W. Carlson, “Sonic Boom Minimization by 

Application of Heat or Force Fields to Airplane Airflow,” NASA TN-D5582 (1969).

John M. Morgenstern, et al., “F-5 Shaped Sonic Boom Demonstrator’s 

Persistence of Boom Shaping Reduction through Turbulence,” AIAA 

2005-12 (2005).
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Marshall E. Mullens, “A Flight Test Investigation of the Sonic Boom,” 

AFFTC TN-56-20 (1956). 

Charles W. Nixon and Harvey H. Hubbard, “Results of the USAF–NASA–

FAA Flight Program to Study Community Response to Sonic Booms 

in the Greater St. Louis Area,” NASA TN-D-2705 (1965). 

Joseph W. Pawlowski, David H. Graham, Charles H. Boccadoro (NGC), 

Peter G. Coen (LaRC), and Domenic J. Maglieri (Eagle Aero.), 

“Origins and Overview of the Shaped Sonic Boom Demonstration 

Program,” AIAA Paper 2005-5 (2005), pp. 3–7. 

Kenneth J. Plotkin, “State of the Art of Sonic Boom Modeling,” Journal 

of the Acoustical Society of America, vol. 111, No. 1, pt. 3 (Jan. 2002), 

pp. 530–536.

Kenneth J. Plotkin and Domenic J. Maglieri, “Sonic Boom Research: 

History and Future,” AIAA Paper 2003-3575 (2003).

Kenneth J. Plotkin, et al., “Ground Data Collection of Shaped Sonic 

Boom Experiment Aircraft Pressure Signatures,” AIAA Paper 2005-

10 (2005).

Hugh W. Poling, “Sonic Boom Propagation Codes Validated by Flight 

Test,” NASA CR-201634 (1996).

John O. Powers, J.M. Sands, and Domenic Maglieri, “Survey of United States 

Sonic Boom Overflight Experimentation,” NASA TM-X-66339 (1969).

Johnny M. Sands, “Sonic Boom Research (1958-1968),” FAA, Nov. 1968, 

Defense Technical Information Center (DTIC) document AD-684806. 

Ira R. Schwartz, ed., Sonic Boom Research, Second Conference, 

Washington, DC, May 9–10, 1968, NASA SP-180 (1968). 

Ira R. Schwartz, ed., Third Conference on Sonic Boom Research. . . 

Washington, DC, Oct. 29–30, 1970, NASA SP-255 (1971).
NASA’s Contributions to Aeronautics

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4

A.R. Seebass, ed., Sonic Boom Research: Proceedings of a Conference. . . 

Washington, DC, Apr. 12, 1967, NASA SP-147 (1967).

A.R. Seebass and A.R. George, “Sonic Boom Minimization through 

Aircraft Design and Operation,” AIAA Paper 73-241 (1973). 

A.R. Seebass and A.R. George, “The Design and Operation of Aircraft 

to Minimize Their Sonic Boom,” Journal of Aircraft, vol. 11, no. 9 

(Sept. 1974), pp. 509–517. 

Harriet J. Smith, “Experimental and Calculated Flow Fields Produced 

by Airplanes Flying at Supersonic Speeds,” NASA TN-D-621 (1960).

James W. Smolka, et al., “Flight Testing of the Gulfstream Quiet Spike 

on a NASA F-15B,” in The Society of Experimental Test Pilots, 2007 

Report to the Aerospace Profession (Lancaster, CA: SETP, 2007).

R.T. Sturgielski, et al., “The Development of a Sonic Boom Simulator 

with Detonable Gases,” NASA CR-1844 (1971).

Rudolph J. Swigart, “An Experimental Study in the Validity of the HeatField Concept for Sonic Boom Alleviation,” NASA CR-2381 (1974).

J.P. Taylor and E.R. Taylor, “A Brief Legal History of the Sonic Boom 

in America,” Aircraft Engine Noise and Sonic Boom (Neuilly Sur 

Seine, France: NATO Advisory Group for Aerospace Research and 

Development [AGARD], 1969), Conference Proceedings (CP) No. 

42, Paris, May 1969, pp. 2-1–2-11.

Charles L. Thomas, “Extrapolation of Sonic Boom Pressure Signatures 

by the Waveform Parameter Method,” NASA TN-D-6823 (1972).

Roger Tomboulian, “Research and Development of a Sonic Boom 

Simulation Device,” NASA CR-1378 (1969).

F. Walkden, “The Shock Pattern of a Wing-Body Combination Far from the 

Flight Path,” Aeronautical Quarterly, vol. 9, pt. 2 (May 1958), pp. 164–194.
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G.B. Whitham, “The Flow Pattern of a Supersonic Projectile,” 

Communications on Pure and Applied Mathematics, vol. 5, No. 3 

(1952), pp. 301–348.

G.B. Whitham, “On the Propagation of Weak Shock Waves,” Journal of 

Fluid Dynamics, vol. 1, no. 3 (Sept. 1956), pp. 290–318.

Books and Monographs:

Donald D. Baals and William R. Corliss, Wind Tunnels of NASA, SP-440 

(Washington, DC: NASA, 1981).

Joseph R. Chambers, Innovation in Flight; Research of the NASA Langley 

Research Center on Revolutionary Concepts for Aeronautics, NASA 

SP-2005-4539 (Washington, DC: NASA, 2005).

Erik M. Conway, High Speed Dreams: NASA and the Technopolitics of 

Supersonic Transportation, 1945–1999 (Baltimore: Johns Hopkins, 2005).

Richard P. Hallion, Supersonic Flight: Breaking the Sound Barrier and 

Beyond—The Story of the Bell X-1 and Douglas D-558 (New York: The 

Macmillan Co., 1977).

Richard P. Hallion and Michael H. Gorn, On the Frontier: Experimental 

Flight at NASA Dryden (Washington, DC: Smithsonian, 2003).

Mel Horwitch, Clipped Wings: The American SST Conflict (Cambridge: 

MIT, 1982).

Richard J. Kent, Jr., Safe, Separated, and Soaring: A History of Civil 

Aviation Policy, 1961–1972 (Washington, DC: FAA, 1980).

F. Edward McLean, Supersonic Cruise Technology, NASA SP-472 

(Washington, DC: NASA, 1985).

Peter W. Merlin, From Archangel to Senior Crown: Design and Development 

of the Blackbird (Reston, VA: AIAA, 2008).
NASA’s Contributions to Aeronautics

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4

Kenneth Owen, Concorde: Story of a Supersonic Pioneer (London: Science 

Museum, 2001).

Stuart I. Rochester, Takeoff at Mid-Century: Federal Civil Aviation Policy 

in the Eisenhower Years, 1953–1961 (Washington, DC: FAA, 1976).
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X-15 research pilot (and, subsequently, Gemini and Apollo astronaut) Neil A. Armstrong, 

wearing the X-15’s Clark MC-2 full-pressure suit, 1960. NASA.
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Toward Transatmospheric 

Flight: From V-2 to the X-51

T.A. Heppenheimer

CASE

5

The expansion of high-speed aerothermodynamic knowledge enabled 

the attainment of hypersonic speeds, that is, flight at speeds of Mach 5 

and above. Blending the challenge of space flight and flight within the 

atmosphere, this led to the emergence of the field of transatmospherics: 

systems that would operated in the upper atmosphere, transitioning from 

lifting flight to ballistic flight, and back again. NACA–NASA research 

proved essential to mastery of this field, from the earliest days of blunt 

body reentry theory to the advent of increasingly sophisticated transatmospheric concepts, such as the X-15, the Shuttle, the X-43A, and the X-51.

O N DECEMBER 7, 1995, the entry probe of the Galileo spacecraft 

plunged downward into the atmosphere of Jupiter. It sliced into 

the planet’s hydrogen-rich envelope at a gentle angle and entered 

at Mach 50, with its speed of 29.5 miles per second being four times that of 

a return to Earth from the Moon. The deceleration peaked at 228 g’s, equivalent to slamming from 5,000 mph to a standstill in a single second. Yet 

the probe survived. It deployed a parachute and transmitted data from its 

onboard instruments for nearly an hour, until overwhelmed by the increasing pressures it encountered within the depths of the Jovian atmosphere.1

The Galileo probe offered dramatic proof of how well the National 

Aeronautics and Space Administration (NASA) had mastered the field of 

hypersonics, particularly the aerothermodynamic challenges of doubledigit high-Mach atmospheric entries. That level of performance was 

impressive, a performance foreshadowed by the equally impressive (certainly for their time) earlier programs such as Mercury, Gemini, Apollo, 

Pioneer, and Viking. But NASA had, arguably, an even greater challenge 

before it: developing the technology of transatmospheric flight—the ability to transit, routinely, from flight within the atmosphere to flight out 

1. Richard E. Young, Martha A. Smith, and Charles K. Sobeck, “Galileo Probe: In Situ Observations of Jupiter’s Atmosphere,” Science, no. 272 (May 10, 1996), pp. 837–838.
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into space, and to return again. It was a field where challenge and contradiction readily mixed: a world of missiles, aircraft, spacecraft, rockets, ramjets, and combinations of all of these, some crewed by human 

operators, some not.

Transatmospheric flight requires mastery of hypersonics, flight at 

speeds of Mach 5 and higher in which aerodynamic heating predominates over other concerns. Since its inception after the Second World 

War, three problems have largely driven its development.

First, the advent of the nuclear-armed intercontinental ballistic missile (ICBM), during the 1950s, brought the science of reentry physics 

and took the problem of thermal protection to the forefront. Missile 

nose cones had to be protected against the enormous heat of their atmosphere entry. This challenge was resolved by 1960.

Associated derivative problems were dealt with as well, including 

that of protecting astronauts during demanding entries from the Moon. 

Maneuvering hypersonic entry became a practical reality with the Martin 

SV-5D Precision Recovery Including Maneuvering Entry (PRIME) in 

1967. In 1981, the Space Shuttle introduced reusable thermal protection—the “tiles”—that enabled its design as a “cool” aluminum airplane rather than one with an exotic hot structure. Then in 1995, the 

Galileo mission met demands considerably greater than those of a return 

from the Moon.

A second and contemporary problem, during the 1950s, involved 

the expectation that flight speeds would increase essentially without 

limit. This hope lay behind the unpiloted air-launched Lockheed X-7, 

which used a ramjet engine and ultimately reached Mach 4.31. There 

also was the rocket-powered and air-launched North American X-15, 

the first transatmospheric aircraft. One X-15 achieved Mach 6.70 (4,520 

mph) in October 1967. This set a record for winged hypersonic flight 

that stood until the flight of the Space Shuttle Columbia in 1981. The 

X-15 introduced reaction thrusters for aircraft attitude, and they subsequently became standard on spacecraft, beginning with Project Mercury. 

But the X-15 also used a “rolling tail” with elevons (combined elevators and ailerons) in the atmosphere and had to transition to and from 

space flight. The flight control system that did this later flew aboard the 

Space Shuttle. The X-15 also brought the first spacesuit that was flexible when pressurized rather than being rigid like an inflated balloon. 

It too became standard. In aviation, the X-15 was first to use a simulator as a basic tool for development, which became a critical instrument 
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for pilot training. Since then, simulators have entered general use and 

today are employed with all aircraft.2

A third problem, emphasized during the era of President Ronald 

Reagan’s Strategic Defense Initiative (SDI) in the 1980s, involved the 

prospect that hypersonic single-stage-to-orbit (SSTO) air-breathing 

vehicles would shortly replace the Shuttle and other multistage rocketboosted systems. This concept depended upon the scramjet, a variant 

of the ramjet engine that sustained a supersonic internal airflow to run 

cool. But while scramjets indeed outperformed conventional ramjets and 

rockets, their immaturity and higher drag made their early application 

as space access systems impossible. The abortive National Aero-Space 

Plane (NASP) program consumed roughly a decade of development 

time. It ballooned enormously in size, weight, complexity, and cost as 

time progressed and still lacked, in the final stages, the ability to reach 

orbit. Yet while NASP faltered, it gave a major boost to computational 

fluid dynamics, which use supercomputers to study airflows in aviation. 

This represents another form of simulation that also is entering general use. NASP also supported the introduction of rapid-solidification 

techniques in metallurgy. These enhance alloys’ temperature resistance, 

resulting in such achievements as the advent of a new type of titanium 

that can withstand 1,500 degrees Fahrenheit (°F).3 Out of it have come 

more practical and achievable concepts, as evidenced by the NASA X-43 

program and the multiparty X-51A program of the present.

Applications of practical hypersonics to the present era have been 

almost exclusively within reentry and thermal protection. Military hypersonics, while attracting great interest across a range of mission areas, 

such as surveillance, reconnaissance, and global strike, has remained 

the stuff of warhead and reentry shape research. Ambitious concepts for 

transatmospheric aircraft have received little support outside the laboratory environment. Concepts for global-ranging hypersonic “cruisers” 

withered in the face of the cheaper and more easily achievable rocket.

2. Mark Wolverton, “The Airplane That Flew Into Space,” American Heritage of Invention and 

Technology, (summer 2001), pp. 12–20.

3. J. Sorensen, “Titanium Matrix Composites—NASP Materials and Structures Augmentation 

Program,” AIAA Paper 90-5207 (1990); Stanley W. Kandebo, “Boeing 777 to Incorporate New 

Alloy Developed for NASP,” Aviation Week, May 3, 1993, p. 36; “NASP Materials and Structures 

Program, Titanium Matrix Composites,” McDonnell-Douglas, Dec. 31, 1991, DTIC ADB-192559, 

Defense Technical Information Center.
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Moving Beyond the V-2: John Becker Births American Hypersonics

During the Second World War, Germany held global leadership in highspeed aerodynamics. The most impressive expression of its technical 

interest and competence in high-speed aircraft and missile design was 

the V-2 terror weapon, which introduced the age of the long-range rocket. 

It had a range of over 200 miles at a speed of approximately Mach 5.4 

A longer-range experimental variant tested in 1945, the A-4b, sported 

swept wings and flew at 2,700 mph, reentering and leveling off in the 

upper atmosphere for a supersonic glide to its target. In its one semisuccessful flight, it completed a launch and reentry, though one wing 

broke off during its terminal Mach 4+ glide.5 One appreciates the ambitious nature and technical magnitude of the German achievement given 

that the far wealthier and more technically advantaged United States 

pursued a vigorous program in piloted rocket planes all through the 

1950s without matching the basic performance sought with the A-4b.

Key to the German success was a strong academic-industry partnership and, particularly, a highly advanced complex of supersonic 

wind tunnels. The noted tunnel designer Carl Wieselsberger (who died 

of cancer during the war) introduced a blow-down design that initially 

operated at Mach 3.3 and later reached Mach 4.4. The latter instrument 

supported supersonic aerodynamic and dynamic stability studies of various craft, including the A-4b. German researchers had ambitious plans 

for even more advanced tunnels, including an Alpine complex capable of 

attaining Mach 10. This tunnel work inspired American emulation after 

the war and, in particular, stimulated establishment of the Air Force’s 

Arnold Engineering Development Center at Tullahoma, TN.6

4. Walter Dornberger, V-2 (New York: The Viking Press, 1958 ed.), relates its history from the point 

of view of the German military commander of V-2 development and its principal research facility.

5. Michael J. Neufeld, The Rocket and the Reich: Peenemünde and the Coming of the Ballistic Missile Era (Cambridge: Harvard University Press, 1995), pp. 250–251.

6. Ronald Smelt, “A Critical Review of German Research on High-Speed Airflow,” Journal of the 

Royal Aeronautical Society, vol. 50, No. 432 (Dec. 1946), pp. 899–934; Theodore von Kármán, 

“Where We Stand: First Report to General of the Army H. H. Arnold on Long Range Research Problems of the AIR FORCES with a Review of German Plans and Developments,” Aug. 22, 1945, vol. 

II-1, Copy No. 13, including Hsue-shen Tsien, “Reports on the Recent Aeronautical Developments 

of Several Selected Fields in Germany and Switzerland,” July 1945; Hsue-shen Tsien, “High Speed 

Aerodynamics,” Dec. 1945; and F.L. Wattendorf, “Reports on Selected Topics of German and 

Swiss Aeronautical Developments,” June 1945; Peter P. Wegener, The Peenemünde Wind Tunnels: 

A Memoir (New Haven: Yale University Press, 1996), pp. 22–24, 70.
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The German A-4b, being readied for a test flight, January 1945. USAF.

At war’s end, America had nothing comparable to the investment 

Germany had made in high-speed flight, either in rockets or in wind tunnels and other specialized research facilities. The best American wartime 

tunnel only reached Mach 2.5. As a stopgap, the Navy seized a German 

facility, transported it to the United States, and ran it at Mach 5.18, but 
NASA’s Contributions to Aeronautics

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5

The layout of the Langley 11-inch hypersonic tunnel advocated by John V. Becker. NASA.

it did this only beginning in 1948.7 Even so, aerodynamicist John Becker, 

a young and gifted engineer working at the National Advisory Committee 

for Aeronautics (NACA) Langley Laboratory, took the initiative in introducing Agency research in hypersonics. He used the V-2 as his rationale. 

In an August 1945 memo to Langley’s chief of research, written 3 days 

before the United States atom-bombed Hiroshima, he noted that planned 

NACA facilities were to reach no higher than Mach 3. With the V-2 having 

already flown at Mach 5, he declared, this capability was clearly inadequate.

He outlined an alternative design concept for “a supersonic tunnel 

having a test section four-foot square and a maximum test Mach number 

of 7.0.”8 A preliminary estimate indicated a cost of $350,000. This was no 

mean sum. It was equivalent six decades later to approximately $4.2 million. Becker sweetened his proposal’s appeal by suggesting that Langley 

7. William B. Anspacher, Betty Gay, Donald Marlowe, Paul Morgan, and Samuel Raff, The Legacy 

of the White Oak Laboratory (Dahlgren, VA: Naval Surface Warfare Center, 2000), pp. 209–210; 

Donald D. Baals and William R. Corliss, Wind Tunnels of NASA, SP-440 (Washington, DC: NASA, 

1981), pp. 51–52; James R. Hansen, Engineer in Charge: A History of the Langley Aeronautical 

Laboratory, 1917–1958, SP-4305 (Washington, DC: NASA, 1987), p. 467.

8. Quoted in Hansen, Engineer in Charge, pp. 344–345.
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begin modestly with a small demonstration wind tunnel. It could be built 

for roughly one-tenth of this sum and would operate in the blow-down 

mode, passing flow through a 1-foot-square test section. If it proved successful and useful, a larger tunnel could follow. His reasoned idea received 

approval from the NACA’s Washington office later in 1945, and out of this 

emerged the Langley 11-Inch Hypersonic Tunnel. Slightly later, Alfred J. 

Eggers began designing a hypersonic tunnel at the NACA’s West Coast Ames 

Aeronautical Laboratory, though this tunnel, with a 10-inch by 14-inch test 

section, used continuous, not blow-down, flow. Langley’s was first. When 

the 11-inch tunnel first demonstrated successful operation (to Mach 6.9) 

on November 26, 1947, American aeronautical science entered the hypersonic era. This was slightly over a month after Air Force test pilot Capt. 

Charles E. Yeager first flew faster than sound in the Bell XS-1 rocket plane.9

Though ostensibly a simple demonstration model for a larger tunnel, the 11-inch tunnel itself became an important training and research 

tool that served to study a wide range of topics, including nozzle development and hypersonic flow visualization. It made practical contributions to aircraft development as well. Research with the 11-inch tunnel 

led to a key discovery incorporated on the X-15, namely that a wedgeshaped vertical tail markedly increased directional stability, eliminating the need for very large stabilizing surfaces. So useful was it that it 

remained in service until 1973, staying active even with a successor, the 

larger Continuous Flow Hypersonic Tunnel (CFHT), which entered service in 1962. The CFHT had a 31-inch test section and reached Mach 

10 but took a long time to become operational. Even after entering service, it operated much of the time in a blow-down mode rather than in 

continuous flow.10

9. John V. Becker, “Results of Recent Hypersonic and Unsteady Flow Research at the Langley Aeronautical Laboratory,” Journal of Applied Physics, vol. 21 (July 1950), pp. 619–628; Patrick J. Johnston and 

Wallace C. Sawyer, “An Historical Perspective on Hypersonic Aerodynamic Research at the Langley 

Research Center,” AIAA Paper 88-0230 (1988). For examples of its research, see Charles H. McLellan, Thomas W. Williams, and Mitchel H. Bertram, “Investigation of a Two-Step Nozzle in the Langley 

11-inch Hypersonic Tunnel,” NACA TN-2171 (1950); Charles H. McLellan and Thomas W. Williams, 

“Liquefaction of Air in the Langley 11-inch Hypersonic Tunnel,” NACA TN-3302 (1954).

10. John V. Becker, “The X-15 Project: Part I—Origins and Research Background,” Astronautics & Aeronautics, vol. 2, No. 2 (Feb. 1964), pp. 52–61; Charles H. McLellan, “A Method for Increasing the 

Effectiveness of Stabilizing Surfaces at High Supersonic Mach Numbers,” NACA RM-L54F21 (1954); 

Baals and Corliss, Wind Tunnels of NASA, pp. 56–57, 94–95; William T. Schaefer, Jr., “Characteristics of Major Active Wind Tunnels at the Langley Research Center,” NASA TM-X-1130 (1965), pp. 12, 27.
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Emergent Hypersonic Technology and the Onset of the Missile Era

The ballistic missile and atomic bomb became realities within a year of 

each other. At a stroke, the expectation arose that one might increase 

the range of the former to intercontinental distance and, by installing an 

atomic tip, generate a weapon—and a threat—of almost incomprehensible destructive power. But such visions ran afoul of perplexing technical issues involving rocket propulsion, guidance, and reentry. Engineers 

knew they could do something about propulsion, but guidance posed a 

formidable challenge. MIT’s Charles Stark Draper was seeking inertial 

guidance, but he couldn’t approach the Air Force requirement, which 

set an allowed miss distance of only 1,500 feet at a range of 5,000 miles 

for a ballistic missile warhead.11

Reentry posed an even more daunting prospect. A reentering 

5,000-mile-range missile would reach 9,000 kelvins, hotter than the solar 

surface, while its kinetic energy would vaporize five times its weight in 

iron.12 Rand Corporation studies encouraged Air Force and industry missile studies. Convair engineers, working under Karel J. “Charlie” Bossart, 

began development of the Atlas ICBM in 1951. Even with this seemingly 

rapid implementation of the ballistic missile idea, time scales remained 

long term. As late as October 1953, the Air Force declared that it would 

not complete research and development until “sometime after 1964.”13

Matters changed dramatically immediately after the Castle Bravo 

nuclear test on March 1, 1954, a weaponizable 15-megaton H-bomb, 

fully 1,000 times more powerful than the atomic bomb that devastated 

Hiroshima less than a decade previously. The “Teapot Committee,” 

chaired by the Hungarian emigree mathematician John von Neumann, 

had anticipated success with Bravo and with similar tests. Echoing 

Bruno Augenstein of the Rand Corporation, the Teapot group recom11. Jacob Neufeld, The Development of Ballistic Missiles in the United States Air Force, 1945–

1960 (Washington, DC: USAF, 1990), p. 293; Col. Edward N. Hall, USAF, “Air Force Missile 

Experience,” in Lt. Col. Kenneth F. Gantz, ed., The United States Air Force Report on the Ballistic 

Missile: Its Technology, Logistics, and Strategy (Garden City, NY: Doubleday & Co., Inc., 1958), 

pp. 47–59; Donald MacKenzie, Inventing Accuracy (Cambridge: MIT Press, 1990).

12. P.H. Rose and W.I. Stark, “Stagnation Point Heat-Transfer Measurements in Dissociated Air,” Journal 

of the Aeronautical Sciences, vol. 25, no. 2 (Feb. 1958), pp. 86–97.

13. John L. Chapman, Atlas: the Story of a Missile (New York: Harper & Brothers, 1960), pp. 

28–34, 74; Neufeld, Development of Ballistic Missiles, pp. 78, 44–50, 68–77; G. Harry Stine, 

ICBM: The Making of the Weapon that Changed the World (New York: Orion Books, 1996), pp. 

140–146, 162–174, 186–188.
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Extract of text from NACA Report 1381 (1953), in which H. Julian Allen and Alfred J. Eggers 

postulated using a blunt-body reentry shape to reduce surface heating of a reentry body. NASA.

mended that the Atlas miss distance should be relaxed “from the present 1,500 feet to at least two, and probably three, nautical miles.”14 This 

was feasible because the new H-bomb had such destructive power that 

such a “miss” distance seemed irrelevant. The Air Force leadership concurred, and only weeks after the Castle Bravo shot, in May 1954, Vice 

Chief of Staff Gen. Thomas D. White granted Atlas the service’s highest 

developmental priority.

But there remained the thorny problem of reentry. Only recently, 

most people had expected an ICBM nose cone to possess the needlenose sharpness of futurist and science fiction imagination. The realities 

of aerothermodynamic heating at near-orbital speeds dictated otherwise. In 1953, NACA Ames aerodynamicists H. Julian Allen and Alfred 

14. Bruno Augenstein, “Rand and North American Aviation’s Aerophysics Laboratory: An Early 

Interaction in Missiles and Space,” International Astronautical Federation, Paper IAA-98-IAA.2.2.06 

(1998); Neufeld, Development of Ballistic Missiles, pp. 259, 102–106, 117; Robert L. Perry, “The 

Atlas, Thor, Titan, and Minuteman,” in Eugene M. Emme, ed., The History of Rocket Technology: 

Essays on Research, Development, and Utility (Detroit: Wayne State University Press, 1964), pp. 

142–161.
NASA’s Contributions to Aeronautics

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Eggers concluded that an ideal reentry shape should be bluntly rounded, 

not sharply streamlined. A sharp nose produced a very strong attached 

shock wave, resulting in high surface heating. In contrast, a blunt nose 

generated a detached shock standing much further off the nose surface, allowing the airflow to carry away most of the heat. What heating 

remained could be alleviated via radiative cooling or by using hot structures and high-temperature coatings.15

There was need for experimental verification of blunt body theory, 

but the hypersonic wind tunnel, previously so useful, was suddenly inadequate, much as the conventional wind tunnel a decade earlier had been 

inadequate to obtaining the fullest understanding of transonic flows. As 

the slotted throat tunnel had replaced it, so now a new research tool, 

the shock tube, emerged for hypersonic studies. Conceived by Arthur 

Kantrowitz, a Langley veteran working at Cornell, the shock tube enabled 

far closer simulation of hypersonic pressures and temperatures. From 

the outset, Kantrowitz aimed at orbital velocity, writing in 1952 that: 

“it is possible to obtain shock Mach numbers in the neighborhood of 

25 with reasonable pressures and shock tube sizes.”16

Despite the advantages of blunt body design, the hypersonic environment remained so extreme that it was still necessary to furnish thermal protection to the nose cone. The answer was ablation: covering the 

nose with a lightweight coating that melts and flakes off to carry away 

the heat.Wernher von Braun’s U.S. Army team invented ablation while 

working on the Jupiter intermediate-range ballistic missile (IRBM), 

though General Electric scientist George Sutton made particularly notable contributions. He worked for the Air Force, which built and successfully protected a succession of ICBMs: Atlas, Titan, and Minuteman.17

15. H. Julian Allen and A.J. Eggers, Jr., “A Study of the Motion and Aerodynamic Heating of Ballistic Missiles Entering the Earth’s Atmosphere at High Supersonic Speeds,” NACA TR-1381 (1953); 

H. Julian Allen, “The Aerodynamic Heating of Atmospheric Entry Vehicles,” in J. Gordon Hall, ed., 

Fundamental Phenomena in Hypersonic Flow: Proceedings of the International Symposium Sponsored by Cornell Aeronautical Laboratory (Ithaca, NY: Cornell University Press, 1966), pp. 6–10; 

Edwin P. Hartman, Adventures in Research: A History of the Ames Research Center, 1940–1965,

NASA SP-4302 (Washington, DC: NASA, 1970), pp. 215–218.

16. E.L. Resler, Shao-Chi Lin, and Arthur Kantrowitz, “The Production of High Temperature Gases in 

Shock Tubes,” Journal of Applied Physics, vol. 23 (Dec. 1952), p. 1397.

17. Frank Kreith, Principles of Heat Transfer (Scranton, PA: International Textbook Co., 1965), pp. 

538–545; George W. Sutton, “The Initial Development of Ablation Heat Protection: an Historical 

Perspective,” Journal of Spacecraft and Rockets, vol. 19 (1982), pp. 3–11.
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A Jupiter IRBM launches from Cape Canaveral on May 18, 1958, on an ablation reentry test. 

U.S. Army.

Flight tests were critical for successful nose cone development, and 

they began in 1956 with launches of the multistage Lockheed X-17. It 

rose high into the atmosphere before firing its final test stage back at 

Earth, ensuring the achievement of a high-heat load, as the test nose 

cone would typically attain velocities of at least Mach 12 at only 40,000 

feet. This was half the speed of a satellite, at an altitude typically traversed by today’s subsonic airliners. In the pre-ablation era, the warheads typically burned up in the atmosphere, making the X-17 effectively 

a flying shock tube whose nose cones only lived long enough to return 

data by telemetry. Yet out of such limited beginnings (analogous to the 
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rudimentary test methodologies of the early transonic and supersonic 

era just a decade previously) came a technical base that swiftly resolved 

the reentry challenge.18

Tests followed with various Army and Air Force ballistic missiles. 

In August 1957, a Jupiter-C (an uprated Redstone) returned a nose cone 

after a flight of 1,343 miles. President Dwight D. Eisenhower subsequently showed it to the public during a TV appearance that sought to 

bolster American morale a month after Sputnik had shocked the world. 

Two Thor-Able flights went to 5,500 miles in July 1958, though their nose 

cones both were lost at sea. But the agenda also included Atlas, which 

first reached its full range of 6,300 miles in November 1958. Two nose 

cones built by GE, the RVX-1 and –2, flew subsequently as payloads. An 

RVX-2 flew 5,000 miles in July 1959 and was recovered, thereby becoming the largest object yet to be brought back. Attention now turned to a 

weaponized nose cone shape, GE’s Mark 3. Flight tests began in October, 

with this nose cone entering operational service the following April.19

Success in reentry now was a reality, yet there was much more for the 

future. The early nose cones were symmetric, which gave good ballistic characteristics but made no provision for significant aerodynamic maneuver and 

cross-range. The military sought both as a means of achieving greater operational flexibility. An Air Force experimental uncrewed lifting body design, 

the Martin SV-5D (X-23) PRIME, flew three flights between December 1966 

and April 1967, lofted over the Pacific Test Range by modified Atlas boosters. The first flew 4,300 miles, maneuvering in pitch (but not in cross-range), 

and missed its target aim point by only 900 feet. The third mission demonstrated a turning cross-range of 800 miles, the SV-5D impacting within 4 

miles of its aim point and subsequently was recovered.20

Other challenges remained. These included piloted return from the 

Moon, reusable thermal protection for the Shuttle, and planetary entry 

into the Jovian atmosphere, which was the most demanding of all. Even 

18. “Re-Entry Research: The Lockheed X-17,” Flight (Feb. 6, 1959), p. 181.

19. James M. Grimwood and Francis Strowd, History of the Jupiter Missile System (Huntsville, AL: 

U.S. Army Ordnance Missile Command, July 27, 1962), pp. 18–20; Time (Nov. 18, 1957, pp. 

19–20, and Dec. 8, 1958, p. 15); Joel W. Powell, “Thor-Able and Atlas Able,” Journal of the British Interplanetary Society, vol. 37, No. 5 (May 1984), pp. 219–225; General Electric, “Thermal 

Flight Test Summary Report for Mark 3 Mod 1 Re-Entry Vehicles” (1960), Defense Technical Information Center [DTIC] Report AD-362539; Convair, “Flight Test Evaluation Report, Missile 7D” (1959), 

DTIC AD-832686.

20. Martin Marietta, “SV-5 PRIME Final Flight Test Summary,” Report ER 14465 (1967).
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so, by the time of PRIME in 1967, the reentry problem had been resolved, 

manifested by the success of both ballistic missile nose cone development 

and the crewed spacecraft effort. The latter was arguably the most significant expression of hypersonic competency until the return to Earth 

from orbit by the Space Shuttle Columbia in 1981.

Transitioning from the Supersonic to the Hypersonic: X-7 to X-15

During the 1950s and early 1960s, aviation advanced from flight at high 

altitude and Mach 1 to flight in orbit at Mach 25. Within the atmosphere, a number of these advances stemmed from the use of the ramjet, at a time when turbojets could barely pass Mach 1 but ramjets could 

aim at Mach 3 and above. Ramjets needed an auxiliary rocket stage as 

a booster, which brought their general demise after high-performance 

afterburning turbojets succeeded in catching up. But in the heady days 

of the 1950s, the ramjet stood on the threshold of becoming a mainstream engine. Many plans and proposals existed to take advantage of 

their power for a variety of aircraft and missile applications.

The burgeoning ramjet industry included Marquardt and Wright 

Aeronautical, though other firms such as Bendix developed them as well. 

There were also numerous hardware projects. One was the Air ForceLockheed X-7, an air-launched high-speed propulsion, aerodynamic, 

and structures testbed. Two were surface-to-air ramjet-powered missiles: the Navy’s ship-based Mach 2.5+ Talos and the Air Force’s Mach 

3+ Bomarc. Both went on to years of service, with the Talos flying “in 

anger” as a MiG-killer and antiradiation SAM-killer in Vietnam. The Air 

Force also was developing a 6,300-mile-range Mach 3+ cruise missile—

the North American SM-64 Navaho—and a Mach 3+ interceptor fighter—

the Republic XF-103. Neither entered the operational inventory. The Air 

Force canceled the troublesome Navaho in July 1957, weeks after the first 

flight of its rival, Atlas, but some flight hardware remained, and Navaho 

flew in test for as far as 1,237 miles, though this was a rare success. The 

XF-103 was to fly at Mach 3.7 using a combined turbojet-ramjet engine. 

It was to be built largely of titanium, at a time when this metal was little 

understood; it thus lived for 6 years without approaching flight test. Still, 

its engine was built and underwent test in December 1956.21

21. Marcelle Size Knaack, Post-World War II Fighters, vol. 1 of Encyclopedia of U.S. Air Force 

Aircraft and Missile Systems (Washington, DC: Office of Air Force History, 1978), p. 329; Richard 

A. DeMeis, “The Trisonic Titanium Republic,” Air Enthusiast, vol. 7 (1978), pp. 198–213.
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The steel-structured X-7 proved surprisingly and consistently productive. The initial concept of the X-7 dated to December 1946 and constituted a three-stage vehicle. A B-29 (later a B-50) served as a “first stage” 

launch aircraft; a solid rocket booster functioned as a “second stage” 

accelerating it to Mach 2, at which the ramjet would take over. First flying 

in April 1951, the X-7 family completed 100 missions between 1955 and 

program termination in 1960. After achieving its Mach 3 design goal, the 

program kept going. In August 1957, an X-7 reached Mach 3.95 with a 

28-inch diameter Marquardt ramjet. The following April, the X-7 attained 

Mach 4.31—2,881 mph—with a more-powerful 36-inch Marquardt ramjet. This established an air-breathing propulsion record that remains 

unsurpassed for a conventional subsonic combustion ramjet.22

At the same time that the X-7 was edging toward the hypersonic frontier, the NACA, Air Force, Navy, and North American Aviation had a far 

more ambitious project underway: the hypersonic X-15. This was Round 

Two, following the earlier Round One research airplanes that had taken 

flight faster than sound. The concept of the X-15 was first proposed by 

Robert Woods, a cofounder and chief engineer of Bell Aircraft (manufacturer of the X-1 and X-2), at three successive meetings of the NACA’s 

influential Committee on Aerodynamics between October 1951 and June 

1952. It was a time when speed was king, when ambitious technologypushing projects were flying off the drawing board. These included the 

Navaho, X-2, and XF-103, and the first supersonic operational fighters—the Century series of the F-100, F-101, F-102, F-104, and F-105.23

Some contemplated even faster speeds. Walter Dornberger, former 

commander of the Nazi research center at Peenemünde turned senior Bell 

Aircraft Corporation executive, was advocating BoMi, a proposed skipgliding “Bomber-Missile” intended for Mach 12. Dornberger supported 

Woods in his recommendations, which were adopted by the NACA’s 

Executive Committee in July 1952. This gave them the status of policy, 

while the Air Force added its own support. This was significant because 

22. Lee L. Peterson, “Evaluation Report on X-7A,” AFMDC [Holloman AFB], ADJ 57-8184 (1957); 

and William A. Ritchie, “Evaluation Report on X-7A (System 601B),” AFMDC DAS-58-8129 (1959).

23. Robert S. Houston, Richard P. Hallion, and Ronald G. Boston, “Transiting from Air to Space: The 

North American X-15,” and John V. Becker, “The Development of Winged Reentry Vehicles: An Essay 

from the NACA-NASA Perspective, 1952–1963,” in Richard P. Hallion, ed., The Hypersonic Revolution: 

Eight Case Studies in the History of Hypersonic Technology, vol. 1: From Max Valier to Project PRIME, 

1924–1967 (Wright-Patterson AFB: Aeronautical Systems Division, 1987), pp. I–xii, No. 1, 383–386.
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its budget was 300 times larger than that of the NACA.24 The NACA alone 

lacked funds to build the X-15, but the Air Force could do this easily. It 

also covered the program’s massive cost overruns. These took the airframe from $38.7 million to $74.5 million and the large engine from $10 

million to $68.4 million, which was nearly as much as the airframe.25

The Air Force had its own test equipment at its Arnold Engineering 

Development Center (AEDC) at Tullahoma, TN, an outgrowth of the 

Theodore von Kármán technical intelligence mission that Army Air 

Forces Gen. Henry H. “Hap” Arnold had sent into Germany at the end 

of the Second World War. The AEDC, with brand-new ground test and 

research facilities, took care to complement, not duplicate, the NACA’s 

research facilities. It specialized in air-breathing and rocket-engine testing. Its largest installation accommodated full-size engines and provided 

continuous flow at Mach 4.75. But the X-15 was to fly well above this, to 

over Mach 6, highlighting the national facilities shortfall in hypersonic 

test capabilities existing at the time of its creation.26

While the Air Force had the deep pockets, the NACA—specifically 

Langley—conducted the research that furnished the basis for a design. This 

took the form of a 1954 feasibility study conducted by John Becker, assisted 

by structures expert Norris Dow, rocket expert Maxime Faget, configuration and controls specialist Thomas Toll, and test pilot James Whitten. 

They began by considering that during reentry, the vehicle should point 

its nose in the direction of flight. This proved impossible, as the heating 

was too high. He considered that the vehicle might alleviate this problem 

by using lift, which he was to obtain by raising the nose. He found that 

the thermal environment became far more manageable. He concluded 

that the craft should enter with its nose high, presenting its flat undersurface to the atmosphere. The Allen-Eggers paper was in print, and he 

later wrote that: “it was obvious to us that what we were seeing here was 

a new manifestation of H.J. Allen’s ‘blunt-body’ principle.”27

24. Harry Hansen, Engineer in Charge, NASA SP-4305, p. 428; Hansen, ed., The World Almanac and Book of Facts for 1956 (New York: New York World-Telegram Corp., 1956), p. 757.

25. Dennis Jenkins, X-15: Extending the Frontiers of Flight, NASA SP-2007-562 (Washington, DC: 

NASA, 2007), pp. 336–337.

26. U.S. Air Force Systems Command, History of the Arnold Engineering Development Center

(Arnold Air Force Station, TN: AEDC, n.d.); Julius Lukasiewicz, Experimental Methods of Hypersonics

(New York: Marcel Dekker, Inc., 1973), p. 247.

27. Becker, “Development of Winged Reentry Vehicles,” in Hallion, Hypersonic Revolution, vol. 1, p. 

386.
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To address the rigors of the daunting aerothermodynamic environment, Norris Dow selected Inconel X (a nickel alloy from International 

Nickel) as the temperature-resistant superalloy that was to serve for the 

aircraft structure. Dow began by ignoring heating and calculated the skin 

gauges needed only from considerations of strength and stiffness. Then 

he determined the thicknesses needed to serve as a heat sink. He found 

that the thicknesses that would suffice for the latter were nearly the same 

as those that would serve merely for structural strength. This meant that 

he could design his airplane and include heat sink as a bonus, with little 

or no additional weight. Inconel X was a wise choice; with a density of 

0.30 pounds per cubic inch, a tensile strength of over 200,000 pounds 

per square inch (psi), and yield strength of 160,000 psi, it was robust, 

and its melting temperature of over 2,500 °F ensured that the rigors of 

the anticipated 1,200 °F surface temperatures would not weaken it.28

Work at Langley also addressed the important issue of stability. Just 

then, in 1954, this topic was in the forefront because it had nearly cost the 

life of the test pilot Chuck Yeager. On the previous December 12, he had 

flown the X-1A to Mach 2.44 (approximately 1,650 mph). This exceeded 

the plane’s stability limits; it went out of control and plunged out of the 

sky. Only Yeager’s skill as a pilot had saved him and his airplane. The 

problem of stability would be far more severe at higher speeds.29

Analysis, confirmed by experiments in the 11-inch wind tunnel, had 

shown that most of the stability imparted by an aircraft’s tail surfaces 

was produced by its wedge-shaped forward portion. The aft portion 

contributed little to the effectiveness because it experienced lower air 

pressure. Charles McLellan, another Langley aerodynamicist, now 

proposed to address the problem of hypersonic stability by using tail surfaces that would be wedge-shaped along their entire length. Subsequent 

tests in the 11-inch tunnel, as mentioned previously, confirmed that 

this solution worked. As a consequence, the size of the tail surfaces 

shrank from being almost as large as the wings to a more nearly conventional appearance.30

28. Becker, “The X-15 Project,” pp. 52–61. Technical characteristics of Inconel X are from “Inconel 

X-750 Technical Data” (Sylmar, CA: High Temp Metals, Inc., 2009).

29. Richard P. Hallion, On the Frontier: Flight Research at Dryden, 1946–1982, SP-4303 (Washington, DC: NASA, 1984), pp. 70–71.

30. McLellan, “A Method for Increasing the Effectiveness of Stabilizing Surfaces,” NACA RML54F21 (1954).
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A schematic drawing of the X-15’s internal layout. NASA.

This study made it possible to proceed toward program approval and 

the award of contracts both for the X-15 airframe and its powerplant, a 

57,000-pound-thrust rocket engine burning a mix of liquid oxygen and 

anhydrous ammonia. But while the X-15 promised to advance the research 

airplane concept to over Mach 6, it demanded something more than the 

conventional aluminum and stainless steel structures of earlier craft such 

as the X-1 and X-2. Titanium was only beginning to enter use, primarily 

for reducing heating effects around jet engine exhausts and afterburners. 

Magnesium, which Douglas favored for its own high-speed designs, was 

flammable and lost strength at temperatures higher than 600 °F. Inconel 

X was heat-resistant, reasonably well known, and relatively easily worked. 

Accordingly, it was swiftly selected as the structural material of choice when 

Becker’s Langley team assessed the possibility of designing and fabricating 

a rocket-boosted air-launched hypersonic research airplane. The Becker 

study, completed in April 1954, chose Mach 6 as the goal and proposed 

to fly to altitudes as great as 350,000 feet. Both marks proved remarkably 

prescient: the X-15 eventually flew to 354,200 feet in 1963 and Mach 6.70 

in 1967. This was above 100 kilometers and well above the sensible atmosphere. Hence, at that early date, more than 3 years before Sputnik, Becker 

and his colleagues already were contemplating piloted flight into space.31

The X-15: Pioneering Piloted Hypersonics

North American Aviation won the contract to build the X-15. It first flew 

under power in September 1959, by which time an Atlas had hurled an 

31. John V. Becker, Norris F. Dow, Maxime A. Faget, Thomas A. Toll, and J.B. Whitten, “Research 

Airplane Study,” NACA Langley (April 1954).
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The North American X-15 at NASA’s Flight Research Center (now the Dryden Flight Research 

Center) in 1961. NASA.

RVX-2 nose cone to its fullest range. However, as a hypersonic experiment, 

the X-15 was a complete airplane. It thus was far more complex than a 

simple reentry body, and it took several years of cautious flight-testing 

before it reached peak speed of above Mach 6, and peak altitude as well.

Testing began with two so-called “Little Engines,” a pair of vintage 

Reaction Motors XLR11s that had earlier served in the X-1 series and the 

Douglas D-558-2 Skyrocket. Using these, the X-15 topped the records of 

the earlier X-2, reaching Mach 3.50 and 136,500 feet. Starting in 1961, 

using the “Big Engine”—the Thiokol XLR99 with its 57,000 pounds of 

thrust—the X-15 flew to its Mach 6 design speed and 50+ mile design altitude, with test pilot Maj. Robert White reaching Mach 6.04 and NASA 

pilot Joseph Walker an altitude of 354,200 feet. After a landing accident, 

the second X-15 was modified with external tanks and an ablative coating, 

with Air Force Maj. William “Pete” Knight subsequently flying this variant 

to Mach 6.70 (4,520 mph) in 1967. However, it sustained severe thermal 

damage, partly as a result of inadequate understanding of the interactions of impinging hypersonic shock-on-shock flows. It never flew again.32

The X-15’s cautious buildup proved a wise approach, for this gave leeway when problems arose. Unexpected thermal expansion leading to localized buckling and deformation showed up during early high-Mach flights. 

The skin behind the wing leading edge exhibited localized buckling after 

the first flight to Mach 5.3, but modifications to the wings eliminated hot 

32. Johnny G. Armstrong, “Flight Planning and Conduct of the X-15A-2 Envelope Expansion Program,” AFFTC TD-69-4 (1969).
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spots and prevented subsequent problems, enabling the airplane to reach 

beyond Mach 6. In addition, a flight to Mach 6.04 caused a windshield to 

crack because of thermal expansion. This forced redesign of its frame to 

incorporate titanium, which has a much lower coefficient of expansion. 

The problem—a rare case in which Inconel caused rather than resolved 

a heating problem—was fixed by this simple substitution.33

Altitude flights brought their own problems, involving potentially 

dangerous auxiliary power unit (APU) failures. These issues arose in 

1962 as flights began to reach well above 100,000 feet; the APUs began 

to experience gear failure after lubricating oil foamed and lost its lubricating properties. A different oil had much less tendency to foam; it now 

became standard. Designers also enclosed the APU gearbox within a 

pressurized enclosure. The gear failures ceased.34

The X-15 substantially expanded the use of flight simulators. These 

had been in use since the famed Link Trainer of Second World War and 

now included analog computers, but now they also took on a new role 

as they supported the development of control systems and flight equipment. Analog computers had been used in flight simulation since 1949. 

Still, in 1955, when the X-15 program began, it was not at all customary to use flight simulators to support aircraft design and development. 

But program managers turned to such simulators because they offered 

effective means to study new issues in cockpit displays, control systems, 

and aircraft handling qualities. A 1956 paper stated that simulation had 

“heretofore been considered somewhat of a luxury for high-speed aircraft,” but now “has been demonstrated as almost a necessity,” in all 

three axes, “to insure [sic] consistent and successful entries into the 

atmosphere.” Indeed, pilots spent much more time practicing in simulators than they did in actual flight, as much as an hour per minute of 

actual flying time.35

33. William H. Dana, “The X-15 Airplane—Lessons Learned,” AIAA Paper 93-0309 (1993); 

Joseph Weil, “Review of the X-15 Program,” NASA TN-D-1278 (1962).

34. Perry V. Row and Jack Fischel, “X-15 Flight Test Experience,” Astronautics and Aerospace Engineering, vol. 1 (June 1963), pp. 25–32.

35. Quotes from “Research Airplane Committee Report on Conference on the Progress of the X-15 

Project,” NACA Langley Aeronautical Laboratory, 1956, p. 84; James I. Kilgore, “The Planes that Never 

Leave the Ground,” American Heritage of Invention and Technology (winter 1989), pp. 60–62; John P. 

Smith, Lawrence J. Schilling, and Charles A. Wagner, “Simulation at Dryden Flight Research Facility from 

1957 to 1982,” NASA TM-101695 (1989), p. 4; Milton O. Thompson, At the Edge of Space: The 

X-15 Flight Program (Washington, DC: Smithsonian Institution Press, 1992), pp. 70–71.
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The most important flight simulator was built by North American. 

Located originally in Los Angeles, Paul Bikle, the Director of NASA’s 

Flight Research Center, moved it to that Center in 1961. It replicated the 

X-15 cockpit and included actual hydraulic and control-system hardware. Three analog computers implemented equations of motion that 

governed translation and rotation of the X-15 about all three axes, transforming pilot inputs into instrument displays.36

The North American simulator became critical in training X-15 pilots 

as they prepared to execute specific planned flights. A particular mission 

might take little more than 10 minutes, from ignition of the main engine 

to touchdown on the lakebed, but a test pilot could easily spend 10 hours 

making practice runs in this facility. Training began with repeated trials 

of the normal flight profile with the pilot in the simulator cockpit and a 

ground controller close at hand. The pilot was welcome to recommend 

changes, which often went into the flight plan. Next came rehearsals of 

off-design missions: too much thrust from the main engine, too high a 

pitch angle when leaving the stratosphere.

Much time was spent practicing for emergencies. The X-15 had an 

inertial reference unit that used analog circuitry to display attitude, altitude, velocity, and rate of climb. Pilots dealt with simulated failures in 

this unit as they worked to complete the normal mission or, at least, to 

execute a safe return. Similar exercises addressed failures in the stability augmentation system. When the flight plan raised issues of possible 

flight instability, tests in the simulator used highly pessimistic assumptions concerning stability of the vehicle. Other simulations introduced 

in-flight failures of the radio or Q-ball multifunction sensor. Premature 

engine shutdown imposed a requirement for safe landing on an alternate lakebed that was available for emergency use.37

The simulations indeed had realistic cockpit displays, but they left 

out an essential feature: the g-loads, produced both by rocket thrust and 

by deceleration during reentry. In addition, a failure of the stability augmentation system, during reentry, could allow the airplane to oscillate 

36. NASA FRC, “Experience with the X-15 Adaptive Flight Control System,” NASA TN-D-6208 

(1971); Perry V. Row and Jack Fischel, “Operational Flight-test Experience with the X-15 Airplane,” 

AIAA Paper 63-075 (1963).

37. Wendell H. Stillwell, X-15 Research Results, NASA SP-60 (Washington, DC: NASA, 1965), 

pp. 37–38; Robert G. Hoey and Richard E. Day, “Mission Planning and Operational Procedures 

for the X-15 Airplane,” NASA TN-D-1158 (1962), NTRS Document ID 19710070140.
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in pitch and yaw. This changed the drag characteristics and imposed a 

substantial cyclical force.

To address such issues, investigators installed a flight simulator 

within the gondola of an existing centrifuge at the Naval Air Development 

Center in Johnsville, PA. The gondola could rotate on two axes while 

the centrifuge as a whole was turning. It not only produced g-forces; 

its g-forces increased during the simulated rocket burn. The centrifuge 

imposed such forces anew during reentry while adding a cyclical component to give the effect of an oscillation in yaw or pitch.38

There also were advances in pressure suits, under development 

since the 1930s. Already an early pressure suit had saved the life of Maj. 

Frank K. Everest during a high-altitude flight in the X-1, when it had 

suffered cabin decompression from a cracked canopy. Marine test pilot 

Lt. Col. Marion Carl had worn another during a flight to 83,235 feet in 

the D-558-2 Skyrocket in 1953, as had Capt. Iven Kincheloe during his 

record flight to 126,200 feet in the Bell X-2 in 1956. But these early suits, 

while effective in protecting pilots, were almost rigid when inflated, 

nearly immobilizing them. In contrast, the David G. Clark Company, a 

girdle manufacturer, introduced a fabric that contracted in circumference while it stretched in length. An exchange between these effects created a balance that maintained a constant volume, preserving a pilot’s 

freedom of movement. The result was the Clark MC-2 suit, which, in 

addition to the X-15, formed the basis for American spacesuit development from Project Mercury forward. Refined as the A/P22S-2, the X-15’s 

suit became the standard high-altitude pressure suit for NASA and the 

Air Force. It formed the basis for the Gemini suit and, after 1972, was 

adopted by the U.S. Navy as well, subsequently being employed by pilots 

and aircrew in the SR-71, U-2, and Space Shuttle.39

38. C.C. Clark and C.H. Woodling, “Centrifuge Simulation of the X-15 Research Aircraft,” NADC 

MA-5916 (1959); Jenkins, X-15, p. 279; NASA, “Research Airplane Committee Report on Conference on the Progress of the X-15 Project,” (1958), pp. 107–116.

39. For the Clark suit’s development, see A. Scott Crossfield with Clay Blair, Jr., Always Another 

Dawn: The Story of a Rocket Test Pilot (Cleveland: The World Publishing Co., 1960), pp. 

253–261; Paul Crickmore, Lockheed SR-71 Blackbird (London: Osprey Publishing Ltd., 1986), pp. 

100–102; T.A. Heppenheimer, History of the Space Shuttle, vol. 2, Development of the Shuttle, 

1972–81 (Washington, DC: Smithsonian Institution Press, 2002), pp. 274–277; Jenkins, X-15,

pp. 131–146; Loyd S. Swenson, James M. Grimwood, and Charles Alexander, This New Ocean: 

A History of Project Mercury, NASA SP-4201 (Washington, DC: NASA, 1998), pp. 225–231.
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The X-15 also accelerated development of specialized instrumentation, including a unique gimbaled nose sensor developed by Northrop. It 

furnished precise speed and positioning data by evaluation of dynamic 

pressure (“q” in aero engineering shorthand), and thus was known as 

the Q-ball. The Q-ball took the form of a movable sphere set in the nose 

of the craft, giving it the appearance of the enlarged tip of a ballpoint 

pen. “The Q-ball is a go-no go item,” NASA test pilot Joseph Walker told 

Time magazine reporters in 1961, adding: “Only if she checks okay do 

we go.”40 The X-15 also incorporated “cold jet” hydrogen peroxide reaction controls for maintaining vehicle attitude in the tenuous upper atmosphere, when dynamic air pressure alone would be insufficient to permit 

adequate flight control functionality. When Iven Kincheloe reached 

126,200 feet, his X-2 was essentially a free ballistic object, uncontrollable in pitch, roll, and yaw as it reached peak altitude and then began its 

descent. This situation made reaction controls imperative for the new 

research airplane, and the NACA (later NASA) had evaluated them on 

a so-called “Iron Cross” simulator on the ground and then in flight on 

the Bell X-1B and on a modified Lockheed F-104 Starfighter. They then 

proved their worth on the X-15 and, as with the Clark pressure suit, 

were incorporated on Mercury and subsequent American spacecraft.

The X-15 introduced a side stick flight controller that the pilot would 

utilize during acceleration (when under loads of approximately 3 g’s), 

relying on a fighter-type conventional control column for approach and 

landing. The third X-15 had a very different flight control system than the 

other two, differing greatly from the now-standard stability-augmented 

hydromechanical system carried by operational military and civilian 

aircraft. The third aircraft introduced a so-called “adaptive” flight control system, the MH-96. Built by Minneapolis Honeywell, the MH-96 

relied on rate gyros, which sensed rates of motion in pitch, roll, and 

yaw. It also incorporated “gain,” defined as the proportion between 

sensed rates of angular motion and a deflection of the ailerons or other 

controls. This variable gain, which changed automatically in response 

to flight conditions, functioned to maintain desired handling qualities 

across the spectrum of X-15 performance. This arrangement made it 

possible to introduce blended reaction and aerodynamic controls on the 

same stick, with this blending occurring automatically in response to 

40. Time, Oct. 27, 1961, p. 89.
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the values determined for gain as the X-15 flew out of the atmosphere 

and back again. Experience, alas, would reveal the MH-96 as an immature, troublesome system, one that, for all its ambition, posed significant headaches. It played an ultimately fatal role in the loss of X-15 pilot 

Maj. Michael Adams in 1967.41

The three X-15s accumulated a total of 199 flights from 1959 through 

1968. As airborne instruments of hypersonic research, they accumulated nearly 9 hours above Mach 3, close to 6 hours above Mach 4, 

and 87 minutes above Mach 5. Many concepts existed for X-15 derivatives and spinoffs, including using it as a second stage to launch small 

satellite-lofting boosters, to be modified with a delta wing and scramjet, and even to form the basis itself for some sort of orbital spacecraft; 

for a variety of reasons, NASA did not proceed with any of these. More 

significantly, however, was the strong influence the X-15 exerted upon 

subsequent hypersonic projects, particularly the National Hypersonic 

Flight Research Facility (NHFRF, pronounced “nerf”), intended 

to reach Mach 8.

A derivative of the Air Force Flight Dynamics Laboratory’s X-24C 

study effort, NHFRF was also to cruise at Mach 6 for 40 seconds. A joint 

Air Force-NASA committee approved a proposal in July 1976 with an 

estimated program cost of $200 million, and NHFRF had strong support 

from NASA’s hypersonic partisans in the Langley and Dryden Centers. 

Unfortunately, its rising costs, at a time when the Shuttle demanded an 

ever-increasing proportion of the Agency’s budget and effort, doomed 

it, and it was canceled in September 1977. Overall, the X-15 set speed 

and altitude records that were not surpassed until the advent of 

the Space Shuttle.42

41. Dana, “The X-15 Airplane—Lessons Learned,” AIAA Paper 93-0309 (1993); Thompson, At the 

Edge of Space, pp. 200–202; Lawrence W. Taylor and George B. Merrick, “X-15 Airplane Stability Augmentation Systems,” NASA TN-D-1157 (1962); Robert A. Tremant, “Operational Experience 

and Characteristics of the X-15 Flight Control System,” NASA TN-D-1402 (Dec. 1962), Donald R. 

Bellman, et al., Investigation of the Crash of the X-15-3 Aircraft on November 15, 1967 (Edwards: 

NASA Flight Research Center, Jan. 1968), pp. 8–15.

42. Kenneth E. Hodge, et al., Proceedings of the X-15 First Flight 30th Anniversary Celebration, CP 

3105 (Edwards: NASA, June 8, 1989); Hallion, On the Frontier, pp. 170–172; Donald P. Hearth 

and Albert E. Preyss, “Hypersonic Technology: Approach to an Expanded Program,” Astronautics and 

Aeronautics, (Dec. 1976), pp. 20–37; “NASA to End Hypersonic Effort,” Aviation Week and Space 

Technology (Sept. 26, 1977).
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The X-20 Dyna-Soar

During the 1950s, as the X-15 was taking shape, a parallel set of initiatives sought to define a follow-on hypersonic program that could 

actually achieve orbit. They were inspired in large measure by the 

1938–1944 Silbervögel (“Silver Bird”) proposal of Austrian space flight 

advocate Eugen Sänger and his wife, mathematician Irene Sänger-Bredt, 

which greatly influenced postwar Soviet, American, and European thinking about hypersonics and long-range “antipodal” flight. Influenced by 

Sänger’s work and urged onward by the advocacy of Walter Dornberger, 

Bell Aircraft Corporation in 1952 proposed the BoMi, intended to fly 

3,500 miles. Bell officials gained funding from the Air Force’s Wright 

Air Development Center (WADC) to study longer-range 4,000-mile and 

6,000-mile systems under the aegis of Air Force project MX-2276.

Support took a giant step forward in February 1956, when Gen. 

Thomas Power, Chief of the Air Research and Development Command 

(ARDC, predecessor of Air Force Systems Command) and a future Air 

Force Chief of Staff, stated that the service should stop merely considering such radical craft and instead start building them. With this level 

of interest, events naturally moved rapidly. A month later, Bell received 

a study contract for Brass Bell, a follow-on Mach 15 rocket-lofted boostglider for strategic reconnaissance. Power preferred another orbital 

glider concept, RoBo (for Rocket Bomber), which was to serve as a 

global strike system. To accelerate transition of hypersonics from the 

research to the operational community, the ARDC proposed its own 

concept, Hypersonic Weapons Research and Development Supporting 

System (HYWARDS). With so many cooks in the kitchen, the Air Force 

needed a coordinated plan. An initial step came in December 1956, as 

Bell raised the velocity of Brass Bell to Mach 18. A month later, a group 

headed by John Becker, at Langley, recommended the same design goal 

for HYWARDS. RoBo still remained separate, but it emerged as a longterm project that could be operational by the mid-1970s.43

NACA researchers split along centerlines over the issue of what kind 

of wing design to employ for HYWARDS. At NACA Ames, Alfred Eggers 

and Clarence Syvertson emphasized achieving maximum lift. They proposed a high-wing configuration with a flat top, calculating its hypersonic 

43. Clarence J. Geiger, “Strangled Infant: The Boeing X-20A Dyna-Soar,” in Hallion, Hypersonic Revolution, vol. 1, pp. 189–201; Capt. Roy F. Houchin, “The Rise and Fall of Dyna-Soar: A History of Air 

Force Hypersonic R&D, 1944–1963,” Air Force Institute of Technology (1995), DTIC ADA-303832.
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life-to-drag (L/D) as 6.85 and measuring a value of 6.65 during hypersonic tunnel tests. Langley researchers John Becker and Peter Korycinski 

argued that Ames had the configuration “upside down.” Emphasizing 

lighter weight, they showed that a flat-bottom Mach 18 shape gave a 

weight of 21,400 pounds, which rose only modestly at higher speeds. 

By contrast, the Ames “flat-top” weight was 27,600 pounds and rising 

steeply. NASA officials diplomatically described the Ames and Langley 

HYWARDS concepts respectively as “high L/D” and “low heating,” but 

while the imbroglio persisted, there still was no acceptable design. It 

fell to Becker and Korycinski to break the impasse in August 1957, and 

they did so by considering heating. It was generally expected that such 

craft required active cooling. But Becker and his Langley colleagues 

found that a glider of global range achieved peak uncooled skin temperatures of 2,000 °F, which was survivable by using improved materials. Accordingly, the flat-bottom design needed no coolant, dramatically 

reducing both its weight and complexity.44

This was a seminal conclusion that reshaped hypersonic thinking 

and influenced all future development down to the Space Shuttle. In 

October 1957, coincident with the Soviet success with Sputnik, the 

ARDC issued a coordinated plan that anticipated building HYWARDS 

for research at 18,000 feet per second, following it with Brass Bell for 

reconnaissance at the same speed and then RoBo, which was to carry 

nuclear bombs into orbit. HYWARDS now took on the new name of 

Dyna-Soar, for “Dynamic Soaring,” an allusion to the Sänger-legacy 

skip-gliding hypersonic reentry. (It was later designated X-20.) To the 

NACA, it constituted a Round Three following the Round One X-1, X-2, 

and Skyrocket, and the Round Two X-15.

The flat-bottom configuration quickly showed that it was robust 

enough to accommodate flight at much higher speeds. In 1959, Herbert 

York, the Defense Director of Research and Engineering, stated that 

Dyna-Soar was to fly at 15,000 mph, lofted by the Martin Company’s 

Titan I missile, though this was significantly below orbital speed. But 

44. John V. Becker, “The Development of Winged Reentry Vehicles: An Essay from the NACANASA Perspective, 1952–1963,” in Hallion, Hypersonic Revolution, vol. 1, pp. 391–407; Alvin 

Seiff and H. Julian Allen, “Some Aspects of the Design of Hypersonic Boost-Glide Aircraft,” NACA 

RM-A55E26 (1955); Alfred J. Eggers and Clarence Syvertson, “Aircraft Configurations Developing High Lift-Drag Ratios at High Supersonic Speeds,” NACA RM-A55L05 (1956); Hansen, 

Engineer in Charge, pp. 467–473.
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This 1957 Langley trade-study shows weight advantage of flat-bottom reentry vehicles at higher 

Mach numbers. This led to abandonment of high-wing designs in favor of flat-bottom ones such 

as the X-20 Dyna-Soar and the Space Shuttle. NASA.

during subsequent years it changed to the more-capable Titan II and 

then to the powerful Titan III-C. With two solid-fuel boosters augmenting its liquid hypergolic main stage, it could easily boost Dyna-Soar to 

the 18,000 mph necessary for it to achieve orbit. A new plan of December 

1961 dropped suborbital missions and called for “the early attainment 

of orbital flight.”45

By then, though, Dyna-Soar was in deep political trouble. It had 

been conceived initially as a prelude to the boost-glider Brass Bell for 

45. Capt. Roy Houchin, “Hypersonic Technology and Aerospace Doctrine,” Air Power History,

vol. 46, no. 3 (fall 1999), pp. 4–17; Terry L. Sunday and John R. London, “The X-20 Space Plane: 

Past Innovation, Future Vision,” in John Becklake, ed., History of Rocketry and Astronautics, vol. 17 

(San Diego: American Astronautical Society/Univelt, 1995), pp. 253–284.
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This full-size mockup of the X-20 gives an indication of its small, compact design. USAF.

reconnaissance and to the orbital RoBo for bombardment. But Brass 

Bell gave way to a purpose-built concept for a small-piloted station, the 

Manned Orbiting Laboratory (MOL), which could carry more sophisticated reconnaissance equipment. (Ironically, though a team of MOL 

astronauts was selected, MOL itself likewise was eventually canceled.) 

RoBo, a strategic weapon, fell out of the picture completely, for the 

success of the solid-propellant Minuteman ICBM established the silolaunched ICBM as the Nation’s prime strategic force, augmented by the 

Navy’s fleet of Polaris-launching ballistic missile submarines.46

In mid-196l, Secretary of Defense Robert S. McNamara directed the 

Air Force to justify Dyna-Soar on military grounds. Service advocates 

responded by proposing a host of applications, including orbital reconnaissance, rescue, inspection of Soviet spacecraft, orbital bombardment, 

46. Wyndham D. Miles, “The Polaris,” in Emme, History of Rocket Technology, pp. 162–175.
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and use of the craft as a ferry vehicle. McNamara found these rationalizations unconvincing but was willing to allow the program to proceed 

as a research effort, at least for the time being. In an October 1961 memo 

to President John F. Kennedy, he proposed to “re-orient the program to 

solve the difficult technical problems involved in boosting a body of high 

lift into orbit, sustaining man in it and recovering the vehicle at a designated place.”47 This reorientation gave the project 2 more years of life.

Then in 1963, he asked what the Air Force intended to do with it 

after using it to demonstrate maneuvering entry. He insisted he could 

not justify continuing the program if it was a dead-end effort with no 

ultimate purpose. But it had little potential utility, for it was not a cargo 

rocket, nor could it carry substantial payloads, nor could it conduct longduration missions. And so, in December McNamara canceled it, after 6 

years of development time, a Government contract investment of $410 

million, the expenditure of 16 million man-hours by nearly 8,000 contractor personnel, 14,000 hours of wind tunnel testing, 9,000 hours of 

simulator runs, and the preparation of 3,035 detailed technical reports.48

Ironically, by time of its cancellation, the X-20 was so far advanced 

that the Air Force had already set aside a block of serial numbers for 

the 10 production aircraft. Its construction was well underway, Boeing 

having completed an estimated 42 percent of design and fabrication 

tasks.49 Though the X-20 never flew, portions of its principal purposes 

were fulfilled by other programs. Even before cancellation, the Air 

Force launched the first of several McDonnell Aerothermodynamic/

elastic Structural Systems Environmental Test (ASSET) hot-structure 

radiative-cooled flat-bottom cone-cylinder shapes sharing important 

configuration similarities to the Dyna-Soar vehicle. Slightly later, its 

Project PRIME demonstrated cross-range maneuver after atmospheric 

entry. This used the Martin SV-5D lifting body, a vehicle differing 

significantly from the X-20 but which complemented it nonetheless. 

In this fashion, the Air Force succeeded at least partially in obtaining lifting reentry data from winged vehicles and lifting bodies that 

widened the future prospects for reentry.

47. Curtis Peebles, “The Origin of the U.S. Space Shuttle—1,” Spaceflight, vol. 21, no. 11 (Nov. 

1979), pp. 435–442.

48. Geiger, “Strangled Infant,” in Hallion, Hypersonic Revolution, vol. 1, pp. 313, 319–320.

49. Ibid., pp. 294–310, 313.
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Hot Structures and Return from Space: X-20’s Legacy and ASSET

Dyna-Soar never flew, but it sharply extended both the technology and 

the temperature limits of hot structures and associated aircraft elements, at a time when the American space program was in its infancy.50

The United States successfully returned a satellite from orbit in April 

1959, while ICBM nose cones were still under test, when the Discoverer 

II test vehicle supporting development of the National Reconnaissance 

Office’s secret Corona spy satellite returned from orbit. Unfortunately, it 

came down in Russian-occupied territory far removed from its intended 

recovery area near Hawaii. Still, it offered proof that practical hypersonic reentry and recovery were at hand.

An ICBM nose cone quickly transited the atmosphere, whereas recoverable satellite reentry took place over a number of minutes. Hence a satellite encountered milder aerothermodynamic conditions that imposed 

strong heat but brought little or no ablation. For a satellite, the heat of 

ablation, measured in British thermal units (BTU) per pound of protective material, was usually irrelevant. Instead, insulative properties were 

more significant: Teflon, for example, had poor ablative properties but 

was an excellent insulator.51

Production Dyna-Soar vehicles would have had a four-flight service life before retirement or scrapping, depending upon a hot structure 

comprised of various materials, each with different but complementary 

properties. A hot structure typically used a strong material capable of 

withstanding intermediate temperatures to bear flights loads. Set off 

from it were outer panels of a temperature-resistant material that did 

not have to support loads but that could withstand greatly elevated temperatures as high as 3,000 °F. In between was a lightweight insulator (in 

Dyna-Soar’s case, Q-felt, a silica fiber from the firm of Johns Manville). 

It had a tendency to shrink, thus risking dangerous gaps where high 

heat could bypass it. But it exhibited little shrinkage above 2,000 °F 

50. Boeing, “Summary of Technical Advances: X-20 Program,” Report D2-23418 (July 1964).

51. Robert L. Perry, Management of the National Reconnaissance Program, 1960–1965 (Chantilly, 

VA: NRO, 2001 edition of a Jan. 1969 work), p. 9; Jeffrey Richelson, American Espionage and the 

Soviet Target (New York: William Morrow, 1987), p. 184; F.R. Riddell and J.D. Teare, “The Differences Between Satellite and Ballistic Missile Re-Entry Problems,” in Morton Alperin and Hollingsworth 

F. Gregory, eds., Vistas in Aeronautics, vol. 2 (New York: Pergamon Press, 1959), pp. 174–190; 

Leo Steg, “Materials for Re-Entry Heat Protection of Satellites,” American Rocket Society Journal (Sept. 

1960), pp. 815–822.
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and could withstand 3,000 °F. By “preshrinking” this material, it qualified for operational use.52

For its primary structure, Dyna-Soar used René 41, a nickel alloy that 

included chromium, cobalt, and molybdenum. Its use was pioneered by 

General Electric for hot-section applications in its jet engines. The alloy 

had room temperature yield strength of 130,000 psi, declining slightly at 

1,200 °F, and was still strong at 1,800 °F. Some of the X-20’s panels were 

molybdenum alloy, which offered clear advantages for such hot areas as 

the wing leading edges. D-36 columbium alloy covered most other areas 

of the vehicle, including the flat underside of the wings.

These panels had to resist flutter, which brought a risk of cracking 

because of fatigue, as well as permitting the entry of superheated hypersonic 

flows that could destroy the internal structure within seconds. Because of 

the risks to wind tunnels from hasty and ill-considered flutter testing (where 

a test model for example can disintegrate, damaging the interior of the tunnel), X-20 flutter testing consumed 18 months of Boeing’s time. Its people 

started testing at modest stress levels and reached levels that exceeded the 

vehicle’s anticipated design requirements.53

The X-20’s nose cap had to function in a thermal and dynamic pressure environment more extreme even than that experienced by the X-15’s 

Q-ball. It was a critical item that faced temperatures of 3,680 °F, accompanied by a daunting peak heat flux of 143 BTU per square foot per 

second. Both Boeing and its subcontractor Chance Vought pursued independent approaches to development, resulting in two different designs. 

Vought built its cap of siliconized graphite with an insulating layer of 

a temperature-resistant zirconium oxide ceramic tiles. Their melting 

point was above 4,500 °F, and they covered its forward area, being held 

in place by thick zirconium oxide pins. The Boeing design was simpler, 

using a solid zirconium oxide nose cap reinforced against cracking with 

two screens of platinum-rhodium wire. Like the airframe, the nose caps 

were rated through four orbital flights and reentries.54

52. Geiger, “Strangled Infant,” in Hallion, Hypersonic Revolution, vol. 1, pp. 347–370.

53. Aeronautical Systems Division, Proceedings of 1962 X-20A (Dyna-Soar) Symposium, vol. 3: 

Structures and Materials (Wright-Patterson AFB, OH: USAF, Mar. 1963), DTIC AD-346192; Howard J. Middendorf, “Materials and Processes for X-20A (Dyna-Soar),” Air Force Systems Command 

(June 1964), DTIC AD-449685; and William Cowie, “Utilization of Refractory Metals on the X-20A 

(Dyna-Soar),” Air Force Systems Command (June 1964), DTIC AD-609169.

54. ASD, X-20A Proceedings, vol. 3, DTIC AD-346192.
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Generally, the design of the X-20 reflected the thinking of Langley’s 

John Becker and Peter Korycinski. It relied on insulation and radiation of the accumulated thermal load for primary thermal protection. 

But portions of the vehicle demanded other approaches, with specialized areas and equipment demanding specialized solutions. Ball bearings, facing a 1,600 °F thermal environment, were fabricated as small 

spheres of René 41 nickel alloy covered with gold. Antifriction bearings 

used titanium carbide with nickel as a binder. Antenna windows had to 

survive hot hypersonic flows yet be transparent to radio waves. A mix 

of oxides of cobalt, aluminum, and nickel gave a coating that showed a 

suitable emittance while furnishing requisite temperature protection.

The pilot looked through five clear panes: three that faced forward 

and two on the sides. The three forward panes were protected by a jettisonable protective shield and could only be used below Mach 5 after reentry, but the side ones faced a less severe aerothermodynamic environment 

and were left unshielded. But could the X-20 be landed if the protective 

shield failed to jettison after reentry? NASA test pilot Neil Armstrong, 

later the first human to set foot upon the Moon, flew approaches using a 

modified Douglas F5D Skylancer. He showed it was possible to land the 

Dyna-Soar using only visual cues obtained through the side windows.

The cockpit, equipment bay, and a power bay were thermally isolated and cooled via a “water wall” using lightweight panels filled with 

a jelled water mix. The hydraulic system was cooled as well. To avoid 

overheating and bursting problems with conventional inflated rubber 

tires, Boeing designed the X-20 to incorporate tricycle-landing skids with 

wire brush landing pads.55 Dyna-Soar, then, despite never having flown, 

significantly advanced the technology of hypersonic aerospace vehicle 

design. Its contributions were many and can be illustrated by examining the confidence with which engineers could approach the design of 

critical technical elements of a hypersonic craft, in 1958 (the year North 

American began fabricating the X-15) and 1963 (the year Boeing began 

fabricating the X-20):56 In short, within the 5 years that took the X-20 

from a paper study to a project well underway, the “art of the possible” 

55. Ibid.; Geiger, “Strangled Infant,” in Hallion, Hypersonic Revolution, vol. 1, pp. 347–349, 

361–370.

56. Geiger, “Strangled Infant,” in Hallion, Hypersonic Revolution, vol. 1, pp. 344–346; R.L. 

Schleicher, “Structural Design of the X-15,” Journal of the Royal Aeronautical Society (Oct. 1963), 

pp. 618–636.
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TABLE 1

INDUSTRY HYPERSONIC “DESIGN CONFIDENCE” 

AS MEASURED BY ACHIEVABLE DESIGN TEMPERATURE CRITERIA, °F

ELEMENT X-15 X-20

Nose cap 3,200 4,300

Surface panels 1,200 2,750

Primary structure 1,200 1,800

Leading edges 1,200 3,000

Control surfaces 1,200 1,800

Bearings 1,200 1,800

in hypersonics witnessed a one-third increase in possible nose cap temperatures, a more than double increase in the acceptable temperatures of 

surface panels and leading edges, and a one-third increase in the acceptable temperatures of primary structures, control surfaces, and bearings.

The winddown and cancellation of Dyna-Soar coincided with the 

first flight tests of the much smaller but nevertheless still very technically ambitious McDonnell ASSET hypersonic lifting reentry test vehicle. 

Lofted down the Atlantic Test Range on modified Thor and Thor-Delta 

boosters, they demonstrated reentry at over Mach 18. ASSET dated to 

1959, when Air Force hypersonic advocates advanced it as a means of 

assessing the accuracy of existing hypersonic theory and predictive techniques. In 1961, McDonnell Aircraft, a manufacturer of fighter aircraft 

and also the Project Mercury spacecraft, began design and fabrication of 

ASSET’s small sharply swept delta wing flat-bottom boost-gliders. They 

had a length of 69 inches and a span of 55 inches.

Though in many respects they resembled the soon-to-be-canceled 

X-20, unlike that larger, crewed transatmospheric vehicle, the ASSET 

gliders were more akin to lifting nose cone shapes. Instead of the X-20’s 

primary reliance upon René 41, the ASSET gliders largely used columbium alloys, with molybdenum alloy on their forward lower heat shield, 

graphite wing leading edges, various insulative materials, and columbium, molybdenum, and graphite coatings as needed. There were also 

three nose caps: one fabricated from zirconium oxide rods, another 

from tungsten coated with thorium, and a third of siliconized graphite 

coated with zirconium oxide. Though all six ASSETs looked alike, they 

were built in two differing variants: four Aerothermodynamic Structural 

Vehicles (ASV) and two Aerothermodynamic Elastic Vehicles (AEV). The 

former reentered from higher velocities (between 16,000 and 19,500 feet 
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5

per second) and altitudes (from 202,000 to 212,000 feet), necessitating 

use of two-stage Thor-Delta boosters. The latter (only one of which flew 

successfully) used a single-stage Thor booster and reentered at 13,000 

feet per second from an altitude of 173,000 feet. It was a hypersonic flutter research vehicle, analyzing as well the behavior of a trailing-edge flap 

representing a hypersonic control surface. Both the ASV and AEV flew 

with a variety of experimental panels installed at various locations and 

fabricated by Boeing, Bell, and Martin.57 The ASSET program conducted 

six flights between September 1963 and February 1965, all successful 

save for one AEV launch in March 1964. Though intended for recovery 

from the Atlantic, only one survived the rigors of parachute deployment, 

descent, and being plunged into the ocean. But that survivor, the ASV3, proved to be in excellent condition, with the builder, International 

Harvester, rightly concluding it “could have been used again.”58 ASV-4, 

the best flight flown, was also the last one, with the final flight-test report 

declaring that it returned “the highest quality data of the ASSET program.” It flew at a peak speed of Mach 18.4, including a hypersonic glide 

that covered 2,300 nautical miles.59

Overall, the ASSET program scored a host of successes. It was all the 

more impressive for the modest investment made in its development: just 

$21.2 million. It furnished the first proof of the magnitude and seriousness of upper-surface leeside heating and the dangers of hypersonic flow 

impingement into interior structures. It dealt with practical issues of fabrication, including fasteners and coatings. It contributed to understanding of hypersonic flutter and of the use of movable control surfaces. It 

also demonstrated successful use of an attitude-adjusting reaction control system, in near vacuum and at speeds much higher than those of the 

X-15. It complemented Dyna-Soar and left the aerospace industry believing that hot structure design technology would be the normative technical approach taken on future launch vehicles and orbital spacecraft.60

57. Richard P. Hallion, “ASSET: Pioneer of Lifting Reentry,” in Hallion, ed., Hypersonic Revolution, 

vol. 1, pp. 451, 461–465, 501–505; USAF Flight Dynamics Laboratory, “Advanced Technology 

Program: Technical Development Plan for Aerothermodynamic/Elastic Structural Systems Environmental Tests (ASSET)” (Sept. 1963), pp. 1–5.

58. Aviation Week (May 24, 1965), p. 62; McDonnell, “ASSET ASV-3 Flight Test Report,” Report 

B251 (65FD-234) (Jan. 4, 1965).

59. McDonnell, “ASSET ASV-4 Flight Test Report,” Report B707 (65FD-938) (June 25, 1965), p. 

156; Hallion, “ASSET,” in Hallion, ed., Hypersonic Revolution, vol. 1, p. 519.

60. USAF Flight Dynamics Laboratory, “ASSET Final Briefing,” Report 65FD-850 (Oct. 5, 1965).
NASA’s Contributions to Aeronautics

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TABLE 2

MCDONNELL ASSET FLIGHT TEST PROGRAM

DATE VEHICLE BOOSTER VELOCITY 

(FEET/

SECOND)

ALTITUDE 

(FEET)

RANGE 

(NAUTICAL 

MILES)

Sept. 18, 

1963

ASV-1 Thor 16,000 205,000 987

Mar. 24, 

1964

ASV-2 Thor-Delta 18,000 195,000 1,800

July 22, 

1964

ASV-3 Thor-Delta 19,500 225,000 1,830

Oct. 27, 

1964

AEV-1 Thor 13,000 168,000 830

Dec. 8, 

1964

AEV-2 Thor 13,000 187,000 620

Feb. 23, 

1965

ASV-4 Thor-Delta 19,500 206,000 2,300

Hypersonic Aerothermodynamic Protection and the Space Shuttle

Certainly over much of the Shuttle’s early conceptual period, advocates 

thought such logistical transatmospheric aerospace craft would employ 

hot structure thermal protection. But undertaking such structures on 

large airliner-size vehicles proved troublesome and thus premature. 

Then, as though given a gift, NASA learned that Lockheed had built a 

pilot plant and could mass-produce silica “tiles” that could be attached 

to a conventional aluminum structure, an approach far more appealing 

than designing a hot structure. Accordingly, when the Agency undertook development of the Space Shuttle in the 1970s, it selected this 

approach, meaning that the new Shuttle was, in effect, a simple aluminum airplane. Not surprisingly, Lockheed received a NASA subcontract 

in 1973 for the Shuttle’s thermal-protection system.

Lockheed had begun its work more than a decade earlier, when 

investigators at Lockheed Missiles and Space began studying ceramic 

fiber mats, filing a patent on the technology in December 1960. Key 

people included R.M. Beasley, Ronald Banas, Douglas Izu, and Wilson 

Schramm. By 1965, subsequent Lockheed work had led to LI-1500, a 

material that was 89 percent porous and weighed 15 pounds per cubic 

foot (lb/ft3). Thicknesses of no more than an inch protected test surfaces during simulations of reentry heating. LI-1500 used methyl methacrylate (Plexiglas), which volatilized when hot, producing an outward 
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flow of cool gas that protected the heat shield, though also compromising its reusability.61

Lockheed’s work coincided with NASA plans in 1965 to build a space 

station as is main post-Apollo venture and, consequently, the first great 

wave of interest in designing practical logistical Shuttle-like spacecraft 

to fly between Earth and the orbital stations. These typically were conceived as large winged two-stage-to-orbit systems with fly-back boosters 

and orbital spaceplanes. Lockheed’s Maxwell Hunter devised an influential design, the Star Clipper, with two expendable propellant tanks and 

LI-1500 thermal protection.62 The Star Clipper also was large enough 

to benefit from the Allen-Eggers blunt-body principle, which lowered 

its temperatures and heating rates during reentry. This made it possible to dispense with the outgassing impregnant, permitting use—and, 

more importantly, reuse—of unfilled LI-1500. Lockheed also introduced 

LI-900, a variant of LI-1500 with a porosity of 93 percent and a weight 

of only 9 pounds per cubic foot. As insulation, both LI-900 and LI-1500 

were astonishing. Laboratory personnel found that they could heat a 

tile in a furnace until it was white hot, remove it, allow its surface to 

cool for a couple of minutes, and pick it up at its edges with their fingers, with its interior still glowing at white heat.63

61. Paul Cooper and Paul F. Holloway, “The Shuttle Tile Story,” Astronautics & Aeronautics, vol. 19, 

no. 1 (Jan. 1981), pp. 24–34; Wilson B. Schramm, Ronald P. Banas, and Y. Douglas Izu, “Space 

Shuttle Tile—The Early Lockheed Years,” Lockheed Horizons, Issue 13 (1983), pp. 2–15; T.A. Heppenheimer, The Space Shuttle Decision, SP-4221 (Washington, DC: NASA, 1999).

62. Lockheed Missiles and Space Corporation (LMSC), “Space Transport and Recovery System 

(Space Shuttle),” LMSC A946332 (Mar. 1969), Shuttle Historical Documents Collection, N. 

SHHDC-0048, NASA Marshall Space Flight Center; LMSC, “Final Report: Integral Launch and ReEntry Vehicle,” LMSC A959837 (Dec. 1969), Center for Aerospace Information 70N-31831.

63. Richard C. Thuss, Harry G. Thibault, and Arnold Hiltz, “The Utilization of Silica Based Surface 

Insulation for the Space Shuttle Thermal Protection System,” SAMPE National Technical Conference on 

Space Shuttle Materials, Huntsville, AL (Oct. 1971), pp. 453–464, Center for Aerospace Information 

72A-10764; Schramm, et al., “Space Shuttle Tile”; L.J. Korb, C.A. Morant, R.M. Calland, and C.S. 

Thatcher, “The Shuttle Orbiter Thermal Protection System,” and Wilson Schramm, “HRSI and LRSI—

The Early Years,” both in American Ceramic Society Bulletin, vol. 60 (1981), pp. 1188–1195; L.J. 

Graham, F.E. Sugg, and W. Gonzalez, “Nondestructive Evaluation of Space Shuttle tiles,” Ceramic 

Engineering and Science Proceedings, vol. 3 (1982), pp. 680–697; Robert L. Dotts, Donald M. 

Curry, and Donald L. Tillian, “Orbiter Thermal Protection System,” and William C. Schneider and Glenn 

J. Miller, “The Changing ‘Scales of the Bird’ (Shuttle Tile Structural Integrity),” in Norman Chaffee, ed., 

“Space Shuttle Technical Conference,” NASA Conference Publication 2343 (1983).
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Previous company work had amounted to general materials 

research. But Lockheed now understood in 1971 that NASA wished to 

build the Shuttle without simultaneously proceeding with the station, 

opening a strong possibility that the company could participate. The 

program had started with a Phase A preliminary study effort, advancing 

then to Phase B, which was much more detailed. Hot structures were initially ascendant but posed serious challenges, as NASA Langley researchers found when they tried to build a columbium heat shield suitable for 

the Shuttle. The exercise showed that despite the promise of reusability 

and long life, coatings were fragile and damaged easily, leading to rapid 

oxygen-induced embrittlement at high temperatures. Unprotected columbium oxidized particularly readily and, when hot, could burst into flame. 

Other refractory metals were available, but they were little understood 

because they had been used mostly in turbine blades.

Even titanium amounted literally to a black art. Only one firm, 

Lockheed, had significant experience with a titanium hot structure. 

That experience came from the Central Intelligence Agency-sponsored 

Blackbird strategic reconnaissance program, so most of the pertinent 

shop-floor experience was classified. The aerospace community knew 

that Lockheed had experienced serious difficulties in learning how to 

work with titanium, which for the Shuttle amounted to an open invitation to difficulties, delays, and cost overruns.

The complexity of a hot structure—with large numbers of clips, 

brackets, standoffs, frames, beams, and fasteners—also militated 

against its use. Each of the many panel geometries needed their own 

structural analysis that was to show with confidence that the panel 

could withstand creep, buckling, flutter, or stress under load, and in 

the early computer era, this posed daunting analytical challenges. Hot 

structures were also known generally to have little tolerance for “overtemps,” in which temperatures exceeded the structure’s design point.64

Thus, having taken a long look at hot structures, NASA embraced 

the new Lockheed pilot plant and gave close examination to Shuttle 

designs that used tiles, which were formally called Reusable Surface 

Installation (RSI). Again, the choice of hot structures versus RSI 

reflected the deep pockets of the Air Force, for hot structures were 

64. Korb, et al., “Shuttle Orbiter TPS”; L. J. Korb and H. M. Clancy, “The Shuttle Thermal Protection 

System—A Material and Structural Overview,” SAMPE 26th National Symposium, Los Angeles, CA 

(Apr. 1981), pp. 232–249 (Center for Aerospace Information 81A-44344).
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costly and complex. But RSI was inexpensive, flexible, and simple. 

It suited NASA’s budget while hot structures did not, so the Agency 

chose it.

In January 1972, President Richard M. Nixon approved the Shuttle 

as a program, thereby raising it to the level of a Presidential initiative. 

Within days, Dale Myers, a senior official, announced that NASA had 

made the basic decision to use RSI. The North American Rockwell concept that won the $2.6 billion prime contract in July therefore specified 

RSI as well—but not Lockheed’s. North American Rockwell’s version 

came from General Electric and was made from mullite.65

Which was better, the version from GE or the one from Lockheed? 

Only tests would tell—and exposure to temperature cycles of 2,300 °F 

gave Lockheed a clear advantage. NASA then added acoustic tests that 

simulated the loud roars of rocket flight. This led to a “sudden-death 

shootout,” in which competing tiles went into single arrays at NASA 

Johnson. After 20 cycles, only Lockheed’s entrants remained intact. 

In separate tests, Lockheed’s LI-1500 withstood 100 cycles to 2,500 °F 

and survived a thermal overshoot to 3,000 °F as well as an acoustic 

overshoot to 174 decibels (dB).

Lockheed won the thermal-protection subcontract in 1973, with 

NASA specifying LI-900 as the baseline RSI. The firm responded by preparing to move beyond the pilot-plant level and to construct a full-scale 

production facility in Sunnyvale, CA. With this, tiles entered the mainstream of thermal protection systems available for spacecraft design, 

in much the same way that blunt bodies and ablative approaches had 

before them, first flying into space aboard the Space Shuttle Columbia 

in April 1981. But getting them operational and into space was 

far from easy.66

65. “NASA Space Shuttle Technology Conference,” vol. 2: “Structures and Materials,” NASA TM-X2273 (1971); Heppenheimer, The Space Shuttle Decision, pp. 341–346.

66. Schramm, et al., “Space Shuttle Tile.” See also Donald H. Humes, “Hypervelocity Impact Tests 

on Space Shuttle Orbiter Thermal Protection Material,” NASA TM-X-74039 (1977); M.J. Suppans 

and C.J. Schroeder, “Space Shuttle Orbiter Thermal Protection Development and Verification Test 

Program,” AIAA Paper 78-485 (1978); R. Jeffrey Smith, “Shuttle Problems Compromise Space 

Program,” Science (Nov. 23, 1979), pp. 910–912, 914; Mitch Waldrop, “Space Shuttle Tiles: 

A Question of Bonding,” Chemical and Engineering News, vol. 58 (May 12, 1980), pp. 27–29; 

W.C. Rochelle, et al., “Orbiter TPS Development and Certification testing at the NASA/JSC 10 MW 

Atmospheric Reentry Materials and Structures Evaluation Facility,” AIAA Paper 83-0147 (1983).
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The Tiles Become Operational

Manufacture of the silica tiles was straightforward, at least in its basic 

steps. The raw material consisted of short lengths of silica fiber of l.0-micron 

diameter. A measured quantity of fibers, mixed with water, formed a slurry. 

The water was drained away, and workers added a binder of colloidal silica, 

then pressed the material into rectangular blocks that were 10 to 20 inches 

in diameter and more than 6 inches thick. These blocks were the crudest 

form of LI-900, the basic choice of RSI for the entire Shuttle. They sat for 

3 hours to allow the binder to jell, then were dried thoroughly in a microwave oven. The blocks moved through sintering kilns that baked them at 

2,375 °F for 2 hours, fusing binder and fibers together. Band saws trimmed 

distortions from the blocks, which were cut into cubes and then carved into 

individual tiles using milling machines driven by computer. The programs 

contained data from Rockwell International on the desired tile dimensions.

Next, the tiles were given a spray-on coating. After being oven-dried, 

they returned to the kilns for glazing at temperatures of 2,200 °F for 90 

minutes. To verify that the tiles had received the proper amount of coating, technicians weighed samples before and after the coating and glazing. The glazed tiles then were made waterproof by vacuum deposition of 

a silicon compound from Dow Corning while being held in a furnace at 

350 °F. These tiles were given finishing touches before being loaded into 

arrays for final milling.67

Although the basic LI-900 material showed its merits during 1972, 

it was another matter to produce it in quantity, to manufacture tiles that 

were suitable for operational use, and to provide effective coatings. To 

avoid having to purify raw fibers from Johns Manville, Lockheed asked 

that company to find a natural source of silica sand with the necessary 

purity. The amount needed was small, about 20 truckloads, and was not 

of great interest to quarry operators. Nevertheless, Johns Manville found 

a suitable source in Minnesota.

Problems arose when shaping the finished tiles. Initial plans called 

for a large number of identical flat tiles, varying only in thickness and 

trimmed to fit at the time of installation. But flat tiles on the curved surface of the Shuttle produced a faceted surface that promoted the onset 

of turbulence in the airflow, resulting in higher rates of heating. The tiles 

67. Richard G. O’Lone, “Thermal Tile Production Ready to Roll,” Aviation Week (Nov. 8, 1976), 

pp. 51–54; L.J. Korb, C.A. Morant, R.M. Calland, and C.S. Thatcher, “The Shuttle Orbiter Thermal 

Protection System,” American Ceramic Society Bulletin, vol. 60, 1981, pp. 1188–1193.
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then would have had to be thicker, which threatened to add weight. The 

alternative was an external RSI contour closely matching that of the orbiter’s outer surface. Lockheed expected to produce 34,000 tiles for each 

orbiter, grouping most of them in arrays of two dozen or so and machining their back faces, away from the glazed coating, to curves matching 

the contours of the Shuttle’s aluminum skin. Each of the many thousands of tiles was to be individually numbered, and none had precisely 

the same dimensions. Instead, each was defined by its own set of dimensions. This cost money, but it saved weight.

Difficulties also arose in the development of coatings. The first good 

one, LI-0042, was a borosilicate glass that used silicon carbide to enhance 

its high-temperature thermal emissivity. It dated to the late 1960s; a variant, LI-0050, initially was the choice for operational use. This coating 

easily withstood the rated temperature of 2,300 °F, but in tests, it persistently developed hairline cracks after 20 to 60 thermal cycles. This was 

unacceptable; it had to stand up to 100 such cycles. The cracks were too 

small to see with the unaided eye and did not grow large or cause tile 

failure. But they would have allowed rainstorms to penetrate the tiles 

during the weeks that an orbiter was on the ground between missions, 

with the rain adding to the launch weight. Help came from NASA Ames, 

where researchers were close to Lockheed, both in their shared interests 

and in their facilities being only a few miles apart. Howard Goldstein 

at Ames, a colleague of the branch chief, Howard Larson, set up a task 

group and brought in a consultant from Stanford University, which also 

was just up the road. They spent less than $100,000 in direct costs and 

came up with a new and superior coating called reaction-cured glass. 

Like LI-0050, it was a borosilicate, consisting of more than 90 percent 

silica along with boria or boron oxide along with an emittance agent. 

The agent in LI-0050 had been silicon carbide; the new one was silicon 

tetraboride, SiB4. During glazing, it reacted with silica in a way that 

increased the level of boria, which played a critical role in controlling 

the coating’s thermal expansion. This coating could be glazed at lower 

temperature than LI-0050 could, reducing the residual stress that led to 

the cracking. SiB4 oxidized during reentry, but in doing so, it produced 

boria and silica, the ingredients of the glass coating itself.68

68. L.J. Korb and H.M. Clancy, “Symposium on Reusable Surface Insulation for Space Shuttle,” vol. 

1, NASA TM-X-2719 (1973), pp. 14–15; “The Shuttle Thermal Protection System—A Material and 

Structural Overview,” Apr. 1981, pp. 232–249, CASI 81A-44344. 
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The Shuttle’s distinctive mix of black-and-white tiles was all designed 

as standard LI-900 with its borosilicate coating, but the black ones had 

SiB4 and the white ones did not. Still, they all lacked structural strength 

and were brittle. They could not be bonded directly to the orbiter’s aluminum skin, for they would fracture and break because of their inability to 

follow the flexing of this skin under its loads. Designers therefore placed 

an intermediate layer between tiles and skin, called a strain isolator pad 

(SIP). It was a felt made of Nomex nylon from DuPont, which would neither melt nor burn. It had useful elasticity and could stretch in response 

to Shuttle skin flexing without transmitting excessive strain to the tiles.69

Testing of tiles and other thermal-protection components continued 

through the 1970s, with NASA Ames being particularly active. A particular challenge lay in creating turbulent flows, which demanded close study 

because they increased the heat-transfer rates many times over. During 

reentry, hypersonic flow over a wing is laminar near the leading edge, transitioning to turbulence at some distance to the rear. No hypersonic wind 

tunnel could accommodate anything resembling a full-scale wing, and 

it took considerable power as well as a strong airflow to produce turbulence in the available facilities. Ames had a 60-megawatt arc-jet, but even 

that facility could not accomplish this. Ames succeeded in producing such 

flows by using a 20-megawatt arc-jet that fed its flow into a duct that was 

9 inches across and 2 inches deep. The narrow depth gave a compressed 

flow that readily produced turbulence, while the test chamber was large 

enough to accommodate panels with size of 8 by 20 inches. This facility supported the study of coatings that led to the use of reaction-cured 

glass. Tiles of LI-900, 6 inches square and treated with this coating, survived 100 simulated reentries at 2,300 °F in turbulent flow.70

The Ames 20-megawatt arc-jet facility made its own contribution in 

a separate program that improved the basic silica tile. Excessive temperatures caused these tiles to fail by shrinking and becoming denser. 

69. David H. Greenshields, “Orbiter Thermal Protection System Development” (Apr. 1977), 

pp. 1-28–1-42, CASI 77A-35304; North American-Rockwell, “Space Shuttle System Summary 

Briefing,” Report SV 72-19 (July 8, 1972); Korb, et al., “Shuttle”; Robert M. Powers, Shuttle: The 

World’s First Spaceship (Harrisburg, PA: Stackpole Books, 1979), p. 241.

70. Frank Kreith, Principles of Heat Transfer (Scranton, PA: International Textbook Co., 1965), pp. 

534–538; Benjamin M. Elson, “New Unit to Test Shuttle Thermal Guard,” Aviation Week (Mar. 

31, 1975), pp. 52–53; H.K. Larson and H.E. Goldstein, “Space Shuttle Orbiter Thermal Protection 

Material Development and Testing,” (Mar. 1978), pp. 189–194, CASI 79A-17673.
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Investigators succeeded in reducing the shrinkage by raising the tile density 

and adding silicon carbide to the silica, rendering it opaque and reducing 

internal heat transfer. This led to a new grade of silica RSI with density of 22 

lb/ft3 that had greater strength as well as improved thermal performance.71

The Ames researchers carried through with this work during 1974 

and 1975, with Lockheed taking this material and putting it into production as LI-2200. Its method of manufacture largely followed that of standard LI-900, but whereas that material relied on sintered colloidal silica 

to bind the fibers together, LI-2200 dispensed with this and depended 

entirely on fiber-to-fiber sintering. LI-2200 was adopted in 1977 for operational use on the Shuttle, where it found application in specialized areas. 

These included regions of high concentrated heat near penetrations such 

as landing-gear doors as well as near interfaces with the carbon-carbon 

nose cap, where surface temperatures could reach 2,600 °F.72

Testing proceeded in four overlapping phases. Material selection ran 

through 1973 and 1974 into 1975; the work that led to LI-2200 was an 

example. Material characterization proceeded concurrently and extended 

midway through 1976. Design development tests covered 1974 through 

1977; design verification activity began in 1977 and ran through subsequent years. Materials characterization called for some 10,000 test specimens, with investigators using statistical methods to determine basic 

material properties. These were not the well-defined properties that engineers find listed in handbooks; they showed ranges of values that often 

formed a Gaussian distribution, with its bell-shaped curve. This activity 

addressed such issues as the lifetime of a given material, the effects of 

changes in processing, or the residual strength after a given number of 

flights. A related topic was simple but far-reaching: to be able to calculate the minimum tile thickness, at a given location, that would hold the 

skin temperature below the maximum allowable.73

Design development tests used only 350 articles but spanned 4 

years, because each of them required close attention. An important 

goal involved validating the specific engineering solutions to a number 

71. CASI 79A-17673; CASI 81A-44344.

72. Korb, et al., “Shuttle”; Elizabeth A. Muenger, Searching the Horizon: A History of Ames 

Research Center, 1940–1976, NASA SP-4304 (Washington, DC: NASA, 1985).

73. Aviation Week (Mar. 31, 1975), pp. 52–53; CASI 81A-44344; Gregory P. McIntosh and 

Thomas P. Larkin, “The Space Shuttle’s Testing Gauntlet,” Astronautics and Aeronautics (Jan. 1976), 

pp. 60, 62–64.
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of individual thermal-protection problems. Thus the nose cap and wing 

leading edges were made of carbon-carbon, in anticipation of their being 

subjected to the highest temperatures. Their attachments were exercised 

in structural tests that simulated flight loads up to design limits, with 

design temperature gradients.

Design development testing also addressed basic questions of the tiles 

themselves. There were narrow gaps between them, and while Rockwell 

had ways to fill them, these gap-fillers required their own trials by fire. 

A related question was frequently asked: What happens if a tile falls off? 

A test program addressed this and found that in some areas of intense 

heating, the aluminum skin indeed would burn through. The only way 

to prevent this was to be sure that the tiles were firmly bonded in place, 

and this meant all those located in critical areas.74

Design verification tests used fewer than 50 articles, but these represented substantial portions of the vehicle. An important test article, 

evaluated at NASA Johnson, reproduced a wing leading edge and measured 5 by 8 feet. It had two leading-edge panels of carbon-carbon set 

side by side, a section of wing structure that included its principal spars, 

and aluminum skin covered with RSI. It could not have been fabricated 

earlier in the program, for its detailed design drew on lessons from 

previous tests. It withstood simulated air loads, launch acoustics, and 

mission-temperature-pressure environments, not once, but many times.75

The testing ranged beyond the principal concerns of aerodynamics, 

heating, and acoustics. There also was concern that meteoroids might 

not only put craters in the carbon-carbon but also cause it to crack. At 

NASA Langley, the researcher Donald Humes studied this by shooting small glass and nylon spheres at target samples using a light-gas 

gun driven by compressed helium. Helium is better than gunpowder, 

as it can expand at much higher velocities. Humes wrote that carboncarbon: “does not have the penetration resistance of the metals on a thickness basis, but on a weight basis, that is, mass per unit area required to 

stop projectiles, it is superior to steel.”76

74. McIntosh and Larkin, “Space Shuttle’s Testing Gauntlet,” pp. 60, 62–64; Aviation Week (Mar. 

31, 1975), p. 52; M.J. Suppans and C.J. Schroeder, “Space Shuttle Orbiter Thermal Protection 

Material,” AIAA Paper 78-485 (1978).

75. McIntosh and Larkin, “Space Shuttle’s Testing Gauntlet,” pp. 60, 62–64.

76. Donald H. Humes, “Hypervelocity Impact Tests on Space Shuttle Orbiter Thermal Protection 

Material,” NASA TM X-74039 (1977), p. 12.
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Yet amid the advanced technology of arc-jets, light-gas guns, and 

hypersonic wind tunnels, one of the most important tests was also one 

of the simplest. It involved nothing more than taking tiles that were 

bonded with adhesive to the SIP and the underlying aluminum skin and 

physically pulling them off.

It was no new thing for people to show concern that the tiles might 

not stick. In 1974, a researcher at Ames noted that aerodynamic noise 

was potentially destructive, telling a reporter for Aviation Week that: 

“We’d hate to shake them all off when we’re leaving.” At NASA Johnson, 

a 10-MW arc-jet saw extensive use in lost-tile investigations. Tests indicated there was reason to believe that the forces acting to pull off a tile 

would be as low as 2 psi, just some 70 pounds for a tile measuring 6 by 6 

inches square. This was low indeed; the adhesive, SIP, and RSI material 

all were considerably stronger. The thermal-protection testing therefore 

had given priority to thermal rather than to mechanical work, essentially taking it for granted that the tiles would stay on. Thus, attachment 

of the tiles to the Shuttle lacked adequate structural analysis, failing to 

take into account the peculiarities in the components. For example, the 

SIP had some fibers oriented perpendicular to the cemented tile undersurface. The tile was made of ceramic fibers, with these fibers concentrating the loads. This meant that the actual stresses they faced were 

substantially greater than anticipated.77

Columbia orbiter OV-102 was the first to receive working tiles. 

Columbia was also slated to be first into space. It underwent final assembly at the Rockwell plant in Palmdale, CA, during 1978. Checkout of 

onboard systems began in September, and installation of tiles proceeded 

concurrently, with Columbia to be rolled out in February 1979. But 

mounting the tiles was not at all like laying bricks. Measured gaps were to 

separate them; near the front of the orbiter, they had to be positioned to 

within 0.17 inches of vertical tolerance to form a smooth surface that 

77. W.C. Rochelle, et al., “Orbiter TPS Development and Certification Testing at the NASA/

JSC 10 MW Atmospheric Reentry Materials and Structures Evaluation Facility,” AIAA Paper 83-

0147 (1983); Richard G. O’Lone, “Shuttle Test Pace Intensifies at Ames,” Aviation Week (June 

24, 1974), p. 71; Mitch Waldrop, “Space Shuttle Tiles: A Question of Bonding,” Chemical and 

Engineering News, vol. 58 (May 12, 1980), pp. 27–29; Paul A. Cooper and Paul F. Holloway, 

“Shuttle Tile Story,” pp. 27, 29; William C. Schneider and Glenn J. Miller, “The Challenging 

‘Scales of the Bird’ (Shuttle Tile Structural Integrity)” in Norman Chaffee, ed., “Space Shuttle Technical Conference,” NASA Conference Publication CP-2342 (1983), pp. 403–404.
NASA’s Contributions to Aeronautics

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would not trip the airflow into turbulence. This would not have been 

difficult if the tiles had rested directly on the aluminum skin, but they 

were separated from that skin by the spongy SIP. The tiles were also fragile. An accidental tap with a wrench, a hard hat, even a key chain could 

crack the glassy coating. When that happened, the damaged tile had to be 

removed and the process of installation had to start again with a new one.78

The tiles came in arrays, each array numbering about three-dozen 

tiles. It took 1,092 arrays to cover this orbiter, and NASA reached a high 

mark when technicians installed 41 of them in a single week. But unfortunate news came midway through 1979 as detailed studies showed that 

in many areas the combined loads due to aerodynamic pressure, vibration, and acoustics would produce excessively large forces on the tiles. 

Work to date had treated a 2-psi level as part of normal testing, but now 

it was clear that only a small proportion of the tiles already installed 

faced stresses that low. Over 5,000 tiles faced force levels of 8.5 to 13 

psi, with 3,000 being in the range of 2 to 6.5 psi. The usefulness of tiles 

as thermal protection was suddenly in doubt.79

What caused this? The fault lay in the nylon felt SIP, which had 

been modified by “needling” to increase its through-the-thickness tensile 

strength and elasticity. This was accomplished by punching a barbed needle through the felt fabric, some 1,000 times per square inch, which oriented fiber bundles transversely to the SIP pad. Tensile loads applied across 

the SIP pad, acting to pull off a tile, were transmitted into the SIP at discrete regions along these transverse fibers. This created localized stress 

concentrations, where the stresses approached twice the mean value. These 

local areas failed readily under load, causing the glued bond to break.80

There also was a clear need to increase the strength of the tiles’ 

adhesive bonds. The solution came during October and involved modifying a thin layer at the bottom of each tile to make it denser. The process was called, quite logically, “densification.” It used DuPont’s Ludox 

78. “First Shuttle Launch Vehicle Being Assembled at Palmdale,” Aviation Week (Nov. 27, 1978), 

p. 64; Waldrop, “Tiles”; Craig Covault, “Thermal Tile Application Accelerated,” Aviation Week 

(May 21, 1979), pp. 59–63.

79. Aviation Week (Nov. 27, 1978), p. 64; Waldrop, “Tiles”; Craig Covault, “Administration 

Backs Shuttle Fund Rise,” Aviation Week (Sept. 17, 1979), pp. 22–23.

80. L.J. Graham, F.E. Sugg, and W. Gonzalez, “Nondestructive Evaluation of Space Shuttle Tiles,” 

Ceramic Engineering and Science Proceedings, vol. 3, (1982), pp. 681–683; Schneider and 

Miller, “Challenging”; Astronautics and Aeronautics (Jan. 1981), p. 29.
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with a silica “slip.” Ludox was colloidal silica stirred into water and 

stabilized with ammonia; the slip had fine silica particles dispersed in 

water. The Ludox acted like cement; the slip provided reinforcement, 

in the manner of sand in concrete. It worked: the densification process 

clearly restored the lost strength.81

By then, Columbia had been moved to the Kennedy Space Center. 

The work nevertheless went badly during 1979, for as people continued 

to install new tiles, they found more and more that needed to be removed 

and replaced. Orderly installation procedures broke down. Rockwell 

had received the tiles from Lockheed in arrays and had attached them 

in well-defined sequences. Even so, that work had gone slowly, with 550 

tiles in a week being a good job. But now Columbia showed a patchwork 

of good ones, bad ones, and open areas with no tiles. Each individual 

tile had been shaped to a predetermined pattern at Lockheed using that 

firm’s numerically controlled milling machines. But the haphazardness 

of the layout made it likely that any precut tile would fail to fit into its 

assigned cavity, leaving too wide a gap with the adjacent ones.

Many tiles therefore were installed one by one, in a time-consuming 

process that fitted two into place and then carefully measured space for 

a third, designing it to fill the space between them. The measurements 

went to Sunnyvale, CA, where Lockheed carved that tile to its unique 

specification and shipped it to the Kennedy Space Center (KSC). Hence, 

each person took as long as 3 weeks to install just 4 tiles. Densification 

also took time; a tile removed from Columbia for rework needed 2 weeks 

until it was ready for reinstallation.82

How could these problems have been avoided? They all stemmed 

from the fact that the tile work was well advanced before NASA learned 

that the tile-SIP-adhesive bonds had less strength than the Agency 

needed. The analysis that disclosed the strength requirements was neither costly nor demanding; it might readily have been in hand during 

1976 or 1977. Had this happened, Lockheed could have begun shipping 

densified tiles at an early date. Their development and installation would 

have occurred within the normal flow of the Shuttle program, with the 

change amounting perhaps to little more than an engineering detail.

81. “Densification Process Applied to Shuttle Tiles,” Aviation Week (Feb. 25, 1980), p. 22; Astronautics and Aeronautics (Jan. 1981), pp. 29–30.

82. Aviation Week (Feb. 25, 1980), pp. 22–24; Craig Covault, “Mated Shuttle Reaches Pad 

39,” Aviation Week (May 7, 1979), p. 14.
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The Space Shuttle Columbia descends to land at Edwards following its hypersonic reentry from 

orbit in April 1981. NASA.

The reason this did not happen was far-reaching, for it stemmed 

from the basic nature of the program. The Shuttle effort followed “concurrent development,” with design, manufacture, and testing proceeding in parallel rather than in sequence. This approach carried risk, but 

the Air Force had used it with success during the 1960s. It allowed new 

technologies to enter service at the earliest possible date. But within 

the Shuttle program, funds were tight. Managers had to allocate their 

budgets adroitly, setting priorities and deferring what they could put 

off. To do this properly was a high art, calling for much experience and 

judgment, for program executives had to be able to conclude that the 

low-priority action items would contain no unpleasant surprises. The calculation of tile strength requirements was low on the action list because 

it appeared unnecessary; there was good reason to believe that the tiles 

would face nothing worse than 2 psi. Had this been true, and had the 

main engines been ready, Columbia might have flown by mid-1980. It 

did not fly until April 1981, and, in this sense, tile problems brought a 

delay of close to 1 year.

The delay in carrying through the tile-strength computation was not 

mandatory. Had there been good reason to upgrade its priority, it could 

readily have been done earlier. The budget stringency that brought this 
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deferral (along with many others) thus was false economy par excellence, for the program did not halt during that year of launch delay. It 

kept writing checks for its contractors and employees. The missing tilestrength analysis thus ramified in its consequences, contributing substantially to a cost overrun in the Shuttle program.83

During 1979, NASA gave the same intense level of attention to the 

tiles’ mechanical problems that it had previously reserved for their thermal development. The effort nevertheless continued to follow the pattern of three steps forward and two steps back, and, for a while, more 

tiles were removed than were put on in a given week. Even so, by the 

fall of 1980, the end was in sight.84

During the spring of 1979, before the main tile problems had come to 

light, the schedule had called for the complete assembly of Columbia, with 

its external tank and solid boosters, to take place on November 24, 1979. 

Exactly 1 year later, a tow vehicle pulled Columbia into the Vehicle Assembly 

Building as a large crowd watched and cheered. Within 2 days, Columbia was 

mounted to its tank, forming a live Shuttle in flight configuration. Kenneth 

Kleinknecht, an X-series and space flight veteran and now Shuttle manager at NASA Johnson, put it succinctly: “The vehicle is ready to launch.”85

Shuttle Aerodynamics and Structures

The Shuttle was one of the last major aircraft to rely almost entirely on 

wind tunnels for studies of its aerodynamics. There was much interest 

in an alternative: the use of supercomputers to derive aerodynamic data 

through solution of the governing equations of airflow, known as the 

Navier-Stokes equations. Solution of the complete equations was out of 

the question, for they carried the complete physics of turbulence, with 

turbulent eddies that spanned a range of sizes covering several orders of 

magnitude. But during the 1970s, investigators made headway by dropping the terms within these equations that contained viscosity, thereby 

suppressing turbulence.86

83. Waldrop, “Tiles.”

84. “Shuttle Engine, Tile Work Proceeding on Schedule,” Aviation Week (Sept. 15, 1980), p. 26.

85. “NASA Finishes Shuttle Mating,” Aviation Week (Dec. 1, 1980), pp. 18–19; Aviation Week

(Sept. 17, 1979), p. 22; “NASA Presses to Hold Tight Shuttle Schedule,” Aviation Week (Aug. 4, 

1980), p. 24.

86. John D. Anderson, A History of Aerodynamics (New York: Cambridge University Press, 1997), 

pp. 441–443.
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People pursued numerical simulation because it offered hope of 

overcoming the limitations of wind tunnels. Such facilities usually tested 

small models that failed to capture important details of the aerodynamics of full-scale aircraft. Other errors arose from tunnel walls and model 

supports. Hypersonic flight brought its own restrictions. No installation 

had the power to accommodate a large model, realistic in size, at the 

velocity and temperatures of reentry.87

By piecing together results from specialized facilities, it was possible to gain insights into flows at near-orbital speeds. The Shuttle reentered at Mach 27. NASA Langley had a pair of wind tunnels that used 

helium, which expands to very high flow velocities. These attained Mach 

20, Mach 26, and even Mach 50. But their test models were only a few 

inches in size, and their flows were very cold and could not duplicate the 

high temperatures of atmosphere entry. Shock tunnels, which heated and 

compressed air using shock waves, gave true temperature up to Mach 

17 while accommodating somewhat larger models. Yet their flow durations were measured in milliseconds.88

During the 1970s, the largest commercially available mainframe 

computers included the Control Data 7600 and the IBM 370-195.89 These 

sufficed to treat complete aircraft—but only at the lowest level of approximation, which used linearized equations and treated the airflow over an 

airplane as a small disturbance within a uniform free stream. The full 

Navier-Stokes equations contained 60 partial derivatives; the linearized 

approximation retained only 3 of these terms. It nevertheless gave good 

accuracy in computing lift, successfully treating such complex configurations as a Shuttle orbiter mated to its 747. The next level of approximation 

restored the most important nonlinear terms and treated transonic and 

hypersonic flows, which were particularly difficult to simulate in wind 

tunnels. The inadequacies of wind tunnel work had brought such errors 

as faulty predictions of the location of shock waves along the wings of 

the C-141, an Air Force transport. In flight test, this plane tended to nose 

downward, and its design had to be modified at considerable expense.

87. Dean R. Chapman, Hans Mark, and Melvin W. Pirtle, “Computers vs. Wind Tunnels for Aerodynamic Flow Simulations,” Astronautics and Aeronautics (Apr. 1975), p. 26.

88. T.A. Heppenheimer, Hypersonic Technologies and the National Aerospace Plane (Arlington, 

VA: Pasha Publications, 1990), pp. 128–134.

89. William D. Metz, “Midwest Computer Architect Struggles with the Speed of Light,” Science 

(Jan. 27, 1978), pp. 404–405.
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Computers such as the 7600 could not treat complete aircraft in 

transonic flow, for the equations were more complex and the computation requirements more severe. HiMAT, a highly maneuverable NASA 

experimental aircraft, flew at Dryden and showed excessive drag at 

Mach 0.9. Redesign of its wing used a transonic-flow computational code 

and approached the design point. The same program, used to reshape 

the wing of the Grumman Gulfstream, gave considerable increases 

in range and fuel economy while reducing the takeoff distance and 

landing speed.90

During the 1970s, NASA’s most powerful computer was the Illiac IV, 

at Ames Research Center. It used parallel processing and had 64 processing units, achieving speeds up to 25 million operations per second. Built 

by Burroughs Corporation with support from the Pentagon, this machine 

was one of a kind. It entered service at Ames in 1973 and soon showed 

that it could run flow-simulation codes an order of magnitude more rapidly than a 7600. Indeed, its performance foreshadowed the Cray-1, a 

true supercomputer that became commercially available only after 1976.

The Illiac IV was a research tool, not an instrument of mainstream 

Shuttle development. It extended the reach of flow codes, treating threedimensional inviscid problems while supporting simulations of viscous 

flows that used approximate equations to model the turbulence.91 In the 

realm of Space Shuttle studies, Ames’s Walter Reinhardt used it to run a 

three-dimensional inviscid code that included equations of atmospheric 

chemistry. Near-peak-entry heating of the Shuttle would be surrounded 

by dissociated air that was chemically reacting and not in chemical equilibrium. Reinhardt’s code treated the full-scale orbiter during entry and 

gave a fine example of the computational simulation of flows that were 

impossible to reproduce in ground facilities.92

90. Gina Bari Kolata, “Who Will Build the Next Supercomputer?” Science, Jan. 16, 1981, pp. 

268–269; Randolph A. Graves, Jr., “Computational Fluid Dynamics: The Coming Revolution,” 

Astronautics and Aeronautics (Mar. 1982), pp. 20–28; Astronautics and Aeronautics (Apr. 1975), 

pp. 22–30. For numbers of partial derivatives, see Dean R. Chapman, “Computational Aerodynamics Development and Outlook,” AIAA Journal, vol. 17, No. 12 (1979), p. 1294.

91. Science, Jan. 16, 1981, pp. 268–269; Benjamin B. Elson, “Computer Seen Assuming Shuttle 

Tasks.” Aviation Week (Sept. 3, 1973), pp. 14–16; D.L. Slotnick, “The Fastest Computer,” Scientific American (Feb. 1971), pp. 76–87.

92. Walter A. Reinhardt, “Parallel Computation of Unsteady, Three-Dimensional, Chemically 

Reacting, Nonequilibrium Flow Using a Time-Split Finite Volume Method on the Illiac IV,” Journal of 

Physical Chemistry, vol. 81, no. 25 (1977), pp. 2427–2435.
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Such exercises gave tantalizing hints of what would be done with 

computers of the next generation. Still, the Shuttle program was at least 

a decade too early to use computational simulations both routinely 

and effectively. NASA therefore used its wind tunnels. The wind tunnel program gave close attention to low-speed flight, which included 

approach and landing as well as separation from the 747 during the 

1977 flight tests of Enterprise.

In 1975, Rockwell built a $1 million model of the orbiter at 0.36 

scale, lemon yellow in color and marked with the blue NASA logo. 

It went into the 40- by 80-foot test section of Ames’s largest tunnel, 

which was easily visible from the adjacent freeway. It gave parameters for the astronauts’ flight simulators, which previously had used 

data from models at 3-percent scale. The big one had grooves in its 

surface that simulated the gaps between thermal protection tiles, 

permitting assessment of the consequences of the resulting roughness of the skin. It calibrated and tested systems for making aerodynamic measurements during flight test and verified the design of the 

elevons and other flight control surfaces as well as of their actuators.93

Other wind tunnel work strongly influenced design changes that 

occurred early in development. The most important was the introduction of the lightweight delta wing late in 1972, which reduced the size 

of the solid boosters and chopped 1 million pounds from the overall 

weight. Additional results changed the front of the external tank from 

a cone to an ogive and moved the solid boosters rearward, placing their 

nozzles farther from the orbiter. The modifications reduced drag, minimized aerodynamic interference on the orbiter, and increased stability 

by moving the aerodynamic center aft.

The activity disclosed and addressed problems that initially had not 

been known to exist. Because both the liquid main engines and the solids had nozzles that gimbaled, it was clear that they had enough power 

to provide control during ascent. Aerodynamic control would not be 

necessary, and managers believed that the orbiter could set its elevons 

in a single position through the entire flight to orbit. But work in wind 

tunnels subsequently showed that aerodynamic forces during ascent 

would impose excessive loads on the wings. This required elevons to 

move while in powered flight to relieve these loads. Uncertainties in the 

93. Richard G. O’Lone, “Tunnel Tests Yield New Orbiter Data,” Aviation Week (June 30, 1975), 

pp. 43–44.
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wind tunnel data then broadened this requirement to incorporate an 

active system that prevented overloading the elevon actuators. This system also helped the Shuttle to fly a variety of ascent trajectories, which 

imposed different elevon loads from one flight to the next.94

Much wind tunnel work involved issues of separation: Enterprise 

from its carrier aircraft, solid boosters from the external tank after 

burnout. At NASA Ames, a 14-foot transonic tunnel investigated problems of Enterprise and its 747. Using the same equipment, engineers 

addressed the separation of an orbiter from its external tank. This was 

supposed to occur in near-vacuum, but it posed aerodynamic problems 

during an abort.

The solid boosters brought their own special issues and nuances. 

They had to separate cleanly; under no circumstances could a heavy steel 

casing strike a wing. Small solid rocket motors, mounted fore and aft on 

each booster, were to push them away safely. It then was necessary to 

understand the behavior of their exhaust plumes, for these small motors 

were to blast into onrushing airflow that could blow their plumes against 

the orbiter’s sensitive tiles or the delicate aluminum skin of the external tank. Wind tunnel tests helped to define appropriate angles of fire 

while also showing that a short, sharp burst from the motors was best.95

Prior to the first orbital flight in 1981, the program racked up 46,000 

wind tunnel hours. This consisted of 24,900 hours for the orbiter, 17,200 

for the mated launch configuration, and 3,900 for the carrier aircraft 

program. During the 9 years from contract award to first flight, this 

was equivalent to operating a facility 16 hours a day, 6 days a week. 

Specialized projects demanded unusual effort, such as an ongoing 

attempt to minimize model-to-model and tunnel-to-tunnel discrepancies. This work alone conducted 28 test series and used 14 wind tunnels.96

Structural tests complemented the work in aerodynamics. The 

mathematics of structural analysis was well developed, with computer 

94. Charlie C. Dill, et al., “The Space Shuttle Ascent Vehicle Aerodynamic Challenges: Configuration Design and Data Base Development,” in Norman Chaffee, ed., “Space Shuttle Technical 

Conference,” NASA CP-2342, (1983), pp. 151–152, 161.

95. Richard G. O’Lone, “Shuttle Task Pace Intensifies at Ames,” Aviation Week (June 24, 1974), p. 

71; Craig Covault, “Thermal, Weight Concerns Force Changes to Shuttle,” Aviation Week (Dec. 9, 

1974), p. 19; Dill, et al., “Ascent,” pp. 154, 165.

96. James C. Young, et al., “The Aerodynamic Challenges of the Design and Development of the 

Space Shuttle Orbiter,” in Chaffee, “Conference,” pp. 217–220.
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programs called NASTRAN that dealt with strength under load while 

addressing issues of vibration, bending, and flexing. The equations of 

NASTRAN were linear and algebraic, which meant that in principle they 

were easy to solve. The problem was that there were too many of them, 

for the most detailed mathematical model of the orbiter’s structure had 

some 50,000 degrees of freedom. Analysts introduced abridged versions 

that cut this number to 1,000 and then relied on experimental tests for 

data that could be compared with the predictions of the computers.97

There were numerous modes of vibration, with frequencies that 

changed as the Shuttle burned its propellants. Knowledge of these frequencies was essential, particularly in dealing with “pogo.” This involved 

a longitudinal oscillation like that of a pogo stick, with propellant flowing in periodic surges within its main feed line. Such surges arose when 

their frequency matched that of one of the structural modes, producing 

resonance. The consequent variations in propellant-flow rate then caused 

the engine thrust to oscillate at that same rate. This turned the engines 

into sledgehammers, striking the vehicle structure at its resonant frequency, and made the pogo stronger. It weakened only when consumption of propellant brought a further change in the structural frequency 

that broke the resonance, allowing the surges to die out.

Pogo was common; it had been present on earlier launch vehicles. 

It had brought vibrations with acceleration of 9 g’s in a Titan II, which 

was unacceptably severe. Engineering changes cut this to below 0.25 

g, which enabled this rocket to launch the manned Gemini spacecraft. 

Pogo reappeared in Apollo during the flight of a test Saturn V in 1968. 

For the Shuttle, the cure was relatively simple, calling for installation 

of a gas-filled accumulator within the main oxygen line. This damped 

the pogo oscillations, though design of this accumulator called for close 

understanding of the pertinent frequencies.98

97. C. Thomas Modlin, Jr., and George A. Zupp, Jr., “Shuttle Structural Dynamics Characteristics—The 

Analysis and Verification,” in Chaffee, “Conference,” p. 326; “COSMIC Software Catalog,” NASA 

CR-191005 (1993).

98. Dennis Jenkins, Space Shuttle: The History of the National Space Transportation System (Stillwater, MN: Voyageur Press), 2001, p. 416; James M. Grimwood, Barton C. Hacker, and Peter J. 

Vorzimmer, Project Gemini Technology and Operations: A Chronology, NASA SP-4002 (Washington, DC: NASA, 1969), pp. 68, 76, 121; Roger E. Bilstein, Stages to Saturn: A Technological 

History of the Apollo/Saturn Launch Vehicles, NASA SP-4206 (Washington, DC: NASA, 1980), 

pp. 360, 362–363.
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The most important structural tests used actual flight hardware, 

including the orbiter Enterprise and STA-099, a full-size test article 

that later became the Challenger. In 1978, Enterprise went to NASA 

Marshall, where the work now included studies on the external tank. 

For vibrational tests, engineers assembled a complete Shuttle by mating Enterprise to such a tank and to a pair of dummy solid boosters. 

One problem that these models addressed came at lift-off. The ignition 

of the three main engines imposes a sudden load of more than 1 million 

pounds of thrust. This force bends the solid boosters, placing considerable stress at their forward attachments to the tank. If the solid boosters were to ignite at that moment, their thrust would add to the stress.

To reduce the force on the attachment, analysts took advantage of 

the fact that the solid boosters would not only bend but would sway 

back and forth somewhat slowly, like an upright fishing rod. The strain 

on the attachment would increase and decrease with the sway, and it 

was possible to have the solid boosters ignite at an instant of minimum 

load. This called for delaying their ignition by 2.7 seconds, which cut 

the total load by 25 percent. The main engines fired during this interval, which consumed propellant, cutting the payload by 600 pounds. 

Still, this was acceptable.99

While Enterprise underwent vibration tests, STA-099 showed the 

orbiter’s structural strength by standing up to applied forces. Like a 

newborn baby that lacks hair, this nascent form of Challenger had no 

thermal-protection tiles. Built of aluminum, it looked like a large fighter 

plane. For the structural tests, tiles were not only unnecessary; they were 

counterproductive. The tiles had no structural strength of their own that 

had to be taken into account, and they would have received severe damage from the hydraulic jacks that applied the loads and forces.

STA-099 and Columbia had both been designed to accommodate a set 

of loads defined by a database designated 5.1. In 1978, there was a new 

database, 5.4, and STA-099 had to withstand its loads without acquiring strains or deformations that would render it unfit for flight. Yet in 

an important respect, this vehicle was untestable; it was not possible to 

validate the strength of its structural design merely by applying loads 

with those jacks. The Shuttle structure had evolved under such strong 

emphasis on saving weight that it was necessary to take full account 

99. Modlin and Zupp, “Structural,” p. 326; Alden C. Mackey and Ralph E. Gotto, “Structural Load 

Challenges During Space Shuttle Development,” in Chaffee, “Conference,” pp. 335–339.
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of thermal stresses that resulted from temperature differences across 

structural elements during reentry. No facility existed that could impose 

thermal stresses on so large an object as STA-099, for that would have 

required heating the entire vehicle.

STA-099 and Columbia had both been designed to withstand ultimate loads 140 percent greater than those of the 5.1 database. The 

structural tests on STA-099 now had to validate this safety factor for 

the new 5.4 database. Unfortunately, a test to 140 percent of the 5.4 

loads threatened to produce permanent deformations in the structure. 

This was unacceptable, for STA-099 was slated for refurbishment into 

Challenger. Moreover, because thermal stresses could not be reproduced 

over the entire vehicle, a test to 140 percent would sacrifice the prospect of building Challenger while still leaving questions as to whether 

an orbiter could meet the safety factor of 140 percent.

NASA managers shaped the tests accordingly. For the entire vehicle, they used the jacks to apply stresses only up to 120 percent of the 

5.4 loads. When the observed strains proved to match closely the values 

predicted by stress analysis, the 140 percent safety factor was deemed 

to be validated. In addition, the forward fuselage underwent the most 

severe aerodynamic heating, yet it was relatively small. It was subjected 

to a combination of thermal and mechanical loads that simulated the 

complete reentry stress environment in at least this limited region. STA099 then was given a detailed and well-documented posttest inspection. 

After these tests, STA-099 was readied as the flight vehicle Challenger, 

joining Columbia as part of NASA’s growing Shuttle fleet.100

Aerospaceplane to NASP: The Lure of Air-Breathing Hypersonics

The Space Shuttle represented a rocket-lofted approach to hypersonic 

space access. But rockets were not the only means of propulsion contemplated for hypersonic vehicles. One of the most important aspects 

of hypersonic evolution since the 1950s has been the development of 

the supersonic combustion ramjet, popularly known as a scramjet. The 

ramjet in its simplest form is a tube and nozzle, into which air is introduced, mixed with fuel, and ignited, the combustion products passing 

100. Craig Covault, “NASA Evaluating Major Shuttle Orbiter Changes,” Aviation Week (Oct. 10, 

1977), p. 26; Photo of Structural Test Article, Aviation Week, Mar. 6, 1978, p. 13; Philip C. Glynn 

and Thomas L. Moser, “Orbital Structural Design and Verification,” in Chaffee, “Conference,” pp. 

353–356.
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through a classic nozzle and propelling the engine forward. Unlike 

a conventional gas turbine, the ramjet does not have a compressor 

wheel or staged compressor blades, cannot typically function at speeds 

less than Mach 0.5, and does not come into its own until the inlet 

velocity is near or greater than the speed of sound. Then it functions remarkably well as an accelerator, to speeds well in excess of 

Mach 3.

Conventional subsonic-combustion ramjets, as employed by the 

Mach 4.31 X-7, held promise as hypersonic accelerators for a time, but 

they could not approach higher hypersonic speeds because their subsonic internal airflow heated excessively at high Mach. If a ramjet could 

be designed that had a supersonic internal flow, it would run much 

cooler and at the same time be able to accelerate a vehicle to doubledigit hypersonic Mach numbers, perhaps reaching the magic Mach 25, 

signifying orbital velocity. Such an engine would be a scramjet. Such 

engines have only recently made their first flights, but they nevertheless 

are important in hypersonics and point the way toward future practical 

air-breathing hypersonics.

An important concern explored at the NACA’s Lewis Flight 

Propulsion Laboratory during the 1950s was whether it was possible 

to achieve supersonic combustion without producing attendant shock 

waves that slow internal flow and heat it. Investigators Irving Pinkel 

and John Serafini proposed experiments in supersonic combustion under a supersonic wing, postulating that this might afford a 

means of furnishing additional lift. Lewis researchers also studied 

supersonic combustion testing in wind tunnels. Supersonic tunnels 

produced very low air pressure, but it was known that aluminum borohydride could promote the ignition of pentane fuel even at pressures 

as low as 0.03 atmospheres. In 1955, Robert Dorsch and Edward 

Fletcher successfully demonstrated such tunnel combustion, and subsequent research indicated that combustion more than doubled lift at 

Mach 3.

Though encouraging, this work involved flow near a wing, not in a 

ramjet-like duct. Even so, NACA aerodynamicists Richard Weber and 

John MacKay posited that shock-free flow in a supersonic duct could 

be attained, publishing the first open-literature discussion of theoretical 

scramjet performance in 1958, which concluded: “the trends developed 

herein indicate that the [scramjet] will provide superior performance 
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at higher hypersonic flight speeds.”101 The Weber-MacKay study came 

a year after Marquardt researchers had demonstrated supersonic combustion of a hydrogen and air mix. Other investigators working contemporaneously were the manager William Avery and the experimentalist 

Frederick Billig, who independently achieved supersonic combustion 

at the Johns Hopkins University Applied Physics Laboratory (APL), and 

J. Arthur Nicholls at the University of Michigan.102

The most influential of all scramjet advocates was the colorful Italian 

aerodynamicist, partisan leader, and wartime emigree, Antonio Ferri. 

Before the war, as a young military engineer, he had directed supersonic 

wind tunnel studies at Guidonia, Benito Mussolini’s showcase aeronautical research establishment outside Rome. In 1943, after the collapse of 

the Fascist regime and the Nazi assumption of power, he left Guidonia, 

leading a notably successful band of anti-Nazi, anti-Fascist partisans. 

Brought to America by Moe Berg, a baseball player turned intelligence 

agent, Ferri joined NACA Langley, becoming Director of its Gas Dynamics 

Branch. Turning to the academic world, he secured a professorship at 

Brooklyn Polytechnic Institute. He formed a close association with 

101. Richard J. Weber and John S. MacKay, “An Analysis of Ramjet Engines Using Supersonic 

Combustion,” NACA TN-4386 (1958), p. 22; Irving Pinkel and John S. Serafini, “Graphical 

Method for Obtaining Flow Field in Two-Dimensional Supersonic Stream to Which Heat is Added,” 

NACA TN-2206 (1950); Irving Pinkel, John S. Serafini, and John L. Gregg, “Pressure Distribution 

and Aerodynamic Coefficients Associated with Heat Addition to Supersonic Air Stream Adjacent to 

Two-Dimensional Supersonic Wing,” NACA RM-E51K26 (1952). 

102. Alan Newman, “Speed Ahead of its Time,” Johns Hopkins Magazine, vol. 40, no. 4 (Dec. 

1988), pp. 26–31; Harold E. Gilreath, “The Beginning of Hypersonic Ramjet Research at APL,” Johns 

Hopkins Applied Physics Laboratory Technical Digest, vol. 11, no. 3-4 (1990), pp. 319–335; G.L. 

Dugger, F.S. Billig, and W.H. Avery, “Hypersonic Propulsion Studies at the Applied Physics Laboratory, 

The Johns Hopkins University,” Johns Hopkins University Applied Physics Laboratory Report TG 405 (June 

14, 1961), esp. pp. 1–3; Frank D. Stull, Robert A. Jones, and William P. Zima, “Propulsion Concepts 

for High Speed Aircraft,” Paper 75-1092, Society of Automotive Engineers (1975); Paul J. Waltrup, 

Griffin Y. Anderson, and Frank D. Stull, “Supersonic Combustion Ramjet (Scramjet) Engine Development 

in the United States,” Johns Hopkins University Applied Physics Laboratory Paper 76-042 (1976); Paul 

J. Waltrup, “Liquid Fueled Supersonic Combustion Ramjets: A Research Perspective of the Past, Present 

and Future,” AIAA Paper 86-0158 (1986); Paul J. Waltrup, “Hypersonic Airbreathing Propulsion: Evolution and Opportunities,” in Advisory Group for Aeronautical Research and Development, Conference 

Proceedings on the Aerodynamics of Hypersonic Lifting Vehicles (Neuilly sur Seine, France: NATOAGARD, 1987), pp. 1–29; and in Thomas C. Adamson, Jr., “Aeronautical and Aerospace Engineering 

Education at the University of Michigan,” in Barnes McCormick, et al., Aerospace Engineering Education 

During the First Century of Flight (Reston, VA: AIAA, 2004), p. 54.
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Alexander Kartveli, chief designer at Republic Aviation, and designer of 

the P-47, F-84, XF-103, and F-105. Indeed, Kartveli’s XF-103 (which, alas, 

never was completed or flown) employed a Ferri engine concept. In 1956, 

he established General Applied Science Laboratories (GASL), with financial backing from the Rockefellers.103

Ferri emphasized that scramjets could offer sustained performance far 

higher than rockets could, and his strong reputation ensured that people 

listened to him. At a time when shock-free flow in a duct still loomed as a 

major problem, Ferri did not flinch from it but instead took it as a point 

of departure. He declared in September 1958 that he had achieved it, thus 

taking a position midway between the demonstrations at Marquardt and 

APL. Because he was well known, he therefore turned the scramjet from 

a wish into an invention, which might be made practical.

He presented his thoughts publicly at a technical colloquium in Milan 

in 1960 (“Many of the older men present,” John Becker wrote subsequently, 

“were politely skeptical”) and went on to give a far more detailed discussion in May 1964, at the Royal Aeronautical Society in London. This was 

the first extensive public presentation on hypersonic propulsion, and 

the attendees responded with enthusiasm. One declared that whereas 

investigators “had been thinking of how high in flight speed they could 

stretch conventional subsonic burning engines, it was now clear that they 

should be thinking of how far down they could stretch supersonic burning 

engines,” and another added that Ferri now was “assailing the field which 

until recently was regarded as the undisputed regime of the rocket.”104

103. Edward T. Curran, “Scramjet Engines: The First Forty Years,” Journal of Propulsion and Power,

vol. 17, No. 6 (Nov.–Dec. 2001), pp. 1138–1148; Antonio Ferri, “Review of Scramjet Technology,” Journal of Aircraft, vol. 5, no. 1 (Jan. 1968), pp. 3–10; T.A. Heppenheimer, Facing the Heat 

Barrier: A History of Hypersonics, SP-2007-4232 (Washington, DC: NASA, 2007), p. 103; R.R. 

Jamison, “Hypersonic Air Breathing Engines,” in A.R. Collar and J. Tinkler, Hypersonic Flow: Proceedings of the Eleventh Symposium of the Colston Research Society held in the University of Bristol, Apr. 

6–8, 1959 (London: Butterworths Scientific Publications, 1960), pp. 391–408; S.W. Greenwood, 

“Spaceplane Propulsion,” The Aeroplane and Astronautics (May 25, 1961), pp. 597–599.

104. John V. Becker, “Confronting Scramjet: The NASA Hypersonic Ramjet Experiment,” in Richard 

P. Hallion, ed., The Hypersonic Revolution: Eight Case Studies in the History of Hypersonic Technology, vol. 2: From Scramjet to the National Aero-Space Plane, 1964–1986 (Wright-Patterson AFB: 

Aeronautical Systems Division, 1987), p. 752; Antonio Ferri, “Review of Problems in Application of 

Supersonic Combustion,” 7th Lanchester Memorial Lecture, Journal of the Royal Aeronautical Society, 

vol. 68, no. 645 (Sept. 1964), pp. 595, 597; Antonio Ferri, “Supersonic Combustion Progress,” 

Astronautics & Aeronautics, vol. 2, no. 8 (Aug. 1964), pp. 32–37; Heppenheimer, Facing the 

Heat Barrier, pp. 104–105. 
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Scramjet advocates were offered their first opportunity to actually build such propulsion systems with the Air Force’s abortive 

Aerospaceplane program of the late 1950s–mid-1960s. A contemporary 

to Dyna-Soar but far less practical, Aerospaceplane was a bold yet premature effort to produce a logistical transatmospheric vehicle and possible orbital strike system. Conceived in 1957 and initially known as the 

Recoverable Orbital Launch System (ROLS), Aerospaceplane attracted 

surprising interest from industry. Seventeen aerospace companies submitted contract proposals and related studies; Convair, Lockheed, and 

Republic submitted detailed designs. The Republic concept had the 

greatest degree of engine-airframe integration, a legacy of Ferri’s partnership with Kartveli.

By the early 1960s, Aerospaceplane not surprisingly was beset with 

numerous developmental problems, along with a continued debate over 

whether it should be a single– or two-stage system, and what proportion 

of its propulsion should be turbine, scramjet, and pure rocket. Though 

it briefly outlasted Dyna-Soar, it met the same harsh fate. In the fall of 

1963, the Air Force Scientific Advisory Board damned the program in 

no uncertain terms, noting: “Aerospaceplane has had such an erratic 

history, has involved so many clearly infeasible factors, and has been 

subjected to so much ridicule that from now on this name should be 

dropped. It is recommended that the Air Force increase [its] vigilance 

[so] that no new program achieves such a difficult position.”105 The next 

year, Congress slashed its remaining funding, and Aerospaceplane was 

at last consigned to a merciful oblivion.

In the wake of Aerospaceplane’s cancellation, both the Air Force 

and NASA maintained an interest in advancing scramjet propulsion for 

transatmospheric aircraft. The Navy’s scramjet interest, though great, 

was primarily in smaller engines for missile applications. But Air Force 

and NASA partisans formed an Ad-Hoc Working Group on Hypersonic 

Scramjet Aircraft Technology.

105. Air Force Scientific Advisory Board, “Memo-Report of the USAF Scientific Advisory Board Aerospace Vehicles/Propulsion Panels on Aerospaceplane, VTOL, and Strategic Manned Aircraft” (Oct. 

24, 1963), pp. 1, 3, SAB Office files, USAF HQ, Pentagon, Washington, DC; see also F.E. Jariett 

and G. Karel, “Aerospaceplane: The Payload Capabilities of Various Recoverable Systems All Using 

Hydrogen Fuel,” General Dynamics Astronautics Report AE62-0892 (Oct. 25, 1962), esp. pp. 8–23; 

“Aerospaceplane May be a Two-Stage Vehicle,” Aviation Week & Space Technology (July 22, 1963), 

pp. 245–249; Heppenheimer, Facing the Heat Barrier, pp. 112–128.
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Both agencies pursued development programs that sought to 

build and test small scramjet modules. The Air Force Aero-Propulsion 

Laboratory sponsored development of an Incremental Scramjet flighttest program at Marquardt. This proposed test vehicle underwent extensive analysis and study, though without actually flying as a functioning 

scramjet testbed. The first manifestation of Langley work was the socalled Hypersonic Research Engine (HRE), an axisymmetric scramjet 

of circular cross section with a simple Oswatitsch spike inlet, designed 

by Anthony duPont. Garrett AiResearch built this engine, planned for a 

derivative of the X-15. The HRE never actually flew as a “hot” functioning engine, though the X-15A-2 flew repeatedly with a boilerplate test 

article mounted on the stub ventral fin (during its record flight to Mach 

6.70 on October 3, 1967, searing hypersonic shock interactions melted it 

off the plane). Subsequent tunnel tests revealed that the HRE was, unfortunately, the wrong design. A podded and axisymmetric design, like an airliner’s fanjet, it could only capture a small fraction of the air that flowed 

past a vehicle, resulting in greatly reduced thrust. Integrating the scramjet with the airframe, so that it used the forebody to assist inlet performance and the afterbody as a nozzle enhancement, would more than 

double its thrust.106

Investigation of such concepts began at Langley in 1968, with pioneering studies by researchers John Henry, Shimer Pinckney, and others. Their work expanded upon a largely Ferri-inspired base, defining 

what emerged as common basic elements of subsequent Langley scramjet research. It included a strong emphasis upon airframe integration, 

use of fixed geometry, a swept inlet that could readily spill excess airflow, and the use of struts for fuel injection. Early observations, published in 1970, showed that struts were practical for a large supersonic 

combustor at Mach 8. The program went on to construct test scramjets 

and conducted almost 1,000 wind tunnel test runs of engines at Mach 4 

106. M.L. Brown and R.L. Maxwell, Marquardt Corporation, “Scramjet Incremental Flight Test Program,” Summary, Marquardt-AF Aero-Propulsion Laboratory Report AFAPLTR-67-112 (1968), pp. 

3–4; Becker, “Confronting Scramjet,” in Hallion, ed., Hypersonic Revolution, vol. 2, pp. 747–861, 

examines the HRE in detail; John R. Henry and Griffin Y. Anderson, “Design Considerations for the 

Airframe-Integrated Scramjet,” NASA TM-X-2895 (1973); Robert A. Jones and Paul W. Huber, 

“Toward Scramjet Aircraft,” Astronautics and Aeronautics, vol. 16, no. 2 (Feb. 1978), pp. 38–48; 

G. Burton Northam and G.Y. Anderson, “Supersonic Combustion Ramjet Research at Langley,” 

AIAA Paper 86-0159 (1986).
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and Mach 7. Inlets at Mach 4 proved sensitive to “unstarts,” a condition 

where the shock wave is displaced, disrupting airflow and essentially starving the engine of its oxygen. Flight at Mach 7 raised the 

question of whether fuel could mix and burn in the short available combustor length.107

Langley test engines, like engines at GASL, Marquardt, and other 

scramjet research organizations, encountered numerous difficulties. 

Large disparities existed between predicted performance and that actually achieved in the laboratory. Indeed, the scramjet, advanced so boldly 

in the mid-1950s, would not be ready for serious pursuit as a propulsive 

element until 1986. Then, on the eve of the National Aerospace Plane 

development program, Langley researchers Burton Northam and Griffin 

Anderson announced that NASA had succeeded at last in developing a practical scramjet. They proclaimed triumphantly: “At both Mach 4 and Mach 7 

flight conditions, there is ample thrust both for acceleration and cruise.”108

Out of such optimism sprang the National Aero-Space Plane program, 

which became a central feature of the presidency of Ronald Reagan. It was 

linked to other Reagan-era defense initiatives, particularly his Strategic 

Defense Initiative, a ballistic missile shield intended to reduce the threat 

of nuclear war, which critics caustically belittled as “Star Wars.” SDI called 

for the large-scale deployment of defensive arms in space, and it became 

clear that the Space Shuttle would not be their carrier. Experience since 

the Shuttle’s first launch in April 1981 had shown that it was costly and 

took a long time to prepare for relaunch. The Air Force was unwilling 

to place the national eggs in such a basket. In February 1984, Defense 

Secretary Caspar Weinberger approved a document stating that total reliance upon the Shuttle represented an unacceptable risk.

107. NASA Langley completed 963 successful runs of three strut, parametric, and step-strut 

scramjets between 1976–1987; from Edward G. Ruf, “Airframe-Integrated Scramjet Engine tests in 

NASA Langley Scramjet Engine Test Facilities,” at http://hapb-www.larc.nasa.gov/Public/

Engines/engine_tests.html, accessed on May 1, 2009.

108. Northam and Anderson, “Supersonic Combustion Ramjet Research,” AIAA Paper 86-0159, 

p. 7; for NASA and other work, see J. Menzler and T.W. Mertz, “Large Scale Supersonic Combustor Testing at Conditions Simulating Mach 8 Flight,” AIAA Paper 70-715 (1970); Carl A. Trexler, 

“Inlet Performance of the Integrated Langley Scramjet Module (Mach 2.3 to 7.6),” AIAA Paper 

75-1212 (1975); Robert W. Guy and Ernest A. Mackley, “Initial Wind Tunnel Tests at Mach 4 and 

7 of a Hydrogen-Burning, Airframe Integrated Scramjet,” AIAA Paper 79-7045 (1979); and R.C. 

Rogers, D.P. Capriotti, and R.W. Guy, “Experimental Supersonic Combustion Research at NASA 

Langley,” AIAA Paper 2506 (1998).
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An Air Force initiative was under way at the time that looked toward 

an alternative. Gen. Lawrence A. Skantze, Chief of Air Force Systems 

Command (AFSC), had sponsored studies of Trans Atmospheric Vehicles 

(TAVs) by Air Force Aeronautical Systems Division (ASD). These reflected 

concepts advanced by ASD’s chief planner, Stanley A. Tremaine, as well 

as interest from Air Force Space Division (SD), the Defense Advanced 

Research Projects Agency (DARPA), and Boeing and other companies. 

TAVs were SSTO craft intended to use the Space Shuttle Main Engine 

(SSME) and possibly would be air-launched from derivatives of the 

Boeing 747 or Lockheed C-5. In August 1982, ASD had hosted a 3-day 

conference on TAVs, attended by representatives from AFSC’s Space 

Division and DARPA. In December 1984, ASD went further. It established a TAV Program Office to “streamline activities related to longterm, preconceptual design studies.”109

DARPA’s participation was not surprising, for Robert Cooper, head of 

this research agency, had elected to put new money into ramjet research. 

His decision opened a timely opportunity for Anthony duPont, who had 

designed the HRE for NASA. DuPont held a strong interest in “combinedcycle engines” that might function as a turbine air breather, translate 

to ram/scram, and then perhaps use some sophisticated air collection 

and liquefaction process to enable them to boost as rockets into orbit. 

There are several types of these engines, and duPont had patented such 

a design as early as 1972. A decade later, he still believed in it, and he 

learned that Anthony Tether was the DARPA representative who had 

been attending TAV meetings.

Tether sent him to Cooper, who introduced him to DARPA aerodynamicist Robert Williams, who brought in Arthur Thomas, who had been 

studying scramjet-powered spaceplanes as early as Sputnik. Out of this 

climate of growing interest came a $5.5 million DARPA study program, 

Copper Canyon. Its results were so encouraging that DARPA took the 

notion of an air-breathing single-stage-to-orbit vehicle to Presidential 

science adviser George Keyworth and other senior officials, including Air Force Systems Command’s Gen. Skantze. As Thomas recalled: 

“The people were amazed at the component efficiencies that had been 

109. USAF Aeronautical Systems Division news release 84-211; Richard P. Hallion, “Yesterday, 

Today, and Tomorrow: From Shuttle to the National Aero-Space Plane,” in Hallion, ed., Hypersonic 

Revolution, vol. 2, pp. 1336–1337, 1361; S.A. Tremaine and Jerry B. Arnett, “Transatmospheric 

Vehicles—A Challenge for the Next Century,” AIAA Paper 84-2414 (1984).
NASA’s Contributions to Aeronautics

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The National Aero-Space Plane concept in final form, showing its modified lifting body design 

approach. NASA.

assumed in the study. They got me aside and asked if I really believed 

it. Were these things achievable? Tony [duPont] was optimistic everywhere: on mass fraction, on drag of the vehicle, on inlet performance, 

on nozzle performance, on combustor performance. The whole thing, 

across the board. But what salved our consciences was that even if these 

things weren’t all achieved, we still could have something worthwhile. 

Whatever we got would still be exciting.”110

Gen. Skantze realized that SDI needed something better than the 

Shuttle—and Copper Canyon could possibly be it. Briefings were encouraging, but he needed to see technical proof. That evidence came when 

he visited GASL and witnessed a subscale duPont engine in operation. 

Afterward, as DARPA’s Bob Williams recalled subsequently: “the Air Force 

system began to move with the speed of a spaceplane.”111 Secretary of 

Defense Caspar Weinberger received a briefing and put his support behind 

the effort. In January 1986, the Air Force established a joint-service Air 

Force-Navy-NASA National Aero-Space Plane Joint Program Office at 

Aeronautical Systems Division, transferring into it all the personnel 

110. Quote from Heppenheimer interview with Arthur Thomas, Sept. 24, 1987.

111. Quote from Heppenheimer interview with Robert Williams, May 1, 1986.
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previously assigned to the TAV Program Office established previously. 

(The program soon received an X-series designation, as the X-30.) Then 

came the clincher. President Ronald Reagan announced his support for 

what he now called the “Orient Express” in his State of the Union Address 

to the Nation on February 4, 1986. President Reagan’s support was not 

the product of some casual whim: the previous spring, he had ordered 

a joint Department of Defense-NASA space launch study of future space 

needs and, additionally, established a national space commission. Both 

strongly endorsed “aerospace plane development,” the space commission recommending it be given “the highest national priority.”112

Though advocates of NASP attempted to sharply differentiate their 

effort from that of the discredited Aerospaceplane of the 1960s, the NASP 

effort shared some distressing commonality with its predecessor, particularly an exuberant and increasingly unwarranted optimism that afforded 

ample opportunity for the program to run into difficulties. In 1984, with 

optimism at its height, DARPA’s Cooper declared that the X-30 could be 

ready in 3 years. DuPont, closer to the technology, estimated that the 

Government could build a 50,000-pound fighter-size vehicle in 5 years 

for $5 billion. Such predictions proved wildly off the mark. As early as 

1986, the “Government baseline” estimate of the aircraft rose to 80,000 

pounds. Six years later, in 1992, its gross weight had risen eightfold, to 

400,000 pounds. It also had a “velocity deficit” of 3,000 feet per second, 

meaning that it could not possibly attain orbit. By the next year, NASP 

“lay on life support.”113

It had evolved from a small, seductively streamlined speedster to a 

fatter and far less appealing shape more akin to a wooden shoe, entering a death spiral along the way. It lacked performance, so it needed 

greater power and fuel, which made it bigger, which meant it lacked performance so that it needed greater power and fuel, which made it bigger 

. . . and bigger . . . and bigger. X-30 could never attain the “design closure” permitting it to reach orbit. NASP’s support continuously softened, 

112. National Commission on Space, Pioneering the Space Frontier (New York: Bantam Books, 

1986), p. 184; Joint DOD–NASA Task Team Report, National Space Transportation and Support 

Study (Washington, DC: GPO, 1986); President Ronald Reagan, State of the Union Address, Feb. 

4, 1986.

113. Larry Schweikart, The Quest for the Orbital Jet: The National Aero-Space Plane Program 

(1983–1995), vol. 3 of Hallion, ed., The Hypersonic Revolution (Washington, DC: USAF, 1998), 

pp. 349, 279–351.
NASA’s Contributions to Aeronautics

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particularly as technical challenges rose, performance estimates fell, and 

other national issues grew in prominence. It finally withered in the mid1990s, leaving unresolved what, if anything, scramjets might achieve.114

Transatmospherics after NASP

Two developments have paced work in hypersonics since NASP died in 

1995. Continuing advances in computers, aided markedly by advancements in wind tunnels, have brought forth computational fluid dynamics (CFD). Today, CFD simulates the aerodynamics of flight vehicles 

with increasing (though not perfect) fidelity. In addition, NASA and the 

Air Force have pursued a sequence of projects that now aim clearly at 

developing operational scramjet-powered military systems.

Early in the NASP effort, in 1984, Robert Whitehead of the Office 

of Naval Research spoke on CFD to its people. Robert Williams recalls 

that Williams presented the equations of fluid dynamics “so the computer could solve them, then showed that the computer technology was 

also there. We realized that we could compute our way to Mach 25 with 

high confidence.”115 Unfortunately, in reality, DARPA could not do that. 

In 1987, the trade journal Aerospace America reported: “almost nothing is known about the effects of heat transfer, pressure gradient, threedimensionality, chemical reactions, shock waves, and other influences on 

hypersonic transition.”116 (This transition causes a flow to change from 

laminar to turbulent, a matter of fundamental importance.)

114. Stuart O. Schmitt, Theodore J. Wierzbanowski, and Johnny Johnson, “The Challenge of X-30 

Flight Test,” 31st Symposium, Society of Experimental Test Pilots, Beverly Hills, CA, Sept. 26, 1987; 

United States General Accounting Office, “National Aero-Space Plane: A Technology Development 

and Demonstration Program to Build the X-30,” Report GAO/NSIAD-88-122 (Apr. 1988); Alan 

W. Wilhite, et al., “Concepts Leading to the National Aero-Space Plane Program,” AIAA Paper 

90-0294 (1990); Robert B. Barthelemy, “The National Aero Space Plane Program: A Revolutionary 

Concept,” Johns Hopkins Applied Physics Laboratory Technical Digest, vol. 11, no. 2 & 3 (1990), pp. 

312–318; United States General Accounting Office, “National Aero-Space Plane: Key Issues Facing 

the Program,” Report GAO/T-NSIAD-92-26 (Mar. 1992), pp. 4–15; Joseph F. Shea, et al., “Report of 

the Defense Science Board Task Force on National Aero-Space Plane (NASP) Program” (1992); United 

States General Accounting Office, National Aero-Space Plane: Restructuring Future Research and Development Efforts, Report GAO/NSIAD-93-71 (Dec. 1992), p. 4; and Ray L. Chase and Ming H. Tang, 

“A History of the NASP Program from the Formation of the Joint Program Office to the Termination of the 

HySTP Scramjet Performance Demonstration Program,” AIAA Paper 95-6031 (1995).

115. Author interview with Robert Williams, May 1, 1986.

116. Quote from Douglas L. Dwoyer, Paul Kutler, and Louis A. Povinelli, “Retooling CFD for Hypersonic Aircraft,” Aerospace America (Oct. 1987), p. 35.
Case 5 | Toward Transatmospheric Flight: From V-2 to the X-51

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Code development did mature so that it could adequately support 

the next hypersonic system, NASA’s X-43A program. In supporting the 

X-43A effort, NASA’s most important code was GASP. NASP had used 

version 2.0; the X-43A used 3.0.117 Like any flow code, it could not calculate the turbulence directly but had to model it. GASP 3.0 used the 

Baldwin-Lomax algebraic model that Princeton’s Antony Jameson, a 

leading writer of flow codes, describes as: “the most popular model in 

the industry, primarily because it’s easy to program.”118 GASP 3.0 also 

uses “eddy-viscosity” models, which Stanford’s Peter Bradshaw rejects 

out of hand: “Eddy viscosity does not even deserve to be described as a 

‘theory’ of turbulence!” More broadly, he adds, “Even the most sophisticated turbulence models are based on brutal simplifications” of the 

pertinent nonlinear partial differential equations.119

Can increasing computer power make up for this? Calculations of 

the NASP era had been rated in gigaflops, billions of floating point operations per second (FLOPS).120 An IBM computer has recently cracked 

the petaflop mark—at a quadrillion operations per second, and even 

greater performance is being contemplated.121 At Stanford University’s 

Center for Turbulence Research, analyst Krishnan Mahesh studied flow 

within a commercial turbojet and found a mean pressure drop that differs from the observed value by only 2 percent. An earlier computation had given an error of 26 percent, an order of magnitude higher.122 

He used Large Eddy Simulation, which calculates the larger turbulent 

eddies and models the small ones that have a more universal character. But John Anderson, a historian of fluid dynamics, notes that LES 

117. Charles E. Cockrell, Jr., Walter C. Engelund, Robert D. Bittner, Tom N. Jentinck, Arthur D. 

Dilley, and Abdelkader Frendi, “Integrative Propulsive Computational Fluid Dynamics Methodology 

for the Hyper-X Flight Experiment.” Journal of Spacecraft and Rockets (Nov.–Dec. 2001), pp. 838, 

843; S. Srinivasan, R.D. Bittner, and B.J. Bobskill, “Summary of the GASP Code Application and 

Evaluation Effort for Scramjet Combustor Flow-fields,” AIAA Paper 93-1973 (1993).

118. T.A. Heppenheimer, “Some Tractable Mathematics for Some Intractable Physics,” Mosaic 

(National Science Foundation), spring 1991, p. 30.

119. Quote in P. Bradshaw, “Progress in Turbulence Research,” AIAA Paper 90-1480 (1990), p. 3.

120. Jack J. Dongarra, “Performance of Various Computers Using Standard Linear Equations Software in a Fortran Environment,” Argonne National Laboratory, Technical Memorandum 23 (Sept. 

30, 1988).

121. “Super Supercomputers,” Aviation Week (Feb. 16, 2009).

122. Krishnan Mahesh, et al., “Large-Eddy Simulation of Gas Turbine Combustors.” Annual 

Research Briefs (Stanford: Center for Turbulence Research, Stanford University, 2001), pp. 3–17.
NASA’s Contributions to Aeronautics

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“is not viewed as an industry standard.” He sees no prospect other than 

direct numerical simulation (DNS), which directly calculates all scales 

of turbulence. “It’s clear-cut,” he adds. “The best way to calculate turbulence is to use DNS. Put in a fine enough grid and calculate the entire 

flow field, including the turbulence. You don’t need any kind of model 

and the turbulence comes out in the wash as part of the solution.” But 

in seeking to apply DNS, even petaflops aren’t enough. Use of DNS for 

practical problems in industry is “many decades down the road. Nobody 

to my knowledge has used DNS to deal with flow through a scramjet. 

That type of application is decades away.”123 With the limitations as well 

as benefits of CFD more readily apparent, it thus is significant that more 

traditional hypersonic test facilities are also improving. As just one example, NASA Langley’s largest hypersonic facility, the 8-foot High Temperature 

Tunnel (HTT), has been refitted to burn methane and use its combustion 

products, with oxygen replenishment, as the test gas. This heats the gas. As 

reviewed by the Journal of Spacecraft and Rockets: “the oxygen content of 

the freestream gas is representative of flight conditions as is the Mach number, total enthalpy, dynamic pressure, and Reynolds number.”124

One fruitful area with NASP had been its aggressive research on 

scramjets, which benefited substantially because of NASA’s increasing 

investment in high-temperature hypersonic test facilities.125

Table 3 enumerates the range of hypersonic test facilities for scramjet 

and aerothermodynamic research available to researchers at the NASA 

Langley Research Center. Between 1987 and the end of 1994, Langley 

researchers ran over 1,500 tests on 10 NASP engine modules, over 1,200 

in a single 3-year period, from the end of 1987 to 1990. After NASP wound 

down, Agency researchers ran nearly 700 tests on four other configurations between 1994 and 1996. These tests, ranging from Mach 4 to Mach 

123. Author interview, John D. Anderson, Jr., Nov. 19, 2008.

124. Quote: Charles E. Cockrell, Jr., et al., Journal of Spacecraft and Rockets (2001), p. 841; 

L.D. Huebner, K.E. Rock, R.T. Voland, and A.R. Wieting, “Calibration of the Langley 8-Foot High 

Temperature Tunnel for Hypersonic Propulsion Airbreathing Testing,” AIAA Paper 96-2197 (1996).

125. R.W. Guy, et al., “Operating Characteristics of the Langley Mach 7 Scramjet Test Facility,” 

NASA TM-81929 (1981); S.R. Thomas and R.W. Guy, “Expanded Operational Capabilities 

of the Langley Mach 7 Scramjet Test Facility, NASA TP-2186 (1983); E.H. Andrews, Jr., et al., 

“Langley Mach 4 Scramjet Test Facility,” NASA TM-86277 (1985); D.E. Reubush and R.L. Puster, 

“Modification to the Langley 8-Ft. High Temperature Tunnel for Hypersonic Propulsion Testing,” AIAA 

Paper 1987-1887 (1987); D.W. Witte, et al., “1998 Calibration of the Mach 4.7 and Mach 6 

Arc-Heated Scramjet Test Facility Nozzles,” NASA TM-2004-213250 (2004).
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TABLE 3

NASA LRC SCRAMJET PROPULSION AND AEROTHERMODYNAMIC TEST FACILITIES

FACILITY NAME MACH REYNOLDS NUMBER SIZE

8-foot High Temperature Tunnel 4, 5, 7 0.3–5.1 x 106 / ft. 8-ft. dia.

Arc-Heated Scramjet Test Facility 4.7–8.0 0.04–2.2 x 106 / ft. 4-ft. dia.

Combustion-Heated Scramjet 

Test Facility

3.5–6.0 1.0–6.8 x 106 / ft. 42” x 30”

Direct Connect Supersonic 

Combustion Test Facility

4.0–7.5 1.8–31.0 x 106

 / ft. [Note (a)]

HYPULSE Shock Tunnel 

[Note (b)]

5.0–25 0.5–2.5 x 106

 / ft. 7-ft dia.

15-inch Mach 6 High 

Temperature Tunnel

6 0.5–8.0 x 106 / ft. 15” dia.

20-inch Mach 6 CF4 Tunnel 6 0.05–0.7 x 106 / ft. 20” dia.

20-inch Mach 6 Tunnel 6 0.5–8.0 x 106 / ft. 20” x 20”

31-inch Mach 10 Tunnel 10 0.2–2.2 x 106 / ft. 31” x 31”

Source: Data from NASA LRC Facility brochures. 

(a) DCSCTF section varies: 1.52” x 3.46” with a M = 2 nozzle and 1.50” x 6.69” 

with a M = 2.7 nozzle.

(b) LRC’s HYPULSE shock tunnel is at the GASL Division of Allied Aerospace Industries, 

Ronkonkoma, NY. 

8, so encouraged scramjet proponents that they went ahead with plans for 

a much-scaled-back effort, the Hyper-X (later designated X-43A), which 

compared in some respects with the ASSET program undertaken after 

cancellation of the X-20 Dyna-Soar three decades earlier.126

The X-43, managed at Langley Research Center by Vincent Rausch, a 

veteran of the earlier TAV and NASP efforts, began in 1995 as Hyper-X, 

coincident with the winddown of NASP. It combined a GASL scramjet 

engine with a 100-inch-long by 60-inch-span slender lifting body and an 

Orbital Sciences Pegasus booster, this combination being carried to a 

launch altitude of 40,000 feet by NASA Dryden’s NB-52B Stratofortress. 

After launch, the Pegasus took the X-43 to approximately 100,000 feet, 

126. The engines and their total successful test runs were Gov’t Baseline (114); Engine A (69); Engine 

A-1 (55); Engine A-2 (321); Engine A-2+ (72); Engine C (233); Engine B-1 (359); NASP SX-20 (160); 

NASP SXPE (142); and NASP CDE (24), a total of 1,549 successful test runs. See the previously cited 

Ruf, “Airframe-Integrated Scramjet Engine Tests in NASA Langley Scramjet Engine Test Facilities.”
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Schematic layout of the Hyper-X (subsequently X-43A) scramjet test vehicle and its Orbital Sciences 

Pegasus winged booster, itself a hypersonic vehicle. NASA.

where it would separate, demonstrating scramjet ignition (using silane 

and then adding gaseous hydrogen) and operation at velocities as 

high as Mach 10.

The X-43 program cost $230 million and consumed not quite a 

decade of development time. Built by Microcraft, Inc., of Tullahoma, 

TN, the X-43 used the shape of a Boeing study for a Mach 10 global 

reconnaissance and space access vehicle, conceived by a team under 

the leadership of George Orton. Langley Research Center furnished 

vital support, executing nearly 900 test runs of 4 engine configurations 

between 1996 and 2003.127

Microcraft completed three X-43A flight-test vehicles for testing 

by NASA Dryden Flight Research Center. Unfortunately, the first flight 

attempt failed in 2001, when the Pegasus booster shed a control fin after 

launch. A 3-year reexamination and review of the program led to a successful flight on March 27, 2004, the first successful hypersonic flight of 

a scramjet-powered airplane. The Pegasus boosted the X-43A to Mach 

6.8. After separation, the X-43A burned silane, which ignites on contact 

127. Engine tests totaled 876, at Mach 5, 7, 10, and 15. The engine configurations, tunnels, and 

test runs were: DFX, 467 runs (97 in CHSTF and 370 in the AHSTF); HXEM, 146 runs (130 in AHSTF and 16 in the 8-ft. HTT); HXFE, 54 runs (all in 8-ft. HTT); and HSM, 209 runs (all in HYPULSE); 

see the previously cited Ruf, “Airframe-Integrated Scramjet Engine Tests in NASA Langley Scramjet 

Engine Test Facilities.”
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with the air, for 3 seconds. Then it ramped down the silane and began 

injecting gaseous hydrogen, burning this gas for 8 seconds. This was 

the world’s first flight test of such a scramjet.128

That November, NASA did it again with its third X-43A. On November 

16, it separated from its booster at 110,000 feet and Mach 9.7 and its 

engine burned for 10 to 12 seconds with silane off. On its face, this looked 

like the fastest air-breathing flight in history, but this speed (approximately 6,500 mph) resulted from its use of Pegasus, a rocket. The key 

point was that the scramjet worked, however briefly. During the flight, 

the X-43A experienced airframe temperatures as high as 3,600 °F.129

Meanwhile, the Air Force was preparing to take the next step with 

its HyTech program. Within it, Pratt & Whitney, now merged with 

Rocketdyne, has been a major participant. In January 2001, it demonstrated the Performance Test Engine (PTE), an airframe-integrated 

scramjet that operated at hypersonic speeds using the hydrocarbon JP-7. 

Like the X-43A engine, though, the PTE was heavy. Its successor, the 

Ground Demonstrator Engine (GDE), was flight-weight. It also used fuel 

to cool the engine structure. One GDE went to Langley for testing in the 

HTT in 2005. It made the important demonstration that the cooling could 

be achieved using no more fuel than was to be employed for propulsion.

Next on transatmospheric agenda is a new X-test vehicle, the X-51A, 

built by Boeing, with a scramjet by Pratt & Whitney Rocketdyne. These 

firms are also participants in a consortium that includes support from 

NASA, DARPA, and the Air Force. The X-51A scramjet is fuel-cooled, 

with the cooling allowing it to be built of Inconel 625 nickel alloy rather 

than an exotic superalloy. Lofted to Mach 4.7 by a modified Army Tactical 

Missile System (ATACMS) artillery rocket booster, the X-51A is intended 

to fly at Mach 7 for minutes at a time, burning JP-7, a hydrocarbon fuel 

128. Bruce A. Smith, “Elevon Failure Precedes Loss of First X-43A,” Aviation Week, June 11, 2001, 

pp. 50–51; Michael Dornheim: “X-43 to Fly in Fall,” July 28, 2003, pp. 36–37; “A Breath of Fast 

Air,” April 5, 2004, pp. 28–29.

129. Thomas J. Bogar, Edwards A. Eiswirth, Lana M. Couch, James L. Hunt, and Charles R. McClinton, “Conceptual Design of a Mach 10, Global Reach Reconnaissance Aircraft,” AIAA Paper 

96-2894 (1996); Charles R. McClinton, Vincent L. Rausch, Joel Sitz, and Paul Reukauf, “Hyper-X 

Program Status,” AIAA Paper 01-0828, 39th Aerospace Sciences Meeting, Reno, NV, Nov. 8–11, 

2001; David E. Reubush, Luat T. Nguyen, and Vincent L. Rausch, “Review of X-43A Return to Flight 

Activities and Current Status,” AIAA Paper 03-7085 (2003); Jay Levine, “Exploring the Hypersonic 

Realm,” The X-Press, vol. 46, no. 10 (Nov. 26, 2004), pp. 1, 8; Michael Dornheim, “But Now 

What?” Aviation Week, Nov. 22, 2004, pp. 24–26.
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The first Boeing X-51 WaveRider undergoing final preparations for flight, Edwards AFB, California, 

2010. USAF

used previously on the Lockheed SR-71. The X-51A uses ethylene to start 

the combustion. Then the flight continues on JP-7. Following checkout 

trials beginning in late 2009, the X-51 made its first powered flight on 

May 26, 2010. After being air-launched from a B-52, it demonstrated 

successful hydrocarbon scramjet ignition and acceleration. Further tests 

will hopefully advance the era of practical scramjet-powered flight, likely 

beginning with long-range missiles. As this review indicates, the story of 

transatmospherics illustrates the complexity of hypersonics; the tenacity and dedication of NASA’s aerodynamics, structures, and propulsion 

community; and the Agency’s commitment to take on challenges, no 

matter how difficult, if the end promises to be the advancement of flight 

and humanity’s ability to utilize the air and space medium.130

130. W. J. Hennigan, “Test Flight Shatters Records,” Los Angeles Times, May 27, 2010; Matthew 

Shaer, “Scramjet-Powered X-51A WaveRider Missile Breaks Mach 6 Record,” Christian Science 

Monitor, May 27, 2010; “WaveRider Sets Record for Hypersonic Flight,” Associated Press 

Release, May 27, 2010.
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Case 5 | Toward Transatmospheric Flight: From V-2 to the X-51

Recommended Additional Readings

Note: The following list represents significant research by NASA and 

its predecessor, the NACA, in the field of transatmospheric flight. These 

references are readily available through the NASA Technical Reports 

Server. Its complete holdings include over half a million citations, of 

which some 90,000 show full text. Users can access it via http://ntrs.nasa.

gov/search.jsp.

Contributions of Hypersonic Leaders:

H. Julian Allen, “Twenty-first Wright Brothers lecture: Hypersonic flight 

and the re-entry problem,” NASA TM-108690, 1957.

H. Julian Allen and Stanford E. Neice, “Problems of performance and 

heating of hypersonic vehicles,” NASA RM-A55L15, 1956.

H. Julian Allen and Murray Toback, “Dynamic stability of vehicles traversing ascending or descending paths through the atmosphere,” 

NASA TN-4275, 1958.

John V. Becker and Peter F. Korycinski, “Heat transfer and pressure distribution at a Mach number of 6.8 on bodies with conical flares and 

extensive flow separation,” NASA RM-L56F22, 1956.

Alfred J. Eggers, H. Julian Allen, and Stanford E. Neice, “A comparative analysis of the performance of long-range hypervelocity vehicles.” NASA TR-1382, 1958.

Maxime A. Faget and H. Rudolph Dettwyler, “Initial flight investigations of a twin-engine supersonic ramjet,” NASA RM-L50H10, 1950.

Maxime A. Faget and O.G. Smith, “Potential improvements to the Shuttle 

through evolution,” AIAA Paper 93-2417, 1993.

Reentry:

R.A. Allen, J.C. Camm, and P.H. Rose, “Nonequilibrium and equilibrium 

radiation at super-satellite re-entry velocities,” NASA CR-51743, 1962.
NASA’s Contributions to Aeronautics

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5

Kim S. Bey, Robert C. Scott, Robert E. Bartels, William A. Waters, and 

Roger Chen, “Analysis of the Shuttle Tile Overlay repair concept,” 

NASA TM-2007-214857, 2007.

G. Bedjai, R.C. Brumfield, S. Demetriades, P.D. Lenn, and D.L. Ward, 

“Design and performance study on a 500 kilowatt linear crossed-field 

air-plasma accelerator for a re-entry facility,” NASA CR-58543, 1964.

H.R. Bredfeldt and W.E. Scharfman, “Use of the Langmuir probe to 

determine the electron density and temperature surrounding reentry vehicles,” NASA CR-66275, 1966.

A.E. Bryson, Jr., and J.L. Speyer, “A neighboring optimum feedback control scheme based on estimated time-to-go with application to reentry flight paths,” NASA CR-87501, 1967.

D.L. Compton, B.J. Short, and S.C. Sommer, “Free-flight measurements 

of static and dynamic stability models of the Project Mercury reentry capsule at Mach numbers 3 and 9.5,” NASA TM-X-373, 1960.

R. Hermann, “Hypersonic aerodynamic problems at re-entry of space 

vehicles,” NASA CR-69486, 1965.

E.J. Hopkins and A.D. Levin, “Reentry glide maneuvers for recovery of 

a winged first-stage rocket booster,” NASA TN-D-1295, 1962.

P.O. Jarvinen, “On the use of magnetohydrodynamics during high speed 

re-entry,” NASA CR-206, 1965.

E. Kaplan and F.D. Linzer, “Considerations in design of calorimeters for the 

Project Fire superorbital re-entry test vehicle,” NASA CR-52196, 1963.

G.C. Kenyon, “The lateral and directional aerodynamic characteristics of 

a re-entry configuration based on a blunt 13 deg half-cone at Mach 

numbers to 0.90,” NASA TM-X-583, 196l.

G.C. Kenyon and F.B. Sutton, “The longitudinal aerodynamic characteristics of a re-entry configuration based on a blunt 13 deg half-cone 

at Mach numbers to 0.92,” NASA TM-X-571, 1961.
Case 5 | Toward Transatmospheric Flight: From V-2 to the X-51

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A.L. Laganelli, “Analysis of flight test transition and turbulent heating 

data. Part 2: Turbulent heating results,” NASA CR-130251, 1972.

R. Lehnert and B. Rosenbaum, “Plasma effects on Apollo re-entry communication,” NASA TN-D-2732, 1965.

A. Martellucci, B.L. Maguire, and R.S. Neff, “Analysis of flight test transition and turbulent heating data. Part l: Boundary layer transition 

results,” NASA CR-129045, 1972.

R.B. Miles, D.A. Santavicca, and M. Zimmermann, “Evaluation of nonintrusive flow measurement techniques for a re-entry flight experiment,” NASA CR-172142, 1983.

A.B. Miller, “Pilot re-entry guidance and control,” NASA CR-331, 1965.

C.T. Swift, “Radiation from slotted-cylinder antennas in a re-entry plasma 

environment,” NASA TN-D-2187, 1964.

F.J. Tischer, “A rough estimate of the ‘blackout’ time in re-entry communications,” NASA TM-X-55059, 1962.

F.J. Tischer, “Attenuation in re-entry communications plans office,” 

TM-X-51027, NASA 1963.

X-15:

E.H. Andrews, Jr., and R.C. Rogers, “Study of underexpanded exhaust 

jets of an X-15 airplane model and attached ramjet engine simulator at Mach 6.86,” NASA TM-X-1571, 1968.

R.G. Bailey, “High total temperature sensing probe for the X-15 hypersonic aircraft,” NASA CR-116772, 1968.

R.D. Banner, A.E. Kuhl, and R.D. Quinn, “Preliminary results of aerodynamic heating studies on the X-15 airplane,” NASA TM-X-638, 1962.
NASA’s Contributions to Aeronautics

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C. Barret, “Review of our national heritage of launch vehicles using 

aerodynamic surfaces and current use of these by other nations,” 

NASA TP-3615, 1996.

S.R. Bland and L.S. Young, “Transonic flutter investigation of models of 

the X-15 airplane horizontal tail,” NASA TM-X-447, 1961.

R.K. Bogue and L.D. Webb, “Advanced air data sensing techniques,” 

NASA TM-X-61115, 1968.

W.G. Cockayne, “Description of an energy management system for the 

X-15,” NASA CR-96006, 1968.

W.R. Deazley, “A study of two proposed stabilization techniques for the 

X-15 horizontal surface control system,” NASA CR-95955, 1961.

R.W. Dunning, “The control characteristics of two preliminary models 

of the X-15 research airplane at Mach numbers of 2.98 and 4.01,” 

NASA TM-X-212, 1960.

D.E. Fetterman, Jr., and J.A. Penland, “Static longitudinal, directional, 

and lateral stability and control data at a Mach number of 6.83 on 

the final configuration of the X-15 research airplane,” NASA TM-X236, 1960.

G.M. Goranson, “Test of a North American X-15 research-vehicle model 

in the JPL 21-inch hypersonic wind tunnel,” NASA CR-53710, 1963.

K.S. Green and T.W. Putnam, “Measurements of sonic booms generated 

by an airplane flying at Mach 3.5 and 4.8,” NASA TM-X-3126, 1974.

E.C. Holleman. M.O. Thompson, and J. Weil, “An assessment of lifting 

reentry flight control requirements during abort, terminal glide, 

and approach and landing situations,” NASA TM-X-59119, 1967.

E.J. Montoya and M. Palitz, “Wind-tunnel investigation of the flow field 

beneath the fuselage of the X-15 airplane at Mach numbers from 4 

to 8,” NASA TM-X-1469, 1967. 
Case 5 | Toward Transatmospheric Flight: From V-2 to the X-51

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5

A.B. Price, “Full scale test report, X-15A-2 ablative thermal protection 

system,” NASA CR-82004, 1968.

A.B. Price, “Thermal protection system X-15A-2. Design report,” NASA 

CR-82003, 1968.

“Progress of the X-15 research airplane program,” NASA SP-90, 1965.

W.J. Sefic, “Friction characteristic of steel skids equipped with skegs on 

a lakebed surface,” NASA TM-81347, 1979.

M.O. Thompson, “General review of piloting problems encountered during simulation and flights of the X-15,” NASA TM-X-56884, 1964.

L.S. Young, “Transonic flutter investigation of models of proposed horizontal tails for the X-15 airplane,” NASA TM-X-442, 1961.

Hot Structures, X-20 Dyna-Soar:

Bianca Trujillo Anderson, Robert R. Meyer, Jr., and Harry R. Chiles, 

“Techniques used in the F-14 variable-sweep transition flight experiment,” NASA TM-100444, 1988.

R.P. Bielat, “Transonic aerodynamic characteristics of the Dyna-Soar 

glider and Titan 3 launch vehicle configuration with various fin 

arrangements,” NASA TM-X-809, 1963.

C.W. Boquist, “Graphite-metal composites,” NASA CR-74308, 1965.

D.A. Buell and G.B. McCullough, “The wind-induced loads on a dynamically scaled model of the Dyna Soar glider and Titan 2 booster in 

launch position,” NASA TM-X-659, 1962.

Victor A. Canacci and Jose C. Gonsalez, “Flow quality measurements 

in an aerodynamic model of NASA Lewis’ Icing Research Tunnel,” 

NASA CR-1999-202353, 1999.
NASA’s Contributions to Aeronautics

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A.J. Chellman, “Development of powder metallurgy 2XXX series Al alloy 

plate and sheet materials for high temperature aircraft structural 

applications,” NASA CR-172521, 1985.

Kim K. deGroh, Donald A. Jaworske, and Daniela C. Smith, “Optical 

property enhancement and durability evaluation of heat receiver 

aperture shield materials,” NASA TM-1998-206623, 1998.

H.L. Giles and J.W. Thomas, “Analysis of hypersonic pressure and heat 

transfer tests on a flat plate with a flap and a delta wing with body, 

elevons, fins, and rudders,” NASA CR-536, 1966.

A.M. Hall, A.F. Hoenie, and C.J. Slunder, “Thermal and mechanical treatment for precipitation-hardening stainless steels,” NASA SP-5089, 

1967.

A.N. Hammer, G.C. Aigret, T.P. Psichogios, and C. Rodgers, “Fabrication 

of cooled radial turbine rotor,” NASA CR-179503, 1986.

A.K. Hepler and A.R. Swegle, “Design and fabrication of brazed Rene 41 

honeycomb sandwich structural panels for advanced space transportation systems,” NASA CR-165801, 1981.

L.S. Jernell and C.D. Babb, “Effect of booster fins on static stability characteristics of a 0.02 scale model of the Titan 3 launch vehicle with 

the Dyna-Soar glider and a bulbous nose at Mach numbers from 

l.60 to 3.50,” NASA TM-X-885, 1963.

M.H. Leipold and C.M. Kapadia, “The role of anions in mechanical failure,” NASA CR-121937, 1971.

D.L. McDaniels and R.A. Signorelli, “Evaluation of low-cost aluminum composites for aircraft engine structural applications,” NASA 

TM-83357, 1983.

M.O. McKinney and J. Scheiman, “Evaluation of turbulence reduction 

devices for the Langley 8-foot Transonic Pressure Tunnel,” NASA 

TM-81792, 1981.
Case 5 | Toward Transatmospheric Flight: From V-2 to the X-51

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5

Shouichi Ochiai, Nanabu Ueno, and Samu Noguchi, “Martensitic transformation and microstructures in sintered NiAl alloys,” NASA TT-20192, 1988.

J. Scheiman, “Considerations for the installation of honeycomb and 

screens to reduce wind-tunnel turbulence,” NASA TM-81868, 1981.

R.E. Shanks and G.N. Ware, “Investigation of the flight characteristics 

of a 1/5-scale model of a Dyna-Soar glider configuration at low subsonic speeds,” NASA TM-X-683, 1962.

A. Tsuge, “Grain boundary engineering of mechanical strength of silicon nitride (Si3N4),” NASA TM-77433, 1984.

S. Yajima, M. Omori, J. Hayashi, H. Kayano, and M. Hamano, “Process 

for the production of metal nitride sintered bodies and resultant silicon 

nitride and aluminum nitride sintered bodies,” NASA TM-77253, 1983.

Shuttle Tiles:

P.J. Bobbitt, C.L.W. Edwards, and R.W. Barnwell, “Simulation of timevarying loads on arrays of Shuttle tiles in a large transonic tunnel,” 

NASA TM-84529, 1982.

O.L. Flowers and D.A. Stewart, “Catalytic surface effects on contaminated Space Shuttle tile in a dissociated nitrogen stream,” NASA 

TM-86770, 1985.

G.L. Giles, “Substructure procedure for including the flexibility in stress 

analysis of Shuttle thermal protection system,” NASA TM-81864, 1980.

G.L. Giles and M. Wallas, “Computer program for nonlinear static stress 

analysis of Shuttle thermal protection system: User’s manual,” NASA 

TM-81856, 198l.

G.L. Giles and M. Vallas, “Use of an engineering data management system 

in the analysis of Space Shuttle orbiter tiles,” NASA TM-83215, 1981.

L.R. Hunt, “Aerodynamic heating in large cavities in an array of RSI 

tiles,” NASA TN-D-8400, 1977.
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Gregory N. Katnick, “Debris/ice/TPS assessment and integrated photographic 

analysis on Shuttle Mission STS-89,” NASA TM-1998-207684, 1998.

Anita Macdonald and Paul Friederich, “Millimeter wave dielectric measurements of Space Shuttle tiles,” NASA CR-187473, 1990.

J. Marroquin and R.R. Burrows, “Results of wind tunnel test OA253 in 

the AEDC 16-T propulsion wind tunnel using a 0.035 scale Space 

Shuttle launch vehicle model 84-OTS and entry vehicle model 84-O,” 

NASA CR-167369, 1982.

B.A. Marshall, “Space Shuttle AFRSI full-scale credibility test in the 

NASA Ames Research Center 11 x 11 foot wind tunnel using model 

124-0 installed in the 96-0 test fixture,” NASA CR-167651, 1982.

A.B. Mattson and C.J. Schwindt, “FTIR instrument to monitor vapors 

from Shuttle tile waterproofing materials,” NASA CR-199959, 1995.

R. Miserentino, L.D. Pinson, and S.A. Leadbetter, “Some Space Shuttle tile/

strain-isolator-pad sinusoidal vibration tests,” NASA TM-81853, 1980.

Timothy R. Moes and Robert R. Meyer, Jr., “In-flight investigation of 

Shuttle tile pressure orifice installations,” NASA TM-4219, 1990.

G.J. Neuner and C.B. Delano, “Development of an improved coating for 

polybenzimidazole foam,” NASA CR-2697, 1976.

A.P. Shore and R. Garcia, “Effects of substate deformation and SIP thickness on tile/SIP interface stresses for Shuttle thermal protection,” 

NASA TM-81855, 1980.

A.A. Stewart, M. Cuellar, and O. Flowers, “Performance of an ablator 

for Space Shuttle in-orbit repair in an arc-plasma airstream,” NASA 

TP-2150, 1983.

S.S. Tompkins, W.D. Brewer, R.K. Clark, C.M. Pittman, and K.L. Brinkley, 

“An assessment of the readiness of ablative materials for preflight 

application to the Shuttle orbiter,” NASA TM-81823, 1980.
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5

W.L. Wells and J. Hudgins, “Experimental assessment of a computer program used in Space Shuttle orbiter entry heating analysis,” NASA 

TM-84572, 1983.

A.J. Zuckerwar and D.R. Sprinkle, “Proposed dynamic phase difference 

method for the detection of tile debonding from the Space Shuttle 

orbiter,” NASA TM-83140, 1981.

X-30 National Aerospace Plane, Scramjet:

A. Abtahi and P. Dean, “Heat flux sensor research and development. The 

cool film calorimeter,” NASA CR-189789, 1990.

E.H. Andrews, Jr., et al., “Langley Mach 4 Scramjet Test Facility,” NASA 

TM-86277, 1985

Steven L. Baughcum and Stephen C. Henderson, “Aircraft emission scenarios projected in the year 2015 for the NASA Technology Concept 

Aircraft (TCA) High Speed Civil Transport,” NASA CR-1998-207635, 1998.

Edwin C. Cady, “Slush hydrogen technology program,” NASA CR-195353, 

1994.

J.A. Cerro, et al., “A study of facilities and fixtures for testing of a high 

speed civil transport wing component,” NASA CR-198352, 1996.

R. Edelman, “Diffusion controlled combustion for scramjet applications,” NASA CR-66363, 1965.

J.C. Evvard, “The scramjet,” NASA TM-X-56755, 1965.

Roger A. Fields, W. Lance Richards, and Michael V. DeAngelis, “Combined 

loads test fixture for thermal-structural testing aerospace vehicle 

panel concepts,” NASA TM-2005-212039, 2004.

R.W. Guy, et al., “Operating Characteristics of the Langley Mach 7 

Scramjet Test Facility,” NASA TM-81929, 1981.
NASA’s Contributions to Aeronautics

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Terry L. Hardy, “FLUSH: a tool for the design of slush hydrogen flow 

systems,” NASA TM-102467, 1990.

Scott D. Holland, “Computational parametric study of sidewall-compression scramjet inlet performance at Mach 10,” NASA TM-4411, 1993.

Marvin Kussoy, George Huang, and Florian Menter, “”Hypersonic flows 

as related to the National Aerospace Plane,” NASA CR-199365, 1995.

Unmeel B. Mehta, “The Aerospace Plane design challenge: Credible computational fluid dynamic results,” NASA TM-102887, 1990.

“National Aero-Space Plane,” NASA TM-109450, 1990.

“National Aerospace Plane thermal development. (Latest citations from 

the Aerospace Database),” NASA TM-97-113072, 1997.

Surya N. Patniak, James D. Guptill, Dale A. Hopkins, and Thomas M. 

Lavelle, “Neural network and regression approximations in high 

speed civil transport aircraft design optimization,” NASA TM-1998-

206316, 1998.

Terrill W. Putnam and Theodore G. Ayers, “Flight research and testing,” 

NASA TM-100439, 1988.

Rodney H. Ricketts, Thomas E. Noll, and Lawrence J. Huttsell, “An 

overview of aeroelasticity studies for the National Aerospace Plane,” 

NASA TM-107728, 1993.

William C. Rose, “Numerical investigations in three-dimensional internal flows,” NASA CR-183108, 1988.

“Shuttle to Space Station. Heart assist implant. Hubble update. X-30 

mock-up,” NASA TM-110837, 1992.

J.C. Tannehill and G. Wadawadigi, “Development of a 3-D upwind 

PNS code for chemically reacting hypersonic flowfields,” NASA 

CR-190182, 1992.
Case 5 | Toward Transatmospheric Flight: From V-2 to the X-51

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5

S.R. Thomas and R.W. Guy, “Expanded Operational Capabilities of the 

Langley Mach 7 Scramjet Test Facility,” NASA TP-2186, 1983.

M.E. Tuttle and J. Rogacki, “Thermoviscoplastic response of Ti-15-3 

under various loading conditions,” NASA CR-187621, 1991.

D.W. Witte, et al., “1998 Calibration of the Mach 4.7 and Mach 6 ArcHeated Scramjet Test Facility Nozzles,” NASA TM-2004-213250, 2004.

X-43 Hyper-X:

“2000–2001 Research Engineering Annual Report,” NASA TM-2004-

212025, 2004.

Ethan Baumann, Catherine Bahm, Brian Strovers, Roger Beck, and 

Michael Richard, “The X-43A six degree of freedom Monte Carlo 

analysis,” NASA TM-2007-214630, 2007.

Scott A. Berry, Michael DiFulvio, and Matthew K. Kowalkowski, “Forced 

Boundary-Layer Transition on X-43 (Hyper-X) in NASA LaRC 20-Inch 

Mach 6 Air Tunnel,” NASA TM-2000-210316.

Scott A. Berry, Michael DiFulvio, and Matthew K. Kowalkowski, “Forced 

Boundary-Layer Transition on X-43 (Hyper-X) in NASA LaRC 31-Inch 

Mach 6 Air Tunnel,” NASA TM-2000-210315.

Edward H. Glaessgen, et al., “X-43A rudder spindle fatigue life estimate 

and testing,” NASA TM-2005-213525, 2005.

William L. Ko and Leslie Gong, “Thermoelastic analysis of Hyper-X camera windows suddenly exposed to Mach 7 stagnation aerothermal 

shock,” NASA TP-2000-209030, 2000.

J.A. Lee and P.S. Chen, “Aluminum-scandium alloys: material characterization, friction stir welding, and compatibility with hydrogen 

peroxide,” NASA TM-2004-213604, 2004.
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Jessica Lux-Baumann, Ray Dees, and David Fratello, “Control room 

training for the Hyper-X project utilizing aircraft simulation,” NASA 

TM-2006-213685, 2006.

Chan-gi Pak, “Aeroservoelastic stability analysis of the X-43A stack,” 

NASA TM-2008-214365, 2008.

Matthew Redif, Yohan Lin, Courtney Amos Besssent, and Carole Barklow, 

“The Hyper-X flight systems validation program,” NASA TM-2007-

214620, 2007.

Karla S. Shy, Jacob J. Hageman, and Jeanette H. Le, “The aircraft simulation role in improving safety through control room training,” 

NASA TM-2002-210731, 2002.
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360

The Northrop HL-10 turning onto final approach for a lakebed landing. NASA.
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6

Physical Problems, 

Challenges, and 

Pragmatic Solutions

Robert G. Hoey

CASE

6

The advent of the supersonic and hypersonic era introduced a wide 

range of operational challenges that required creative insight by the flight 

research community. Among these were phenomena such as inertial (roll) 

coupling, transonic pitch-up, panel flutter, structural resonances, pilotinduced oscillations, and aerothermodynamic heating. Researchers had 

to incorporate a variety of solutions and refine simulation techniques to 

better predict the realities of flight. The efforts of the NACA and NASA, 

in partnership with other organizations, including the military, enabled 

development and refinement of reliable aerospace vehicle systems.

T

HE HISTORY OF AVIATION is replete with challenges and difficulties overcome by creative scientists and engineers whose insight, 

coupled with often-pragmatic solutions, broke through what had 

appeared to be barriers to future flight. At the dawn of aviation, the 

problems were largely evident to all: for example, simply developing a 

winged vehicle that could take off, sustain itself in the air, fly in a controlled fashion, and then land. As aviation progressed, the problems and 

challenges became more subtle but no less demanding. The National 

Advisory Committee on Aeronautics (NACA) had been created in 1915 

to pursue the “scientific study of the problems of flight, with a view 

to their practical solution,” and that spirit carried over into the aeronautics programs of the National Aeronautics and Space Administration 

(NASA), which succeeded the NACA on October 1, 1958, not quite a year 

after Sputnik had electrified the world. The role of the NACA, and later 

NASA, is mentioned often in the following discussion. Both have been 

instrumental in the discovery and solution to many of these problems.

As aircraft flight speeds moved from the firmly subsonic through the 

transonic and into the supersonic and even hypersonic regimes, the continuing challenge of addressing unexpected interactions and problems 
NASA’s Contributions to Aeronautics

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followed right along. Since an airplane is an integrated system, many of 

these problems crossed multiple discipline areas and affected multiple 

aspects of an aircraft’s performance, or flight safety. Numerous examples 

could be selected, but the author has chosen to examine a representative 

sampling from several areas: experience with flight control systems and 

their design, structures, and their aeroelastic manifestations; flight simulation; flight dynamics (the motions and experience of the airplane in 

flight); and aerothermodynamics, the demanding environment of aerodynamic heating that affects a vehicle and its structure at higher velocities. 

Flight Control Systems and Their Design

During the Second World War, there were multiple documented incidents and several fatalities that occurred when fighter pilots dove their 

propeller-driven airplanes at speeds approaching the speed of sound. 

Pilots reported increasing levels of buffet and loss of control at these 

speeds. Wind tunnels at that time were incapable of producing reliable 

meaningful data in the transonic speed range because the local shock 

waves were reflected off the wind tunnel walls, thus invalidating the data 

measurements. The NACA and the Department of Defense (DOD) created a new research airplane program to obtain a better understanding 

of transonic phenomena through flight-testing. The first of the resulting 

aircraft was the Bell XS-1 (later X-1) rocket-powered research airplane.

On NACA advice, Bell had designed the X-1 with a horizontal tail 

configuration consisting of an adjustable horizontal stabilizer with a 

hinged elevator at the rear for pitch control, at a time when a fixed horizontal tail and hinged elevator constituted the standard pitch control 

configuration for that period.1 The X-1 incorporated this as an emergency 

means to increase its longitudinal (pitch) control authority at transonic 

speeds. It proved a wise precaution because, during the early buildup 

flights, the X-1 encountered similar buffet and loss of control as reported 

by earlier fighters. Analysis showed that local shock waves were forming on the tail surface, eventually migrating to the elevator hinge line. 

When they reached the hinge line, the effectiveness of the elevator was 

significantly reduced, thus causing the loss of control. The X-1 NACA–

U.S. Air Force (USAF) test team determined to go ahead, thanks to the 

1. R.M. Stanley and R.J. Sandstrom, “Development of the XS-1 Airplane,” HQ Air Materiel Command, Air Force Supersonic Research Airplane XS-1, Report No. 1 (Wright Field: Air Materiel 

Command, Jan. 9, 1948), p. 7. 
Case 6 | Physical Problems, Challenges, and Pragmatic Solutions

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X-1 having an adjustable horizontal tail. They subsequently validated 

that the airplane could be controlled in the transonic region by moving 

the horizontal stabilizer and the elevator together as a single unit. This 

discovery allowed Capt. Charles E. Yeager to exceed the speed of sound 

in controlled flight with the X-1 on October 14, 1947.2

An extensive program of transonic testing was undertaken at the 

NACA High-Speed Flight Station (HSFS; subsequently the Dryden 

Flight Research Center) on evaluating aircraft handling qualities using 

the conventional elevator and then the elevator with adjustable stabilizer.3

 As a result, subsequent transonic airplanes were all designed to use a 

one-piece, all-flying horizontal stabilizer, which solved the control problem 

and was incorporated on the prototypes of the first supersonic American jet 

fighters, the North American YF-100A, and the Vought XF8U-1 Crusader, 

flown in 1953 and 1954. Today, the all-moving tail is a standard design element of virtually all high-speed aircraft developed around the globe.4

Resolving the Challenge of Aerodynamic Damping 

Researchers in the early supersonic era also faced the challenges posed 

by the lack of aerodynamic damping. Aerodynamic damping is the natural resistance of an airplane to rotational movement about its center of 

gravity while flying in the atmosphere. In its simplest form, it consists 

of forces created on aerodynamic surfaces that are some distance from 

the center of gravity (cg). For example, when an airplane rotates about 

the cg in the pitch axis, the horizontal tail, being some distance aft of 

the cg, will translate up or down. This translational motion produces 

a vertical lift force on the tail surface and a moment (force times distance) that tends to resist the rotational motion. This lift force opposes 

the rotation regardless of the direction of the motion. The resisting force 

will be proportional to the rate of rotation, or pitch rate. The faster the 

rotational rate, the larger will be the resisting force. The magnitude of 

2. Walter C. Williams, “Instrumentation, Airspeed Calibration, Tests, Results and Conclusions,” HQ 

AMC, Air Force Supersonic Research Airplane XS-1, p. 24. 

3. W.C. Williams and A.S. Crossfield, “Handling Qualities of High-Speed Airplanes,” RM L52A08 

(Jan. 28, 1952).

4. It should be noted, of course, that the all-moving tail was essentially a “rediscovery” of earlier 

design practice. All-moving tails, for very different reasons, had been a feature of early airplanes, 

typified by the Wright Flyer and numerous European examples such as the Blériot and the Fokker 

Eindecker.
NASA’s Contributions to Aeronautics

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the resisting tail lift force is dependent on the change in angle of attack 

created by the rotation. This change in angle of attack is the vector sum 

of the rotational velocity and the forward velocity of the airplane. For 

low forward velocities, the angle of attack change is quite large and the 

natural damping is also large. The high aerodynamic damping associated with the low speeds of the Wright brothers flights contributed a 

great deal to the brothers’ ability to control the static longitudinal instability of their early vehicles.

At very high forward speed, the same pitch rate will produce a much 

smaller change in angle of attack and thus lower damping. For practical 

purposes, all aerodynamic damping can be considered to be inversely 

proportional to true velocity. The significance of this is that an airplane’s 

natural resistance to oscillatory motion, in all axes, disappears as the 

true speed increases. At hypersonic speeds (above Mach 5), any rotational disturbance will create an oscillation that will essentially not 

damp out by itself.

As airplanes flew ever faster, this lightly damped, oscillatory tendency became more obvious and was a hindrance to accurate weapons delivery for military aircraft, and pilot and passenger comfort for 

commercial aircraft. Evaluating the seriousness of the damping challenge in an era when aircraft design was changing markedly (from the 

straight-wing propeller-driven airplane to the swept and delta wing jet 

and beyond). It occupied a great amount of attention from the NACA 

and early NASA researchers, who recognized that it would pose a continuing hindrance to the exploitation of the transonic and supersonic 

region, and the hypersonic beyond.5

In general, aerodynamic damping has a positive influence on handling qualities, because it tends to suppress the oscillatory tendencies 

of a naturally stable airplane. Unfortunately, it gradually disappears as 

the speed increases, indicating the need for some artificial method of 

suppressing these oscillations during high-speed flight. In the preelectronic flight control era, the solution was the modification of flight control systems to incorporate electronic damper systems, often referred to 

as Stability Augmentation Systems (SAS). A damper system for one axis 

consisted of a rate gyro measuring rotational rate in that axis, a gainchanging circuit that adjusted the size of the needed control command, 

5. For cases, see Edwin J. Saltzman and Theodore G. Ayers, Selected Examples of NACA/NASA 

Supersonic Flight Research, SP-513 (Edwards, CA: NASA Dryden Flight Research Center, 1995).
Case 6 | Physical Problems, Challenges, and Pragmatic Solutions

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and a servo mechanism that added additional control surface commands 

to the commands from the pilot’s stick. Control surface commands were 

generated that were proportional to the measured rotational rate (feedback) but opposite in sign, thus driving the rotational rate toward zero.

Damper systems were installed in at least one axis of all of the Centuryseries fighters (F-100 through F-107), and all were successful in stabilizing 

the aircraft in high-speed flight.6 Development of stability augmentation 

systems—and their refinement through contractor, Air Force–Navy, and 

NACA–NASA testing—was crucial to meeting the challenge of developing 

Cold War airpower forces, made yet more demanding because the United 

States and the larger NATO alliance chose a conscious strategy of using 

advanced technology to generate high-leverage aircraft systems that could 

offset larger numbers of less-individually capable Soviet-bloc designs.7

Early, simple damper systems were so-called single-string systems and 

were designed to be “fail-safe.” A single gyro, servo, and wiring system 

were installed for each axis. The feedback gains were quite low, tailored 

to the damping requirements at high speed, at which very little control 

surface travel was necessary. The servo travel was limited to a very small 

value, usually less than 2 degrees of control surface movement. A failure in 

the system could drive the servo to its maximum travel, but the transient 

motion was small and easily compensated by the pilot. Loss of a damper 

at high speed thus reduced the comfort level, or weapons delivery accuracy, but was tolerable, and, at lower speeds associated with takeoff and 

landing, the natural aerodynamic damping was adequate. 

One of the first airplanes to utilize electronic redundancy in the design 

of its flight control system was the X-15 rocket-powered research airplane, which, at the time of its design, faced numerous unknowns. 

Because of the extreme flight conditions (Mach 6 and 250,000-foot altitude), the servo travel needed for damping was quite large, and the 

pilot could not compensate if the servo received a hard-over signal. 

6. Robert. G. Hoey and Capt. Iven C. Kincheloe, “F-104A Stability and Control,” AFFTC TR-56-14, 

April 1958.

7. See Joseph R. Chambers, Partners in Freedom: Contributions of the Langley Research Center 

to U.S. Military Aircraft of the 1990s, SP-2000-4519 (Washington, DC: NASA, 2000), passim; 

Robert K. Geiger, et al., The AGARD History, 1952–1987 (Neuilly sur Seine: NATO Advisory 

Group for Aeronautical Research and Development, 1988 ed.), pp. ix–xxv; and Thomas C. Lassman, Sources of Weapon Systems Innovation in the Department of Defense: The Role of In-House 

Research and Development, 1945–2000 (Washington, DC: Center for Military History, 2008), 

pp. 93–97.
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The solution was the incorporation of an independent, but identical, 

feedback “monitoring” channel in addition to the “working” channel in 

each axis. The servo commands from the monitor and working channel 

were continuously compared, and when a disagreement was detected, 

the system was automatically disengaged and the servo centered. This 

provided the equivalent level of protection to the limited-authority failsafe damper systems incorporated in the Century series fighters. Two 

of the three X-15s retained this fail-safe damper system throughout the 

9-year NASA–Air Force–Navy test program, although a backup roll rate 

gyro was added to provide fail-operational, fail-safe capability in the 

roll axis.8 Refining the X-15’s SAS system necessitated a great amount 

of analysis and simulator work before the pilots deemed it acceptable, 

particularly as the X-15’s stability deteriorated markedly at higher angles 

of attack above Mach 2. Indeed, one of the major aspects of the X-15’s 

research program was refining understanding of the complexities of 

hypersonic stability and control, particularly during reentry at high 

angles of attack.9

The electronic revolution dramatically reshaped design approaches 

to damping and stability. Once it was recognized that electronic assistance was beneficial to a pilot’s ability to control an airplane, the concept evolved rapidly. By adding a third independent channel, and some 

electronic voting logic, a failed channel could be identified and its signal “voted out,” while retaining the remaining two channels active. 

If a second failure occurred (that is, the two remaining channels did 

not agree), the system would be disconnected and the damper would 

become inoperable. Damper systems of this type were referred to as failoperational, fail-safe (FOFS) systems. Further enhancement was provided 

by comparing the pilot’s stick commands with the measured airplane 

response and using analog computer circuits to tailor servo commands 

so that the airplane response was nearly the same for all flight conditions. 

These systems were referred to as Command Augmentation Systems 

(CAS). The next step in the evolution was the incorporation of a mathematical model of the desired aircraft response into the analog computer 

circuitry. An error signal was generated by comparing the instantaneous 

8. Personal Experience as an Air Force Flight Planner during the X-15 envelope expansion flight-testing.

9. Robert A. Tremant, “Operational Experience and Characteristics of the X-15 Flight Control System,” NASA Technical Note D-1402 (Dec. 1962), and Wendell H. Stillwell, X-15 Research Results, 

SP-60 (Washington, DC: NASA, 1965), pp. 51–52.
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measured aircraft response with the desired mathematical-model 

response, and the servo commands forced the airplane to fly per the 

mathematical model, regardless of the airplane’s inherent aerodynamic 

tendencies. These systems were called “model-following.”

Even higher levels of redundancy were necessary for safe operation 

of these advanced control concepts after multiple failures, and the failure logic became increasingly more complex. Establishing the proper 

“trip” levels, where an erroneous comparison would result in the exclusion of one channel, was an especially challenging task. If the trip levels 

were too tight, a small difference between the outputs of two perfectly 

good gyros would result in nuisance trips, while trip levels that were 

too loose could result in a failed gyro not being recognized in a timely 

manner. Trip levels were usually adjusted during flight test to provide 

the safest settings. 

NASA’s Space Shuttle orbiter utilized five independent control system 

computers. Four had identical software. This provided fail-operational, 

fail-operational, fail-safe (FOFOFS) capability. The fifth computer used 

a different software program with a “get-me-home” capability as a last 

resort (often referred to as the “freeze-dried” control system computer).

The Challenge of Limit-Cycles

The success of the new electronic control system concepts was based on 

the use of electrical signals from sensors (primarily rate gyros and accelerometers) that could be fed into the flight control system to control aircraft motion. As these electronic elements began to play a larger role, a 

different dynamic phenomenon came into play. “Limit-cycles” are a common characteristic of nearly all mechanical-electrical closed-loop systems and are related to the total gain of the feedback loop. For an aircraft 

flight control system, total loop gain is the product of two variables: (1) 

the magnitude of the aerodynamic effectiveness of the control surface for 

creating rotational motion (aerodynamic gain) and (2) the magnitude of 

the artificially created control surface command to the control surface 

(electrical gain). When the aerodynamic gain is low, such as at very low 

airspeeds, the electrical gain will be correspondingly high to command 

large surface deflections and rapid aircraft response. Conversely, when the 

aerodynamic gain is high, such as at high airspeed, low electrical gains 

and small surface deflections are needed for rapid airplane response.

These systems all have small dead bands, lags, and rate limits (nonlinearities) inherent in their final, real-world construction. When the 
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total feedback gain is increased, the closed-loop system will eventually 

exhibit a small oscillation (limit-cycle) within this nonlinear region. The 

resultant total loop gain, which causes a continuous, undamped limitcycle to begin, represents the practical upper limit for the system gain 

since a further increase in gain will cause the system to become unstable 

and diverge rapidly, a condition which could result in structural failure 

of the system. Typically the limit-cycle frequency for an aircraft control 

system is between two and four cycles per second.

Notice that the limit-cycle characteristics, or boundaries, are dependent upon an accurate knowledge of control surface effectiveness. Ground 

tests for limit-cycle boundaries were first devised by NASA Dryden Flight 

Research Center (DFRC) for the X-15 program and were accomplished by 

using a portable analog computer, positioned next to the airplane, to generate the predicted aerodynamic control effectiveness portion of the feedback path.10 The control system rate gyro on the airplane was bypassed, and 

the analog computer was used to generate the predicted aircraft response 

that would have been generated had the airplane been actually flying. This 

equivalent rate gyro output was then inserted into the control system. The 

total loop gain was then gradually increased at the analog computer until a 

sustained limit-cycle was observed at the control surface. Small stick raps 

were used to introduce a disturbance in the closed-loop system in order to 

observe the damping characteristics. Once the limit-cycle total loop gain 

boundaries were determined, the predicted aerodynamic gains for various 

flight conditions were used to establish electrical gain limits over the flight 

envelope. These ground tests became routine at NASA Dryden and at the 

Air Force Flight Test Center (AFFTC) for all new aircraft.11 For subsequent 

production aircraft, the resulting gain schedules were programmed within 

the flight control system computer. Real-time, direct measurements of airspeed, altitude, Mach number, and angle of attack were used to access and 

adjust the electrical gain schedules while in flight to provide the highest 

safe feedback gain while avoiding limit-cycle boundaries.

Although the limit-cycle ground tests described above had been performed, the NASA–Northrop HL-10 lifting body encountered limit-cycle 

10. L.W. Taylor, Jr., and J.W. Smith, “An Analysis of the Limit-Cycle and Structural-Resonance 

Characteristics of the X-15 Stability Augmentation System,” NASA TN-D-4287 (Dec. 1967).

11. Weneth D. Painter and George J. Sitterle, “Ground and Flight Test Methods for Determining 

Limit Cycle and Structural Resonance Characteristics of Aircraft Stability Augmentation Systems,” 

NASA TN-D-6867 (1972). 
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oscillations on its maiden flight. After launch from the NB-52, the telemetry data showed a large limit-cycle oscillation of the elevons. The oscillations were large enough that the pilot could feel the aircraft motion in 

the cockpit. NASA pilot Bruce Peterson manually lowered the pitch gain, 

which reduced the severity of the limit-cycle. Additional aerodynamic 

problems were present during the short flight requiring that the final 

landing approach be performed at a higher-than-normal airspeed. This 

caused the limit-cycle oscillations to begin again, and the pitch gain was 

reduced even further by Peterson, who then capped his already impressive performance by landing the craft safely at well over 300 mph. NASA 

engineer Weneth Painter insisted the flight be thoroughly analyzed before 

the test team made another flight attempt, and subsequent analysis by 

Robert Kempel and a team of engineers concluded that the wind tunnel 

predictions of elevon control effectiveness were considerably lower than 

the effectiveness experienced in flight.12 This resulted in a higher aerodynamic gain than expected in the total loop feedback path and required 

a reassessment of the maximum electrical gain that could be tolerated.13

Flight Control Systems and Pilot-Induced Oscillations

Pilot-induced oscillations (PIO) occur when the pilot commands the 

control surfaces to move at a frequency and/or magnitude beyond the 

capability of the surface actuators. When a hydraulic actuator is commanded to move beyond its design rate limit, it will lag behind the commanded deflection. If the command is oscillatory in nature, then the 

resulting surface movement will be smaller, and at a lower rate, than 

commanded. The pilot senses a lack of responsiveness and commands 

even larger surface deflections. This is the same instability that can be 

generated by a high-gain limit-cycle, except that the feedback path is 

through the pilot’s stick, rather than through a sensor and an electronic 

servo. The instability will continue until the pilot reduces his gain (ceases 

to command large rapid surface movement), thus allowing the actuator 

to return to its normal operating range. 

12. R.W. Kempel and J.A. Manke, “Flight Evaluation of HL-10 Lifting Body Handling Qualities at 

Mach Numbers from 0.30 to 1.86”, NASA TN-D-7537, (Jan. 1974).

13. Milton O. Thompson with J.D. Hunley, Flight Research: Problems Encountered and What They 

Should Teach Us, SP-2000-4522 (Washington, DC: NASA, 2000), pp. 19–20; see also R. 

Dale Reed with Darlene Lister, Wingless Flight: The Lifting Body Story, SP-4220 (Washington, DC: 

NASA, 1997), pp. 96–102. 
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The prototype General Dynamics YF-16 Lightweight Fighter (LWF) 

unexpectedly encountered a serious PIO problem on a high-speed taxi 

test in 1974. The airplane began to oscillate in roll near the end of the 

test. The pilot, Philip Oestricher, applied large, corrective stick inputs, 

which saturated the control actuators and produced a pilot-induced 

oscillation. When the airplane began heading toward the side of the 

runway, the pilot elected to add power and fly the airplane rather than 

veer into the dirt along side of the runway. Shortly after the airplane 

became airborne, his large stick inputs ceased, and the PIO and limitcycle stopped. Oestricher then flew a normal pattern and landed the airplane safely. Several days later, after suitable modifications to its flight 

control system, it completed its “official” first flight.

The cause of this problem was primarily related to the “force stick” 

used in the prototype YF-16. The control stick was rigidly attached to the 

airplane, and strain gages on the stick measured the force being applied 

by the pilot. This electrical signal was transmitted to the flight control system as the pilot’s command. There was no motion of the stick, thus no 

feedback to the pilot of how much control deflection he was commanding. 

During the taxi test, the pilot was unaware that he was commanding full 

deflection in roll, thus saturating the actuators. The solution was a reduction in the gain of the pilot’s command signal, as well as a geometry change 

to the stick that allowed a small amount of stick movement. This gave the 

pilot some tactile feedback as to the amount of control deflection being 

commanded, and a hard stop when the stick was commanding full deflection.14 The incident offered lessons in both control system design and in 

human factors engineering, particularly on the importance of ensuring 

that pilots receive indications of the magnitude of their control inputs 

via movable sticks. Subsequent fly-by-wire (FBW) aircraft have incorporated this feature, as opposed to the “fixed” stick concept tried on the 

YF-16. As for the YF-16, it won the Lightweight Fighter design competition, was placed in service in more developed form as the F-16 Fighting 

Falcon, and subsequently became a widely produced Western jet fighter.

Another PIO occurred during the first runway landing of the NASA–

Rockwell Space Shuttle orbiter during its approach and landing tests in 

1978. After the flare, and just before touchdown, astronaut pilot Fred 

Haise commanded a fairly large pitch control input that saturated the 

14. Personal recollections from serving as a member of the YF-16 Taxi Test Incident review team.
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The General Dynamics YF-16 prototype Lightweight Fighter (LWF) in flight over the Edwards 

range. USAF.

control actuators. At touchdown, the orbiter bounced slightly and the 

rate-limiting saturation transferred to the roll axis. In an effort to keep the 

wings level, the pilot made additional roll inputs that created a momentary pilot-induced oscillation that continued until the final touchdown. 

At one point, it seemed the orbiter might veer toward spectators, one of 

whom was Britain’s Prince Charles, then on a VIP tour of the United States. 

(Ironically, days earlier, the Prince of Wales had “flown” the Shuttle simulator at the NASA Johnson Space Center, encountering the same kind of 

lateral PIO that Haise did on touchdown.) Again, the cause was related 

to the high sensitivity of the stick in comparison with the Shuttle’s slowmoving elevon actuators. The incident sparked a long and detailed study 

of the orbiter’s control system in simulators and on the actual vehicle. 

Several changes were made to the control system, including a reduced 

sensitivity of the stick and an increase in the maximum actuator rates.15

15. Robert G. Hoey, et al., “Flight Test Results from the Entry and Landing of the Space Shuttle 

Orbiter for the First Twelve Orbital Flights,” AFFTC TR-85-11 (1985), p. 104. Robert G. Hoey, et 

al., “AFFTC Evaluation of the Space Shuttle Orbiter and Carrier Aircraft—NASA Approach and 

Landing Test,” AFFTC TR-78-14, May 1978, pp. 104, 114, 117. See also Richard P. Hallion, On 

the Frontier: Flight Research at Dryden, 1946–1981, SP-4303 (Washington, DC: NASA, 1984), 

pp. 249–250.
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The above discussion of electronic control system evolution has 

sequentially addressed the increasing complexity of the systems. This was 

not necessarily the actual chronological sequence. The North American 

F-107, an experimental nuclear strike fighter derived from the earlier 

F-100 Super Sabre, utilized one of the first fly-by-wire control systems—

Augmented Longitudinal Control System (ALCS)—in 1956. One of the 

three prototypes was used by NASA, thus providing the Agency with its 

first exposure to fly-by-wire technology. Difficult maintenance of the 

one-of-kind subsystems in the F-107 forced NASA to abandon its use 

as a research airplane after about 1 year of flying.

Self-Adaptive Flight Control Systems

One of the more sophisticated electronic control system concepts was 

funded by the AF Flight Dynamics Lab and created by Minneapolis 

Honeywell in the late 1950s for use in the Air Force-NASA-Boeing X-20 

Dyna-Soar reentry glider. The extreme environment associated with 

a reentry from space (across a large range of dynamic pressures and 

Mach numbers) caused engineers to seek a better way of adjusting the 

feedback gains than stored programs and direct measurements of the 

atmospheric variables. The concept was based on increasing the electrical gain until a small limit-cycle was measured at the control surface, 

then alternately lowering and raising the electrical gain to maintain a 

small continuous, but controlled, limit-cycle throughout the flight. This 

allowed the total loop gains to remain at their highest safe value but 

avoided the need to accurately predict (or measure) the aerodynamic 

gains (control surface effectiveness). 

This system, the MH-96 Adaptive Flight Control System (AFCS), was 

installed in a McDonnell F-101 Voodoo testbed and flown successfully 

by Minneapolis Honeywell in 1959–1960. It proved to be fairly robust in 

flight, and further system development occurred after the cancellation 

of the X-20 Dyna-Soar program in 1963. After a ground-test explosion 

during an engine run with the third X-15 in June 1960, NASA and the 

Air Force decided to install the MH-96 in the hypersonic research aircraft when it was rebuilt. The system was expanded to include several 

autopilot features, as well as a blending of the aerodynamic and reaction controls for the entry environment. The system was triply redundant, thus providing fail-operational, fail-safe capability. This was an 

improvement over the other two X-15s, which had only fail-safe features. Because of the added features of the MH-96, and the additional 
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redundancy it provided, NASA and the Air Force used the third X-15 for 

all planned high-altitude flights (above 250,000 feet) after an initial envelope expansion program to validate the aircraft’s basic performance.16

Unfortunately, on November 15, 1967, the third X-15 crashed, killing its pilot, Major Michael J. Adams. The loss of X-15 No. 3 was related 

to the MH-96 Adaptive Flight Control System design, along with several 

other factors. The aircraft began a drift off its heading and then entered 

a spin at high altitude (where dynamic pressure—“q” in engineering 

shorthand—is very low). The flight control system gain was at its maximum when the spin started. The control surfaces were all deflected to 

their respective stops attempting to counter the spin, thus no limit-cycle 

motion—4 hertz (Hz) for this airplane—was being detected by the gain 

changer. Thus, it remained at maximum gain, even though the dynamic 

pressure (and hence the structural loading) was increasing rapidly during entry. When the spin finally broke and the airplane returned to a 

normal angle of attack, the gain was well above normal, and the system commanded maximum pitch rate response from the all-moving 

elevon surface actuators. With the surface actuators operating at their 

maximum rate, there was still no 4-Hz limit-cycle being sensed by the 

gain changer, and the gain remained at the maximum value, driving the 

airplane into structural failure at approximately 60,000 feet and at a 

velocity of Mach 3.93.17

As the accident to the third X-15 indicated, the self-adaptive control system concept, although used successfully for several years, had 

some subtle yet profound difficulties that resulted in it being used in 

only one subsequent production aircraft, the General Dynamics F-111 

multipurpose strike aircraft. One characteristic common to most of the 

model-following systems was a disturbing tendency to mask deteriorating handling qualities. The system was capable of providing good handling qualities to the pilot right up until the system became saturated, 

resulting in an instantaneous loss of control without the typical warning a pilot would receive from any of the traditional signs of impending 

loss of control, such as lightening of control forces and the beginning 

16. Dennis R. Jenkins, X-15: Extending the Frontiers of Flight, SP-2007-562 (Washington, DC: 

NASA, 2007), p. 402. 

17. Donald R. Bellman, et al., Investigation of the Crash of the X-15-3 Aircraft on November 15, 

1967 (Edwards: NASA Flight Research Center, Jan. 1968), pp. 8–15. 
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of control reversal.18 A second serious drawback that affected the F-111 

was the relative ease with which the self-adaptive system’s gain changer 

could be “fooled,” as with the accident to the third X-15. During early 

testing of the self-adaptive flight control system on the F-111, testers discovered that, while the plane was flying in very still air, the gain changer 

in the flight control system could drive the gain to quite high values 

before the limit-cycle was observed. Then a divergent limit-cycle would 

occur for several seconds while the gain changer stepped the gain back 

to the proper levels. The solution was to install a “thumper” in the system that periodically introduced a small bump in the control system to 

start an oscillation that the gain changer could recognize. These oscillations were small and not detectable by the pilot, and thus, by inducing 

a little “acceptable” perturbation, the danger of encountering an unexpected larger one was avoided. 

For most current airplane applications, flight control systems use 

stored gain schedules as a function of measured flight conditions (altitude, airspeed, etc.). The air data measurement systems are already 

installed on the airplane for pilot displays and navigational purposes, 

so the additional complication of a self-adaptive feature is considered 

unnecessary. As the third X-15’s accident indicated, even a well-designed 

adaptive flight control system can be fooled, resulting in tragic consequences.19 The “lesson learned,” of course (or, more properly, the “lesson relearned”) is that the more complex the system, the harder it is to 

identify the potential hazards. It is a lesson that engineers and designers might profitably take to heart, no matter what their specialty.

Induced Structural Resonances

Overall, electronic enhancements introduced significant challenges with 

respect to their practical incorporation in an airplane. Model-following 

systems required highly responsive servos and high gain levels for the 

feedback from the motion sensors (gyros and accelerometers) to the control surfaces. These high-feedback-gain requirements introduced serious 

issues regarding the aircraft structure. An aircraft structure is surprisingly vulnerable to induced frequencies, which, like a struck musical 

tuning fork, can result in resonant motions that may reach the naturally 

18. For more on its strengths and weaknesses, see L.W. Taylor, Jr., and E.J. Adkins, “Adaptive Flight 

Control Systems—Pro and Con,” NASA TM-X-56008 (1964).

19. Personal experience as an X-15 flight planner and X-20 stability and control flight test engineer.
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The General Dynamics F-111A was the first production aircraft to use a self-adaptive flight control system. NASA.

destructive frequency of the structure, breaking it apart. Rapid movement of a control surface could trigger a lightly damped oscillation of 

one of the structural modes of the airplane (first mode tail bending, for 

example). This structural oscillation could be detected by the flight control system sensors, resulting in further rapid movement of the control 

surface. The resulting structural/control surface oscillation could thus 

be sustained, or even amplified. These additive vibrations were typically 

at higher frequencies (5–30 Hz) than the limit-cycle described earlier 

(2–4 Hz), although some of the landing gear modes and wing bending 

modes for larger aircraft are typically below 5 Hz. If seemingly esoteric, 

this phenomenon, called structural resonance, is profoundly serious. 

Even the stiff and dense X-15 encountered serious structural resonance effects. Ground tests had uncovered a potential resonance between 

the pitch control system and a landing gear structural mode. Initially, 

researchers concluded that the effect was related to the ground-test equipment and its setup, and thus would not occur in flight. However, after 

several successful landings, the X-15 did experience a high-frequency 

vibration upon one touchdown. Additionally, a second and more severe 

structural resonance occurred at 13 Hz (coincident with the horizontal tail bending mode) during one entry from high altitude by the third 
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X-15 outfitted with the MH-96 adaptive flight control system.20 The pilot 

would first note a rumbling vibration that swiftly became louder. As the 

structure resonated, the vibrations were transmitted to the gyros in the 

flight control system, which attempted to “correct” for them but actually 

fed them instead. They were so severe that the pilot could not read the 

cockpit instruments and had to disengage the pitch damper in order to 

stop them. As a consequence, a 13 Hz “notch” filter was installed in the 

electrical feedback path to reduce the gain at the observed structural frequency. Thereafter, the third X-15 flew far more predictably21

Structural resonance problems are further complicated by the fact 

that the predicted structural frequencies are often in error, thus the flight 

control designers cannot accurately anticipate the proper filters for the 

sensors. Further, structural resonance is related to a structural feedback 

path, not an aerodynamic one as described for limit-cycles. As a precaution, ground vibration tests (GVT) are usually conducted on a new 

airplane to accurately determine the actual structural mode frequencies of the airplane.22 Researchers attach electrically driven and controlled actuators to various locations on the airplane and perform a small 

amplitude frequency “sweep” of the structure, essentially a “shake test.” 

Accelerometers at strategic locations on the airplane detect and record 

the structural response. Though this results in a more accurate determination of the actual structural frequencies for the control system designer, 

it still does not identify the structural path to the control system sensors. 

The flight control resonance characteristics can be duplicated on the 

ground by placing the airplane on a soft mounting structure, (airbags, 

or deflated tires and struts) then artificially raising the electrical gain in 

each flight control closed loop until a vibration is observed. Based on its 

experience with ground- and flight-testing of research airplanes, NASA 

DFRC established a ground rule that the flight gains could only be allowed 

to reach one-half of the gain that triggered a resonance (a gain margin of 

20. G.B. Merrick and L.W. Taylor, Jr., “X-15 Stability Augmentation System,” NASA Report H-271 

(Jan. 1961); L.W. Taylor, Jr., and J.W. Smith, “An Analysis of the Limit Cycle and Structural Resonance Characteristics of the X-15 Stability Augmentation System,” NASA TN-D-4287 (Dec. 1967). 

21. John P. Smith, Lawrence J. Schilling, and Charles A. Wagner, “Simulation at Dryden Flight 

Research Facility from 1957 to 1982,” NASA TM-101695 (1989), p. 4; Stillwell, X-15 Research 

Results, pp. 61–69. 

22. Weneth D. Painter and George J. Sitterle, “Ground and Flight Test Methods for Determining 

Limit Cycle and Structural Resonance Characteristics of Aircraft Stability Augmentation Systems,” 

NASA TN-D-6867 (June 1972). 
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2.0). This rule-of-thumb ground rule has been generally accepted within 

the aircraft industry, and ground tests to establish resonance gain margins are performed prior to first flights. If insufficient gain margin is present, the solution is sometimes a relocation of a sensor, or a stiffening of 

the sensor mounting structure. For most cases, the solution is the placement of an electronic notch filter within the control loop to reduce the 

system gain at the identified structural frequency. Many times the followup ground test identifies a second resonant frequency for a different 

structural mode that was masked during the first test. A typical notch filter will lower the gain at the selected notch frequency as desired but will 

also introduce additional lag at nearby frequencies. The additional lag 

will result in a lowering of the limit-cycle boundaries. The control system 

designer is thus faced with the task of reducing the gain at one structural 

frequency while minimizing any increase in the lag characteristics at the 

limit-cycle frequency (typically 2–4 Hz). This challenge resulted in the creation of lead-lag filters to minimize the additional lag in the system when 

notch filters were required to avoid structural resonance.23

Fighter aircraft usually are designed for 7–9 g load factors and, as a 

consequence, their structures are quite stiff, exhibiting high natural frequencies. Larger transport and reconnaissance airplanes are designed 

for much lower load factors, and the structures are more limber. Since 

structural frequencies are often only slightly above the natural aerodynamic frequencies—as well as the limit-cycle frequencies—of the 

airplane, this poses a challenge for the flight control system designer 

who is trying to aggressively control the aerodynamic characteristics, avoid limit-cycles, and avoid any control system response at the 

structural mode frequencies.

Structural mode interactions can occur across a range of flight activities. For example, Rockwell and Air Force testers detected a resonant 

vibration of the horizontal stabilizer during early taxi tests of the B-1 

bomber. It was traced to a landing gear structural mode, and a notch 

filter was installed to reduce the flight control gain at that frequency. 

The ground test for resonance is fairly simple, but the structural modes 

that need to be tested can produce a fairly large matrix of test conditions. External stores and fuel loadings can alter the structural frequencies of the airplane and thus change the control system feedback 

23. Ibid. 
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Aircraft frequency spectrum for flight control system design. USAF.

characteristics.24 The frequency of the wing torsion mode of the General 

Dynamics YF-16 Lightweight Fighter (the prototype of the F-16A Fighting 

Falcon) was dramatically altered when AIM-9 Sidewinder air-to-air missiles were mounted at the wingtip. 

The transformed dynamics of the installed missiles induced a serious aileron/wing-twist vibration at 6 Hz (coincident with the wing 

torsion mode), a motion that could also be classified as flutter, but in 

this case was obviously driven by the flight control system. Again, the 

solution was the installation of a notch filter to reduce the aileron 

response at 6 Hz.25

NASA researchers at the Dryden Flight Research Center had an 

unpleasant encounter with structural mode resonance during the 

Northrop–NASA HL-10 lifting body flight-test program. After an aborted 

launch attempt on the HL-10 lifting body, the NB-52B mother ship was 

returning with the HL-10 still mounted under the wing pylon. When the 

HL-10 pilot initiated propellant jettison, the launch airplane immediately experienced a violent vibration of the launch pylon attaching the 

lifting body to the NB-52B. The pilot stopped jettisoning and turned the 

flight control system off, whereupon the vibration stopped. The solution to this problem was strictly a change in operational procedure—in 

24. Personal experience as a member of the B-1 Flight Readiness Review Team.

25. Personal experience as a member of the Light Weight Fighter Joint Test Force.
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the future, the control system was to be disengaged before jettisoning 

while in captive flight.26

Fly-By-Wire: Fulfilling Promise and Navigating Around Nuance

As designers and flightcrews became more comfortable with electronic 

flight control systems and the systems became more reliable, the idea of 

removing the extra weight of the pilot’s mechanical control system began 

to emerge. Pilots resisted the idea because electrical systems do fail, and 

the pilots (especially military pilots) wanted a “get-me-home” capability. One flight-test program received little attention but contributed a 

great deal to the acceptance of fly-by-wire technology. The Air Force initiated a program to demonstrate that a properly designed fly-by-wire 

control system could be more reliable and survivable than a mechanical system. The F-4 Survivable Flight Control System (SFCS) program 

was initiated in the early 1970s. Many of the then-current accepted practices for flight control installations were revised to improve survivability. Four independent analog computer systems provided fail-op, fail-op 

(FOFO) redundancy. A self-adaptive gain changer was also included in 

the control logic (similar to the MH-96 in the X-15). Redundant computers, gyros, and accelerometers were eventually mounted in separate 

locations in the airplane, as were power supplies. Flight control system 

wire bundles for redundant channels were separated and routed through 

different parts of the airplane. Individual surface actuators (one aileron 

for example) could be operated to continue to maintain control when 

the opposite control surface was inoperative. The result was a flight 

control system that was lighter yet more robust than a mechanical system (which could be disabled by a single failure of a pushrod or cable). 

After development flight-testing of the SFCS airplane was completed, 

the standard F-4 mechanical backup system was removed, and the airplane was flown in a completely fly-by-wire configuration.27

The first production fly-by-wire airplane was the YF-16. It used four 

redundant analog computers with FOFO capability. The airplane was 

not only the first production aircraft to use FBW control, it was also the 

first airplane intentionally designed to be unstable in the pitch axis while 

26. Interview with John Manke, HL-10 test pilot. 

27. Maj. Robert Ettinger, Capt. Robert Majoros, and Lt. Col. Cecil W. Powell, “Air Force Evaluation of the Fly-By-Wire Portion of the Surviveable Flight Control System Advanced Development 

Program,” AFFTC TR-73-32 (Aug. 1973). 
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flying at subsonic speeds (“relaxed static stability”). The YF-16 prototype test program allowed the Air Force and General Dynamics to iron 

out the quirks of the FBW control system as well as the airplane aerodynamics before entering the full scale development of the F-16A/B. The 

high gains required for flying the unstable airplane resulted in some 

structural resonance and limit-cycle problems. The addition of external stores (tanks, bombs, and rockets) altered the structural mode frequencies and required fine-tuning of the control laws. Researchers and 

designers learned that flight control system design and aircraft interactions in the emergent FBW era were clearly far more complex and 

nuanced than control system design in the era of direct mechanical 

feedback and the augmented hydromechanical era that had followed.28

The Advent of Digital Flight Control Systems

Digital flight control systems were more nuanced still.29 Analog computers calculate solutions simultaneously, thus producing an instantaneous output for any input. Digital computers, although more precise 

than analog, calculate solutions in sequence, thus introducing a time 

delay between the input and the output, often referred to as “transport 

delay.” Early digital computers were far too slow to function in a realtime, flight control feedback system and could not compute the required 

servo commands fast enough to control the aircraft motions. As digital 

computation become faster and faster, control system designers gave 

serious attention to using them in aircraft flight control systems. NASA 

Dryden undertook the modification and flight-testing of a Vought F-8C 

Crusader Navy fighter to incorporate a digital fly-by-wire (DFBW) control 

system, based on the Apollo Guidance Computer used in the Apollo space 

capsule. The F-8 DFBW’s first flight was in 1972, and the test program 

completed 248 DFBW flights before its retirement at the end of 1985. 

It constituted a very bold and aggressive research program. The 

F-8 used redundant digital computers and was the first airplane relying 

solely on fly-by-wire technology for all of its flights. (Earlier FBW efforts, 

such as the AF F-4 Survivable Flight Control System, used a mechanical backup system for the first few flights.) NASA’s F-8 DFBW program 

28. Maj. James A. Eggers and Maj. William Bryant, Jr., “Flying Qualities Evaluation of the YF-16 

Prototype Light Weight Fighter,” AFFTC TR-75-15 (1975). 

29. The early advent of digital fly-by-wire is the subject of another case study in this volume (Piccirillo) and so is not examined in great detail here.
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not only set the stage for future military and civil digital flight control 

systems and fly-by-wire concepts, it also established the precedent for 

the operational procedures and built-in-test (BIT) requirements for this 

family of flight control systems.30 The ground-testing and general operating methods that were established by NASA DFRC in order to ensure 

safety of their F-8 DFBW airplane are still being used by most modern 

military and civilian airplanes.

After the completion of the basic digital FBW demonstration program, the F8 DFBW airplane was used for additional research testing, 

such as identifying the maximum allowable transport delay for a digital 

system to avoid pilot-induced oscillations. This is a key measurement in 

determining whether digital computations are fast enough to be used 

successfully in a control system. (The number turned out to be quite 

small, on the order of only 100 to 120 milliseconds.) The stimulus for 

this research was the PIO experienced by Shuttle pilot-astronaut Fred 

Haise during the fifth and last of the approach and landing tests flown 

at Edwards by the Space Shuttle orbiter Enterprise on October 26, 1977. 

Afterward, the Shuttle test team asked the DFBW test team if they could 

run in-flight simulations of the Shuttle using the F-8 DFBW testbed, to 

determine the effect of transport delays upon control response. During 

this follow-on research-testing phase, NASA Dryden Flight Research 

Center pilot John Manke experienced a dramatic, and very scary, landing. As he touched down, he added power to execute a “touch and go” 

to fly another landing pattern. But instead of climbing smoothly away, 

the F-8 began a series of violent pitching motions that Manke could not 

control. He disengaged the test system (which then reverted to a digital 

FBW version of the basic F-8 control system) just seconds before hitting the ground. The airplane returned to normal control, and the pilot 

landed safely. The culprit was an old set of control laws resident in the 

computer memory that had never been tested or removed. A momentary high pitch rate during the short ground roll had caused the airplane to automatically switch to these old control laws, which were later 

30. Dwain A. Deets and Kenneth J. Szalai, “Design and Flight Experience with a Digital Fly-By-Wire 

Control System in an F-8 Airplane,” NATO Advisory Group for Aeronautical Research and Development Conference Paper AGARD-CP-137 (1974); see also James E. Tomayko, Computers Take 

Flight: A History of NASA’s Pioneering Digital Fly-By-Wire Project, SP-2000-4224 (Washington, 

DC: NASA, 2000). 
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The Ling-Temco-Vought A-7D DIGITAC testbed was the first U.S. Air Force airplane with a digital flight 

control system. USAF.

determined to be unflyable.31 This event further reinforced the need for 

extensive validation and verification tests of all software used in digital flight control systems, no matter how expensive or time-consuming. 

In 1975, the Air Force began its own flight-testing of a digital flight 

control system, using a Ling-Temco-Vought A-7D Corsair II attack aircraft modified with a digital flight control system (dubbed DIGITAC) to 

duplicate the handling qualities of the analog Command Augmentation 

System of the baseline A-7D aircraft. As well, testers intended to evaluate several multimode features. 

The model-following system was enhanced to allow several models to be selected in flight. The objective was to determine if the pilot 

might desire a different model response during takeoff and landing, for 

example, than during air-to-air or air-to-ground gunnery maneuvers. The 

program was completed successfully in only 1 year of testing, primarily because the airplane was equipped with the standard A-7D mechanical backup system. The airplane used two digital computers that were 

continuously compared. If a disagreement occurred, the entire system 

would disengage, and the backup mechanical system was used to safely 

recover the airplane. The pilot also had a paddle switch on the stick that 

31. Interview with Manke; see also Tomayko, Computers Take Flight, pp. 111–114.
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immediately disconnected the digital system. This allowed software 

changes to be made quickly and safely and avoided most of the necessary, but time-consuming, preflight safety procedures that were associated with NASA’s F-8 DFBW program.32

One of the more challenging flight control system designs was associated with the Grumman X-29 research airplane. The X-29 was designed 

to demonstrate the advantages of a forward-swept wing (FSW), along 

with other new technologies. 

The airplane would fly with an unusually large level of pitch instability. The F-16, while flying at subsonic speeds, had a negative static margin of about 6 percent. The X-29 static margin was 35 percent unstable. 

(In practical terms, this meant that the divergence time to double amplitude was about half a second, effectively meaning that the airplane would 

destroy itself if it went out of control before the pilot could even recognize the problem!) This level of instability required extremely fast control 

surface actuators and state-of-the-art computer software. The primary 

system was a triplex of digital computers, each of which was backed 

up by an analog computer. A failure of one digital channel did not prevent the remaining two digital computers from continuing to function. 

After two digital channel failures, the system reverted to the three allanalog computers, thus maintaining fail-op, fail-op, fail-safe capability.

After completing the limit-cycle and resonance ground tests mentioned earlier, plus a lengthy software validation and verification effort, 

the flight-testing began in 1984 at NASA’s Dryden Flight Research Center.33

The control system handled the high level of instability quite well, and 

the test program on two airplanes was very successful, ending in 1992. 

Although the forward-swept wing concept has not been incorporated 

in any modern airplanes, the successful completion of the X-29 program further boosted the confidence in digital FBW control systems.34

In recent years, the digital FBW systems have become the norm in 

military aircraft. The later models of the F-15, F-16, and F/A-18 were 

32. Capt. Lawrence Damman, Capt. Ronald Grabe, Robert Kennington, and Paul W. Kirsten, 

“Flight Test Development of a Multimode Digital Flight Control System Implemented in an A-7D 

(DIGITAC),” AFFTC TR-76-15 (June 1976). 

33. Personal experience as a member of the X-29 Flight Readiness Review Team.

34. Paul Pellicano, Joseph Krumenacker, and David Van Hoy, “X-29 High Angle-of-Attack Flight Test 

Procedures, Results, and Lessons Learned,” Society of Flight Test Engineers 21st Annual Symposium, 

Aug. 1990.
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equipped with digital FBW flight control systems. The C-17 Globemaster 

III airlifter and F-117 Nighthawk stealth fighter performed their first 

flights with digital FBW systems. The Lockheed Martin F-22 Raptor 

and F-35 Lightning II Joint Strike Fighter exploit later digital FBW 

technology. Each has three digital computers and, for added safety, 

three of each critical component within its control systems. (Such 

“cross-strapping” of the various components allows FOFOFS redundancy.) There are dual-air data systems providing the various state variables that are backed up by an inertial system. The various “survivability” 

features first examined and demonstrated decades previously with the 

F-4 SFCS program (wire-routing, separate component locations, etc.) 

were also included in their basic design. 

Enhanced Electrical Actuators: Critical Enablers for FBW/DFBW

Nearly all high-speed airplanes use hydraulic actuators to operate the 

control surfaces. This provides a significant boost to the pilot’s ability 

to move a large control surface, which is experiencing very high aerodynamic loads. The computers and other electronic devices mentioned 

above merely provided signals to servos, which in turn commanded 

movement of hydraulic actuators. The hydraulic system provided the real 

muscle to move the surfaces. When Lockheed Martin’s “Skunk Works” 

was designing the planned X-33 Research Vehicle (intended to explore 

one possible design for a single-stage-to-orbit logistical spacecraft), 

keeping gross lift-off weight (GLOW) as low as possible was a crucial 

design goal. Because the hydraulic system would have been employed 

only during boost and entry, the entire hydraulic system would have 

been dead weight while the vehicle was in the space environment. Thus, 

control system designers elected to use electro-mechanical actuators to 

move the control surfaces, eliminating any need for a hydraulic system. 

Though X-33 was canceled for a variety of other reasons, its provision 

for electrical actuators clearly pointed toward future design practice. 

Following up on this were a series of three flight-test projects during 1997–1998 as part of the Electrically Powered Actuator Design 

(EPAD) program sponsored by the Naval Air Warfare Center and Air 

Force Research Laboratory. Each project tested a different advanced 

flight control actuator for the left aileron of NASA Dryden’s F/A-18 

Systems Research Aircraft (SRA). The first was the “smart actuator” 

that used fiber optics instead of the normal fly-by-wire system to con-
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trol an otherwise conventional hydraulic actuator.35 The next project 

flew an electro-hydrostatic actuator that used an electric motor to drive 

a small hydraulic pump that actuated the left aileron; the actuator was 

independent of the normal aircraft hydraulic system.36 The third project 

used an electro-mechanical actuator (EMA) that used electrical power 

generated by the F/A-18’s engines to power the left aileron actuator. A 

fiber-optic controller, self-contained control-surface actuator promises a 

significant reduction in weight and complexity over conventional actuation systems for future advanced air and space vehicles.37

Load Feedback for Flight Controls: Imitating the Birds

Among their many distinctive attributes, birds possess a particularly 

unique characteristic not experienced by humans: they are continuously aware of the loads their wings and control feathers bear, and they 

can adjust the wing shape to alleviate or redistribute these loads in real 

time. This allows a bird to optimize its wing shape across its entire 

range of flight; for example, a different wing shape for low-speed soaring than for high-speed cruising. Humans are not so fortunate. In the 

earliest days of flight, most aircraft designers consciously emulated the 

design of birds for both the planform and airfoil cross section of wings. 

Indeed, the frail fabric and wood structure of thin wings used by pioneers such as the Wright brothers, Louis Blériot, the Morane brothers, 

and Anthony Fokker permitted use of aeroelastic wing-warping (twisting) of a wing to bank an airplane, until superseded by the invention of 

the pivoted aileron. Naturally, when thicker wings appeared, the option 

of wing-warping became a thing of the past, not revived until the far 

later jet age and the era of thin composite structures. 

For human-created flight, structural loads can be measured via 

strain gages, and, indeed, the YF-16 utilized strain gages on the main 

wing spar to adjust the g limiter in the control laws for various fuel 

loadings and external store configurations. Though the system worked 

35. Eddie Zavala, “Fiber Optic Experience with the Smart Actuation System on the F-18 Systems 

Research Aircraft,” NASA TM-97-206223, Oct. 1997.

36. Robert Navarro, “Performance of an Electro-Hydrostatic Actuator on the F-18 Systems Research 

Aircraft,” NASA TM-97-206224, Oct. 1997. 

37. Joel R. Sitz, “F-18 Systems Research Aircraft,” NASA TM-4433 (1992); Lane E. Wallace, 

Flights of Discovery: 50 Years at the NASA Dryden Flight Research Center, SP-4309 (Washington, 

DC: NASA, 1996), pp. 124–125.
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and showed great promise, General Dynamics and the Air Force abandoned this approach for the production F-16 out of concern over the 

relatively low reliability of the strain gages. The technology still has not 

yet evolved to the point where designers are willing to put the strain 

gage outputs directly into the flight control system in a load-feedback 

manner.38 Certainly this technology will continue, and changing wing 

shapes based on load measurements will evolve. 

The NASA–Air Force Transonic Aircraft Technology (TACT) program, 

a joint cooperative effort from 1969 to 1988 between the Langley, Ames, 

and Dryden Centers, and the Air Force Flight Dynamics Laboratory, 

led to the first significant test of a so-called mission-adaptive wing 

(MAW), one blending a Langley-designed flexible supercritical wing 

planform joined to complex hydraulic mechanisms that could vary its 

shape in flight. Installed on an F-111A testbed, the MAW could “recontour” itself from a thick supercritical low-speed airfoil section suitable 

for transonic performance to a thinner symmetrical section ideal for 

supersonic flight.39 The MAW, a “first generation” approach to flexible 

skin and wing approaches, inspired follow-on work including tests by 

NASA Dryden on its Systems Research Aircraft, a McDonnell-Douglas 

(now Boeing) F/A-18B Hornet attack fighter using wing deformation as 

a means of achieving transonic and supersonic roll control.40

NASA DFRC is continuing its research on adaptive wing shapes 

and airfoils to improve efficiency in various flight environments. 

Thus, over a century after the Wrights first flew a bird-imitative wingwarping airplane at Kitty Hawk, wing-warping has returned to aeronautics, in a “back to the future—back to nature” technique used by 

the Wright brothers (and birds) to bank, and to perform turns. This 

cutting-edge technology is not yet in use on any operational airplanes, 

but it is only a matter of time before these performance enhancement 

features will increase the efficiency of future military and civilian aircraft.

38. Eggers and Bryant, “Flying Qualities Evaluation of the YF-16,” AFFTC TR-75-15 (1975).

39. Theodore G. Ayers and James B. Hallissy, “Historical Background and Design Evolution of the 

Transonic Aircraft Technology Supercritical Wing,” NASA TM-81356 (1981); Paul W. Phillips and 

Stephen B. Smith, “AFTI/F-111 Mission Adaptive Wing (MAW) Automatic Flight Control System 

Modes Lift and Drag Characteristics,” AFFTC TR-89-03 (1989).

40. Andrew M. Lizotte and Michael J. Allen, “Twist Model Development and Results From the 

Active Aeroelastic Wing F/A-18 Aircraft,” NASA TM-2005-212861 (2005); see also Chambers, 

Partners in Freedom, pp. 78–81.
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Structures and their Aeroelastic Manifestations 

Though an airplane looks rigidly solid, in fact it is a surprisingly flexible 

machine. The loadings it experiences in flight can manifest themselves 

in a variety of ways that affect and “move” the structure, and, as discussed previously, the flight control system itself can adversely affect the 

structure. The convoluted field in which aerodynamics and structures 

collide both statically and dynamically has led to some of the most complex and challenging problems that engineers, researchers, and designers have faced in the history of aeronautics. 

The safety factor for a railroad bridge is usually “10,” meaning that 

the structural members are sized to carry 10 times the design load without failing. Since weight is so crucial to the performance of an airplane, 

however, its structural safety factor is typically “1.5,” that is, the structure can fail if the loads are only 50 percent higher than the design value. 

As a result of the low aircraft design safety factor, aircraft structures 

receive far more attention during the design than do bridge structures 

and are subject to much larger deformations when loaded. This structural deformation can also interact with the aerodynamics of an airplane, both dynamically and statically, independently from the control 

system interaction mentioned earlier.

Flutter: The Insidious Threat

The most dramatic interaction of airplane structure with aerodynamics is “flutter”: a dynamic, high-frequency oscillation of some part of the 

structure. Aeroelastic flutter is a rapid, self-excited motion, potentially 

destructive to aircraft structures and control surfaces. It has been a particularly persistent problem since invention of the cantilever monoplane 

at the end of the First World War. The monoplane lacked the “bridge 

truss” rigidity found in the redundant structure of the externally braced 

biplane and, as it consisted of a single surface unsupported except at 

the wing root, was prone to aerodynamic induced flutter. The simplest 

example of flutter is a free-floating, hinged control surface at the trailing edge of a wing, such as an aileron. The control surface will begin 

to oscillate (flap, like the trailing edge of a flag) as the speed increases. 

Eventually the motion will feed back through the hinge, into the structure, and the entire wing will vibrate and eventually self-destruct. A 

similar situation can develop on a single fixed aerodynamic surface, 

like a wing or tail surface. When aerodynamic forces and moments are 

applied to the surface, the structure will respond by twisting or bending 
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about its elastic axis. Depending on the relationship between the elastic axis of the structure and the axis of the applied forces and moments, 

the motion can become self-energizing and a divergent vibration—one 

increasing in both frequency and amplitude—can follow. The high frequency and very rapid divergence of flutter causes it to be one of the 

most feared, and potentially catastrophic, events that can occur on an 

aircraft. Accordingly, extensive detailed flutter analyses are performed 

during the design of most modern aircraft using mathematical models of the structure and the aerodynamics. Flight tests are usually performed by temporarily fitting the aircraft with a flutter generator. This 

consists of an oscillating mass, or small vane, which can be controlled 

and driven at different frequencies and amplitudes to force an aerodynamic surface to vibrate. Instrumentation monitors and measures the 

natural damping characteristics of the structure when the flutter generator is suddenly turned off. In this way, the flutter mathematical model 

(frequency and damping) can be validated at flight conditions below the 

point of critical divergence.

Traditionally, if flight tests show that flutter margins are insufficient, operational limits are imposed, or structural beef-ups might be 

accomplished for extreme cases. But as electronic flight control technology advances, the prospect exists for so-called “active” suppression 

of flutter by using rapid, computer-directed control surface deflections. 

In the 1970s, NASA Langley undertook the first tests of such a system, 

on a one-seventeenth scale model of a proposed Boeing Supersonic 

Transport (SST) design, in the Langley Transonic Dynamics Tunnel 

(TDT). Encouraged, Center researchers followed this with TDT tests of 

a stores flutter suppression system on the model of the Northrop YF-17, 

in concert with the Air Force Flight Dynamics Laboratory (AFFDL, now 

the Air Force Research Laboratory’s Air Vehicles Directorate), later 

implementing a similar program on the General Dynamics YF-16. Then, 

NASA DFRC researchers modified a Ryan Firebee drone with such a 

system. This program, Drones for Aerodynamic and Structural Testing 

(DAST), used a Ryan BQM-34 Firebee II, an uncrewed aerial vehicle, 

rather than an inhabited system, because of the obvious risk to the pilot 

for such an experiment. 

The modified Firebee made two successful flights but then, in June 

1980, crashed on its third flight. Postflight analysis showed that one of 

the software gains had been inadvertently set three times higher than 

planned, causing the airplane wing to flutter explosively right after launch 
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A Drones for Aerodynamic and Structural Testing (DAST) unpiloted structural test vehicle, derived 

from the Ryan Firebee, during a 1980 flight test. NASA.

from the B-52 mother ship. In spite of the accident, progress was made 

in the definition of various control laws that could be used in the future 

for control and suppression of flutter.41 Overall, NASA research on active 

flutter suppression has been generally so encouraging that the fruits of it 

were applied to new aircraft designs, most notably in the “growth” version of the YF-17, the McDonnell-Douglas (now Boeing) F/A-18 Hornet 

strike fighter. It used an Active Oscillation Suppression (AOS) system 

to suppress flutter tendencies induced by its wing-mounted stores and 

wingtip Sidewinder missiles, inspired to a significant degree by earlier 

YF-17 and YF-16 Transonic Dynamics Tunnel testing.42

41. E. Nissim, “Design of Control Laws for Flutter Suppression Based on the Aerodynamic Energy 

Concept and Comparisons With Other Design Methods,” Technical Report TP-3056, Research 

Engineering, NASA Dryden Flight Research Center (1990) [given also as American Institute of 

Aeronautics and Astronautics Conference Paper 89-1212 (1989)]. 

42. J.T. Foughner, Jr., and C.T. Bensinger, “F-16 Flutter Model Studies With External Wing Stores,” 

NASA TM-74078 (1977); C. Hwang, E. Jonson, G. Mills, T. Noll, and M. Farmer, “Wind Tunnel 

Test of a Fighter Aircraft Wing/Store Flutter Suppression System: An International Effort,” AGARD 

R-689 (1980); R.P. Peloubet, Jr., and R.L. Haller, “Wind-Tunnel Demonstration of Actrive Flutter Suppression Using F-16 Model with Stores,” AFWAL TR-83-3046, vol. 1 (1983); Joseph R. Chambers, 

Innovation in Flight: Research of the NASA Langley Research Center on Revolutionary Advanced 

Concepts for Aeronautics, SP-2005-4539 (Washington, DC: NASA, 2005), pp. 196–203, 

212–215.
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Elastic Aerostructural Effects

The distortion of the shape of an airplane structure because of applied 

loads also creates a static aerodynamic interaction. When air loads are 

applied to an aerodynamic surface, it will bend or twist proportional to 

the applied load, just like a spring. Depending on the surface configuration, the distorted shape can produce different aerodynamic properties 

when compared with the rigid shape. A swept wing, for example, will 

bend upward at the tip and may also twist as it is loaded. 

This new shape may exhibit higher dihedral effect and altered spanwise lift distribution when compared with a rigid shape, impacting the 

performance of the aircraft. Because virtually all fighter aircraft have 

short wings and can withstand 7 to 9 g, their aeroelastic deformation 

is relatively small. In contrast, bomber, cargo, or high-altitude reconnaissance airplanes are typically designed for lower g levels, and 

the resulting structure, particularly its long, high aspect ratio wings, 

is often quite limber. 

Notice that this is not a dynamic, oscillatory event, but a static condition that alters the steady-state handling qualities of the airplane. The 

prediction of these aeroelastic effects is a complex and not altogether 

accurate process, though the trends are usually correct. Because the 

effect is a static condition, the boundaries for safe flight can usually be 

determined during the buildup flight-test program, and, if necessary, 

placards, can be applied to avoid serious incidents once the aircraft 

enters operational service. 

The six-engine Boeing B-47 Stratojet was the first airplane designed 

with a highly swept, relatively thin, high aspect ratio wing. At higher transonic Mach numbers, deflection of the ailerons would cause the wing to 

twist sufficiently to cancel, and eventually exceed, the rolling moment 

produced by the aileron, thus producing an aileron reversal. (In effect, 

the aileron was acting like a big trim tab, twisting the wing and causing 

the exact opposite of what the pilot intended.) Aerodynamic loads are 

proportional to dynamic pressure, so the aeroelastic effects are usually 

more pronounced at high airspeed and low altitude, and this combination caused several fatal accidents with the B-47 during its flight-testing 

and early deployment. After flight-testing determined the magnitude and 

region of reduced roll effectiveness, the airplane was placarded to 425 

knots to avoid roll reversal. In sum, then, an aeroelastic problem forced 

the limiting of the maximum performance achievable by the airplane, 

rendering it more vulnerable to enemy defenses. The B-47’s successor, 
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the B-52, had a much thicker wing root and more robust structure to 

avoid the problems its predecessor had encountered.

The Mach 3.2+ Lockheed SR-71 Blackbird, designed to cruise at 

supersonic speeds at very high altitude, was another aircraft that exhibited significant aeroelastic structural deformation.43 The Blackbird’s 

structure was quite limber, and the aeroelastic predictions for its behavior at cruise conditions were in error for the pitch axis. The SR-71 was 

a blended wing-body design with chines running along the forward 

sides of the fuselage and the engine nacelles, then blending smoothly 

into the rounded delta wing. These chines added lift to the airplane, and 

because they were well forward of the center of gravity, added a significant amount of pitching moment (much like a canard surface on an airplane such as the Wright Flyer or the Saab AJ-37 Viggen). Flight-testing 

revealed that the airplane required more “nose-up” elevon deflection at 

cruise than predicted, adding a substantial amount of trim drag. This 

reduced the range the Blackbird could attain, degrading its operational 

performance. To correct the problem, a small shim was added to the 

production fuselage break just forward of the cockpit. The shim tilted 

the forebody nose cone and its attached chine surfaces slightly upward, 

producing a nose-up pitching moment. This allowed the elevons to be 

returned to their trim faired position at cruise flight conditions, thus 

regaining the lost range capability. 

Sadly, the missed prediction of the aeroelastic effects also contributed to the loss of one of the early SR-71s. While the nose cone 

forebody shim was being designed and manufactured, the contractor 

desired to demonstrate that the airplane could attain its desired range 

if the elevons were faired. To achieve this, Lockheed technicians added 

trim-altering ballast to the third production SR-71, then being used 

for systems and sensor testing. The ballast shifted the center of gravity about 2 percent aft from its normal position and at the aft design 

limit for the airplane. The engineers calculated that this would permit 

the elevons to be set in their faired position at cruise conditions for 

this one flight so that the SR-71 could meet its desired range performance. Instead, the aft cg, combined with the nonlinear aerodynamics 

43. For perspectives on the various members of the Blackbird family, see Peter W. Merlin, From 

Archangel to Senior Crown: Design and Development of the Blackbird, (Reston, VA: American 

Institute for Aeronautics and Astronautics, 2008); and also his Mach 3+: NASA/USAF YF-12 Flight 

Research, 1969–1979, SP-2001-4525 (Washington, DC: NASA, 2001).
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and aeroelastic bending of the fuselage, resulted in the airplane going 

out of control at the start of a turn at a cruise Mach number. The airplane broke in half, catapulting the pilot, who survived, from the cockpit. Unfortunately, his flight-test engineer/navigator perished.44 Shim 

installation, together with other minor changes to the control system 

and engine inlets, subsequently enabled the SR-71 to meet its performance goals, and it became a mainstay of America’s national reconnaissance fleet until its retirement in early 1990.

Lockheed, the Air Force, and NASA continued to study Blackbird 

aeroelastic dynamics. In 1970, Lockheed proposed installation of a 

Loads Alleviation and Mode Suppression (LAMS) system on the YF-12A, 

installing very small canards called “exciter-” or “shaker-vanes” on the 

forebody to induce in-flight motions and subsequent suppression techniques that could be compared with analytical models, particularly 

NASA’s NASTRAN and Boeing’s FLEXSTAB computerized load prediction and response tools. The LAMS testing complemented Air ForceNASA research on other canard-configured aircraft such as the Mach 3+ 

North American XB-70A Valkyrie, a surrogate for large transport-sized 

supersonic cruise aircraft. The fruits of this research could be found on 

the trim canards used on the Rockwell B-1A and B-1B strategic bombers, which entered service in the late 1980s and notably improved their 

high-speed “on the deck” ride qualities, compared with their three lowaltitude predecessors, the Boeing B-52 Stratofortress, Convair B-58 

Hustler, and General Dynamics FB-111.45

The Advent of Fixed-Base Simulation

Simulating flight has been an important part of aviation research since 

even before the Wright brothers. The wind tunnel, invented in the 1870s, 

represented one means of simulating flight conditions. The rudimentary 

Link trainer of the Second World War, although it did not attempt to 

represent any particular airplane, was used to train pilots on the proper 

navigation techniques to use while flying in the clouds. Toward the end 

of the Second World War, researchers within Government, the military 

services, academia, and private industry began experimenting with 

44. Personal experience during SR-71 accident investigation; Ben R. Rich and Leo Janos, Skunk 

Works: A Personal Memoir of My Years of Lockheed (Boston: Little, Brown, and Co., 1994), pp. 

192–237. 

45. Merlin, Mach 3+, pp. 39–42.
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analog computers to solve differential equations in real time. Electronic 

components, such as amplifiers, resistors, capacitors, and servos, were 

linked together to perform mathematical operations, such as arithmetic and integration. By patching many of these components together, it 

was possible to continuously solve the equations of motion for a moving 

object. There are six differential equations that can be used to describe 

the motion of an object. Three rotational equations identify pitching, 

rolling, and yawing motions, and three translational equations identify 

linear motion in fore-and-aft, sideways, and up-and-down directions. 

Each of these equations requires two independent integration processes 

to solve for the vehicle velocities and positions. Prior to the advent of 

analog computers, the integration process was a very tedious, manual 

operation and not amenable to real-time solutions. Analog computers 

allowed the integration to be accomplished in real time, opening the 

door to pilot-in-the-loop simulation. The next step was the addition of 

controlling inputs from an operator (stick and rudder pedals) and output 

displays (dials and oscilloscopes) to permit continuous, real-time control of a simulated moving object. Early simulations only solved three 

of the equations of motion, usually pitch rotation and the horizontal 

and vertical translational equations, neglecting some of the minor coupling terms that linked all six equations. As analog computers became 

more available and affordable, the simulation capabilities expanded 

to include five and eventually all six of the equations of motion (commonly referred to as “six degrees of freedom” or 6DOF).

By the mid-1950s, the Air Force, on NACA advice, had acquired a 

Goodyear Electronic Differential Analyzer (GEDA) to predict aircraft 

handling qualities based on the extrapolation of data acquired from 

previous test flights. One of the first practical applications of simulation was the analysis of the F-100A roll-coupling accident that killed 

North American Aviation (NAA) test pilot George “Wheaties” Welch on 

October 12, 1954, one of six similar accidents that triggered an emergency grounding of the Super Sabre. By programming the pilot’s inputs 

into a set of equations of motion representing the F-100A, researchers 

duplicated the circumstances of the accident. The combination of simulation and flight-testing on another F-100A at the NACA High-Speed 

Flight Station (now the Dryden Center) forced redesign of the aircraft. 

North American increased the size of the vertical fin by 10 percent and, 

when even this proved insufficient, increased it again by nearly 30 percent, modifying existing and new production Super Sabres with the 
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larger tail. Thus modified, the F-100 went on to a very successful career 

as a mainstay Air Force fighter-bomber.46

Another early application of computerized simulation analysis 

occurred during the Air Force-NACA X-2 research airplane program in 

1956. NACA engineer Richard E. Day established a simulation of the 

X-2 on the Air Force’s GEDA analog computer. He used a B-17 bombardier’s stick as an input control and a simple oscilloscope with a line 

representing the horizon as a display along with some voltmeters for 

airspeed, angle of attack, etc. Although the controls and display were 

crude, the simulation did accurately duplicate the motions of the airplane. Day learned that lateral control inputs near Mach 3 could result 

in a roll reversal and loss of control. He showed these characteristics to 

Capt. Iven Kincheloe on the simulator before his flight to 126,200 feet 

on September 7, 1956. When the rocket engine quit near Mach 3, the airplane was climbing steeply but was in a 45-degree bank. Kincheloe remembered the simulation results and did not attempt to right the airplane with 

lateral controls until well into the entry at a lower Mach number, thus avoiding the potentially disastrous coupled motion observed on the simulator.47

Kincheloe’s successor as X-2 project pilot, Capt. Milburn Apt, also 

flew the simulator before his ill-fated high-speed flight in the X-2 on 

September 27, 1956. When the engine exhausted its propellants, Apt 

was at Mach 3.2 and over 65,000 feet, heading away from Edwards and 

apparently concerned that the speeding plane would be unable to turn 

and glide home to its planned landing on Rodgers Dry Lake. When he 

used the lateral controls to begin a gradual turn back toward the base, 

46. Marcelle Size Knaack, Post-World War II Fighters, vol. 1 of Encyclopedia of U.S. Air Force 

Aircraft and Missile Systems (Washington, DC: Office of Air Force History, 1978), pp. 114–116; 

Bill Gunston, Early Supersonic Fighters of the West (New York: Charles Scribner’s Sons, 1975), pp. 

153–157; HSFS, “Flight Experience With Two High-Speed Airplanes Having Violent Lateral-Longitudinal Coupling in Aileron Rolls,” RM H55A13 (1955); Hubert M. Drake and Wendell H. Stillwell, 

“Behavior of the Bell X-1A Research Airplane During Exploratory Flights at Mach Numbers Near 2.0 

and at Extreme Altitudes,” RM H55G25 (1955); Hubert M. Drake, Thomas W. Finch, and James 

R. Peele, “Flight Measurements of Directional Stability to a Mach Number of 1.48 for an Airplane 

Tested with Three Different Vertical Tail Configurations,” RM H55G26 (1955).

47. Hubert M. Drake and Wendell H. Stillwell, “Behavior of the Bell X-1A Research Airplane During 

Exploratory Flights at Mach Numbers Near 2.0 and at Extreme Altitudes,” RM H55G25 (1955); 

Capt. Iven C. Kincheloe, USAF, “Flight Research at High Altitude, Part II,” in Proceedings of the 

Seventh AGARD General Assembly, Nov. 18–26, 1957 (Washington, DC: NATO Advisory Group 

for Aeronautical Research and Development, 1958).
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the X-2 went out of control. Apt was badly battered in the violent motions 

that ensued, was unable to use his personal parachute, and was killed.48

The loss of the X-2 and Apt shocked the Edwards community. The 

accident could be duplicated on the simulator, solidifying the value of 

simulation in the field of aviation and particularly flight-testing.49 The 

X-2 experience convinced the NACA (later NASA) that simulation must 

play a significant role in the forthcoming X-15 hypersonic research aircraft program. The industry responded to the need with larger and more 

capable analog computer equipment.50

The X-15 simulator constituted a significant step in both simulator 

design and flight-test practice. It consisted of several analog computers 

connected to a fixed-base cockpit replicating that of the aircraft, and an 

“iron bird” duplication of all control system hardware (hydraulic actuators, cable runs, control surface mass balances, etc.). Computer output 

parameters were displayed on the normal cockpit instruments, though 

there were no visual displays outside the cockpit. This simulator was first 

used at the North American plant in Inglewood, CA, during the design 

and manufacture of the airplane. It was later transferred to NASA DFRC 

at Edwards AFB and became the primary tool used by the X-15 test team 

for mission planning, pilot training, and emergency procedure definition.

The high g environment and the high pilot workload during the 

10-minute X-15 flights required that the pilot and the operational support team in the control room be intimately familiar with each flight plan. 

There was no time to communicate emergency procedures if an emergency 

occurred—they had to be already imbedded in the memories of the pilot 

and team members. That necessity highlighted another issue underscored 

by the X-15’s simulator experience: the necessity of replicating with great 

fidelity the actual cockpit layout and instrumentation in the simulator. On 

at least two occasions, X-15 pilots nearly misread their instrumentation 

or reached for the wrong switch because of seemingly minor differences 

between the simulator and the instrumentation layout of the X-15 aircraft.51

48. Bell X-2 No. 1 Accident Report, copy in History Office archives, Air Force Flight Test Center, 

Edwards AFB, CA.

49. Ronald Bel Stiffler, The Bell X-2 Rocket Research Aircraft: The Flight Test Program (Edwards 

AFB: Air Force Flight Test Center, Aug. 12, 1957), p. 87; Richard E. Day, “Coupling Dynamics in 

Aircraft: A Historical Perspective,” SP-532 (1997). 

50. Smith, Schilling, and Wagner, “Simulation at Dryden,” p. 1.

51. Ibid., p. 3.
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Overall, test pilots and flight-test engineers uniformly agreed that 

the X-15 program could not have been accomplished safely or productively without the use of the simulator. Once the X-15 began flying, 

engineers updated the simulator using data extracted from actual flight 

experience, steadily refining and increasing its fidelity. An X-15 pilot 

“flew” the simulator an average of 15 hours for every flight, roughly 1 

hour of simulation for every minute of flying time. The X-15 experience 

emphasized the profound value of simulation, and soon, nearly all new 

airplanes and spacecraft were accompanied by fixed-base simulators 

for engineering analysis and pilot/astronaut training.

Updating Simulator Prediction with Flight-Test Experience

Test pilots who “flew” the early simulators were skeptical of the results 

that they observed, because there was usually some aspect of the simulation that did not match the real airplane. Stick forces and control 

surface hinge moments were often not properly matched on the simulator, and thus the apparent effectiveness of the ailerons or elevators 

was often higher or lower than experienced with the airplane. For procedural trainers (used for checking out pilots in new airplanes) mathematical models were often changed erroneously based strictly on pilot 

comments, such as “the airplane rolls faster than the simulator.” Since 

these early simulators were based strictly on wind tunnel or theoretical 

aerodynamic predictions and calculated moments of inertia, the flighttest community began to explore methods for measuring and validating 

the mathematical models to improve the acceptance of simulators as valid 

tools for analysis and training. Ground procedures and support equipment were devised by NASA to measure the moments of inertia of small 

aircraft and were used for many of the research airplanes flown at DFRC.52

A large inertia table was constructed in the Air Force Flight Test 

Center Weight and Balance facility at Edwards AFB for the purpose of 

measuring the inertia of large airplanes. Unfortunately, the system was 

never able to provide accurate results, as fluctuations in temperature 

and humidity adversely affected the performance of the table’s sensitive 

bearings, so the concept was discarded.

During the X-15 flight-test program, NASA researchers at Edwards 

developed several methods for extracting the aerodynamic stability 

52. Capt. John Retelle, “Measured Weight, Balance, and Moments of Inertia of the X-24A Lifting 

Body,” AFFTC TD-71-6 (1971). 
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derivatives from specific flight-test maneuvers. Researchers then compared these results with wind tunnel or theoretical predictions and, 

where necessary, revised the simulator mathematical models to reflect 

the flight-test-derived information. For the X-15, the predictions were 

quite good, and only minor simulator corrections were needed to allow 

flight maneuvers to be replicated quite accurately on the simulator. 

The most useful of these methods was an automatic computer analysis of pulse-type maneuvers, originally referred to as Newton-Raphson 

Parameter Identification.53,54 This system evolved into a very useful tool 

subsequently used as an industry standard for identifying the real-world 

stability and control derivatives during early testing of new aircraft.55

The resulting updates are usually also transplanted into the final training simulators to provide the pilots with the best possible duplication of 

the airplanes’ handling qualities. Bookkeeping methods for determining moments of inertia of a new aircraft (i.e., tracking the weight and 

location of each individual component or structural member during aircraft manufacture) have also been given more attention.

Characteristically, the predicted aerodynamics for a new airplane 

are often in error for at least a few of the derivatives. These errors are 

usually a result of either a discrepancy between the wind tunnel model 

that was tested and the actual airplane that was manufactured, or a 

result of a misinterpretation or poor interpolation of the wind tunnel 

data. In some cases, these discrepancies have been significant and have 

led to major incidents (such as the HL-10 first flight described earlier). 

Another source of prediction errors for simulation is the prediction of 

the aeroelastic effects from applied air loads to the structure. These 

aeroelastic effects are quite complex and difficult to predict for a limber airplane. They usually require flight-test maneuvers to identify or 

validate the actual handling quality effects of structural deformation. 

There have been several small, business aircraft that have been built, 

developed, and sold commercially wherein calculated predictions of 

the aerodynamics were the primary data source, and very little if any 

wind tunnel tests were ever accomplished. Accurate simulators for pilot 

53. K.W. Iliff, B.G. Powers, and L.W. Taylor, Jr., “A Comparison of Newton-Raphson and Other 

Methods for Determining Stability Derivatives from Flight Data,” NASA Report H-544 (Mar. 1969). 

54. K.W. Iliff and L.W. Taylor, Jr., “Determination of Stability Derivatives from Flight Data Using a 

Newton-Raphson Minimization Technique,” NASA TN-D-6579 (Mar. 1972).

55. Kenneth W. Iliff, “Aircraft Parameter Estimation,” AIAA Meeting Paper 1987-0623 (1987). 
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training have been created by conducting a brief flight test of each airplane, performing required test maneuvers, then applying the flight-test 

parameter estimation methods developed by NASA. With a little bit of 

attention during the flight-test program, a highly accurate mathematical 

model of a new airplane can be assembled and used to produce excellent 

simulators, even without wind tunnel data.56

Moving Base Cockpit and Centrifuge Simulators

As the computational capability to accurately model the handling qualities of an airplane improved, there was recognition that the lack of motion 

cues was a distraction to the realism of the simulation. An early attempt 

to simulate motion for the pilot consisted of mounting the entire simulator cockpit on a set of large hydraulic actuators. These actuators would 

generate a small positive or negative bump to simulate g onset, while any 

steady-state acceleration was washed out over time (i.e., back to 1 g). The 

actuators could also tilt the simulator cockpit to produce a side force, 

or fore and aft force, on the crew. When correlated with a horizon on a 

visual screen, the result was a quite realistic sensation of motion. These 

moving-base cockpit systems were rather expensive and difficult to maintain compared with a simple fixed-base cockpit. Since both the magnitude of the g vector and the rotational motion required were false, the 

systems were not widely accepted in the flight-testing community, where 

the goal is to evaluate the pilot’s response and capabilities in a true flight 

environment. They found ready acceptance as airline procedures trainers when the maneuvers are slow and g forces are typically small and 

proved a source of entertainment in amusement parks, aerospace museums, and science centers. 

In the 1950s, the Naval Air Development Center (NADC) at Johnsville, 

PA, developed a large, powerful centrifuge to explore human tolerance 

to high g forces. The centrifuge consisted of a 182-ton electric DC motor 

turning a 50-foot arm with a gondola at the end of the arm. The motor 

could generate g forces at the gondola as high as 40 g’s. The gondola was 

mounted with two controllable gimbals that allowed the g vector to be 

oriented in different directions for the gondola occupant.57

56. David L Kohlman, William G. Schweikhard, and Donald R.L Renz, “Advances in Flight Test 

Instrumentation and Analysis” SAE Doc. No. 871802, Oct. 1987.

57. C.C. Clark and C.H. Woodling, “Centrifuge Simulation of the X-15 Research Aircraft,” NADC 

MA-5916 (1959).
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Test pilot entering the centrifuge gondola at the Naval Air Development Center (NADC) in 

Johnsville, PA. NASA.

Many detailed studies defining human tolerance to g forces were performed on the centrifuge using programmed g profiles. NADC devised a 

method for installing a cockpit in the gondola, connecting it to a large 

analog computer, and allowing the pilot to control the computer simulation, which in turn controlled the centrifuge rotation rate and gimbal angles. This allowed the pilot in the gondola to not only see the pilot 

displays of the simulated flight, but also to feel the associated translational g levels in all three axes. Although the translational g forces were 

correctly simulated, the gimbal rotations necessary to properly align the 

total g vector with the cockpit were artificial and were not representative of a flight environment. 

One of the first applications of this closed-loop, moving base simulation was in support of the X-15 program in 1958. There were two 

prime objectives of the X-15 centrifuge program associated with the 

high g exit and entry: assessment and validation of the crew station (side 

arm controller, head and arm restraints, displays, etc.), and evaluation 

of the handling qualities with and without the Stability Augmentation 

System. The g environment during exit consisted of a forward acceleration (eyeballs-in) increasing from 2 to 4 g, combined with a 2 g pullup 

(eyeballs-down). The entry g environment was more severe, consisting 
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of a deceleration (eyeballs-out) of 3 g combined with a simultaneous 

pullout acceleration of 6 g (eyeballs-down). 

The results of the X-15 centrifuge program were quite useful to the 

X-15’s overall development; however, the pilots felt that the centrifuge did 

not provide a very realistic simulation of an aircraft flight environment. The 

false rotational movement of the gondola was apparent to the pilots and 

was a distraction to the piloting task during entry. The exit phase of an X-15 

flight was a fairly steady acceleration with little rotational motion, and the 

pilots judged the simulation a good representation of that environment.58

The NADC centrifuge was also used in support of the launch phase 

of the Mercury, Gemini, and Apollo space programs. These provided 

valuable information regarding the physiological condition of the 

astronauts and the crew station design but generally did not include 

closed-loop piloting tasks with the pilot controlling the simulated 

vehicle and trajectory. 

A second closed-loop centrifuge simulation was performed in support of the Boeing X-20 Dyna-Soar program. Dyna-Soar constituted an 

ambitious but feasible Air Force effort to develop a hypersonic lofted 

boost-glider capable of an orbital flight. Unfortunately, it was prematurely 

canceled in 1963 by then-Secretary of Defense Robert S. McNamara. 

The Dyna-Soar centrifuge study effort was similar to the X-15 centrifuge 

program, but the acceleration lasted considerably longer and peaked 

at 6 g (eyeballs-in) at burnout of the Titan III booster. The pilots were 

“flying” the vehicle in all three axes during these centrifuge runs, and 

valuable data were obtained relative to the pilot’s ability to function effectively during long periods of acceleration. Some of the piloting demonstrations included alleviating wind spikes during the early ascent phase 

and successfully guiding the booster to an accurate orbital insertion 

using simple backup guidance concepts in the event of a booster guidance failure.59 The Mercury and Gemini programs used automatic guidance during the ascent phase, and the only piloting task during boost 

was to initiate an abort by firing the escape rockets. The Apollo program 

included a backup piloting mode during the boost based on the results 

of the X-20 and other centrifuge programs. 

58. Personal recollections as a flight planning engineer participating in the X-15 centrifuge program. 

Also see Dennis Jenkins, X-15: Extending the Frontiers of Flight. 

59. Robert G. Hoey, Lt. Col. Harry R. Bratt, and Maj. Russell L. Rogers, “A Dynamic Simulation of 

Pilot Controlled Boost for the X-20A Air Vehicle,” AFFTC TDR-63-21 (1964). 
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Variable Stability Airplanes

Although the centrifuge was effective in simulating relatively steady high 

g accelerations, it lacked realism with respect to normal aircraft motions. 

There was even concern that some amount of negative training might be 

occurring in a centrifuge. One possible method of improving the fidelity of 

motion simulation was to install the entire simulation (computational mathematical model, cockpit displays, and controls) in an airplane, then forcing the airplane to reproduce the flight motions of the simulated airplane, 

thus exposing the simulator pilot to the correct motion environment. An 

airplane so equipped is usually referred to as a “variable stability aircraft.”

Since their invention, variable stability aircraft have played a significant role in advancing flight technology. Beginning in 1948, the Cornell 

Aeronautical Laboratory (now Calspan) undertook pioneering work on 

variable stability using conventional aircraft modified in such a fashion 

that their dynamic characteristics reasonably approximated those of different kinds of designs. Waldemar Breuhaus supervised modification of 

a Vought F4U-5 Corsair fighter as a variable stability testbed. From this 

sprang a wide range of subsequent “v-stab” testbeds. NACA Ames researchers modified another Navy fighter, a Grumman F6F-5 Hellcat, so that it 

could fly as if its wing were set at a variety of dihedral angles; this research, 

and that of a later North American F-86 Sabre jet fighter likewise modified 

for v-stab research, was applied to design of early Century series fighters, 

among them the Lockheed F-104 Starfighter, a design with pronounced 

anhedral (negative wing dihedral).60

As the analog simulation capability was evolving, Cornell researchers developed a concept of installing a simulator in one cockpit of a 

60. Edwin P. Hartman, Adventures in Research: A History of Ames Research Center, 1940–1965, 

SP-4302 (Washington, DC: NASA, 1970), pp. 164–166; 257–258; Paul F. Borchers, James A. 

Franklin, Jay W. Fletcher, Flight Research at Ames: Fifty-Seven Years of Development and Validation 

of Aeronautical Technology, SP-3300 (Washington, DC: NASA, 1998), passim; William M. Kauffman, Charles J. Liddell, Jr., G. Allan Smith, and Rudolph D. Van Dyke, Jr., “An Apparatus for Varying 

Effective Dihedral in Flight with Application to a Study of Tolerable Dihedral on a Conventional 

Fighter Airplane,” NACA Report 948 (1949); Walter E. McNeill and Brent Y. Creer, “A Summary 

of Results Obtained during Flight Simulation of Several Aircraft Prototypes with Variable Stability Airplanes,” NACA RM-A56C08 (1956); Richard F. Vomaske, Melvin Sadoff, and Fred J. Drinkwater, 

III, “The Effect of Lateral-Directional Control Coupling on Pilot Control of an Airplane as Determined 

in Flight and a Fixed-Base Flight Simulator,” NASA TN-D-1141 (1961); William M. Kauffman 

and Fred J. Drinkwater, III, “Variable Stability Airplanes in Lateral Stability Research,” Aeronautical 

Engineering Review, vol. 14, No. 8 (Aug. 1955), pp. 29–30.
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two-seat Lockheed NT-33A Shooting Star aircraft. By carefully measuring the stability and controllability characteristics of the “T-Bird” and 

then subtracting those characteristics from the simulated mathematical model, the researchers could program the airplane with a completely 

different dataset that would effectively represent a different airplane.61

Initially the variable stability feature was used to perform general 

research tests by changing various controlled variables and evaluating 

their effect on pilot performance. Eventually mathematical models were 

introduced that represented the complete predicted aerodynamic and 

control system characteristics of new designs. The NT-33A became the 

most-recognized variable-stability testbed in the world, having “modeled” 

aircraft as diverse as the X-15, the B-1 bomber, and the Rockwell Space 

Shuttle orbiter, and flying from the early 1950s until retirement after 

the end of the Cold War. Thanks to its contributions and those of other 

v-stab testbeds developed subsequently,62 engineers and pilots have had 

a greater understanding of anticipated flying qualities and performance 

of new aircraft before the crucial first flight.63 In particular, the variable 

stability aircraft did not exhibit the false rotations associated with the 

centrifuge simulation and were thus more realistic in simulating rapid 

aircraft-like maneuvers. Several YF-22 control law variations were tested 

using the CALSPAN NT-33 prior to the first flight. Before the first flight of 

the F-22, the control laws were tested on the CALSPAN VISTA. Today it 

is inconceivable that a new aircraft would fly before researchers had first 

evaluated its anticipated handling qualities via variable-stability research.

Low L/D Approach and Landing Trainers

In addition to the need to simulate the handling qualities of a new airplane, a need to accurately duplicate the approach and landing performance also evolved. The air-launched, rocket-powered research airplane 

concept, pioneered by the X-1, allowed quick access to high-speed flight 

61. G. Warren Hall, “Research and Development History of USAF Stability T-33,” Journal of the 

American Aviation Historical Society, vol. 19, no. 4 (winter 1974).

62. Mostly notably of these were a North American JF-100C Super Sabre (another Ames project), 

a Martin-Air Force v-stab Convair F-106 Delta Dart; the NASA FRC General Purpose Airborne 

Simulator (a modified Lockheed Jetstar executive jet transport); the CALSPAN–Air Force Convair 

NC-131H Total In-Flight Simulator (TIFS), retired in late 2008; the CALSPAN variable stability 

Douglas B-26 Invader; its successor, the CALSPAN v-stab Learjet; and the most recent, the CALSPAN VISTA Lockheed Martin NF-16.

63. Shafer, “In-Flight Simulation Studies at the NASA Dryden Flight Research Facility.”
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for research purposes. It also brought with it unpowered, gliding landings, after the rocket fuel was expended. For the X-1 series of airplanes, 

the landings were not particular stressful because most landings were on 

the 7-mile dry lakebed at Edwards AFB and the approach glide angles 

were 8 degrees or less (lift-to-drag (L/D) ratios of about 8). As the rocketpowered airplanes reached toward higher speeds and altitudes, the landing 

approach angles increased rather dramatically. The approach glide angle 

for the X-15 was predicted to be between 15 and 20 degrees (lift-to-drag 

ratios between 2.8 and 4.25) primarily because of the larger base area at 

the rear of the fuselage. The L/D was further reduced to about 2.5 after 

landing gear and flap deployment. These steep unpowered approaches 

prompted a reassessment of the piloting technique to be used. Higherthan-normal approach speeds were suggested as well as a delay of the landing gear and flap deployment until after completion of the landing flare. 

These new landing methods also indicated a need for a training “simulator” that could duplicate the landing performance of the X-15 in order to 

explore different landing techniques and train test pilots.

Out-of-the-cockpit, simulated visual displays available at that time 

were of very poor quality and were not even considered for the X-15 

fixed-base simulator. Simulated missions on the X-15 fixed-base simulator were flown to a high-key location over the lakebed using the cockpit 

instruments, but the simulation was not considered valid for the landing 

pattern or the actual landing, which was to be done using visual, out-ofthe-window references. 

North American added a small drag chute to one of its F-100s to allow 

its pilots to fly landing approaches simulating the X-15. Additionally, both 

the Air Force and NASA began to survey available jet aircraft that could 

match the expected X-15 landing maneuver so that the Government pilots 

could develop a consistent landing method and identify what external 

cues were necessary to perform accurate landings. The F-104 had just 

entered the inventory at the AFFTC and NASA. Flight-testing showed that 

it was an excellent candidate for duplicating the X-15 landing pattern.64

64. Gene J. Matranga and Neil A. Armstrong, “Approach and Landing Investigation at Lift-Drag 

Ratios of 2 to 4 Utilizing a Straight-Wing Fighter Airplane,” NASA TM-X-31 (1959); Gene J. 

Matranga and Neil A. Armstrong, “Approach and Landing Investigation at Lift-Drag Ratios of 2 to 4 

Utilizing a Delta-Wing Fighter Airplane,” NASA TM-X-125 (1959); Stillwell, X-15 Research Results, 

pp. 38–39; Milton O. Thompson, At the Edge of Space: The X-15 Flight Program (Washington: 

Smithsonian Institution Press, 1992).
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Various combinations of landing gear and flap settings, plus partial 

power on the engine, could be used to simulate the entire X-15 landing trajectory from high key to touchdown. F-104s were used throughout the program for chase, for training new X-15 pilots, for practicing 

approaches prior to each flight, and also for practicing approaches into 

uprange emergency lakebeds. The combination of the X-15 fixed-base 

simulator and the F-104 in-flight landing simulation worked very well 

for pilot training and emergency planning over the entire X-15 test program, and the F-104 did yeoman work supporting the subsequent lifting body research effort as well, through the X-24B. 

In the late 1960s, engineers at the Air Force Flight Dynamics 

Laboratory had evolved a family of reentry shapes (particularly the 

AFFDL 5, 7, and 8) that blended a lifting body approach with an extensible variable-sweep wing for terminal approach and landing. In support 

of these studies, in 1969, the Air Force Flight Test Center undertook a 

series of low L/D approach tests using a General Dynamics F-111A as a 

surrogate for a variable-sweep Space Shuttle-like craft returning from 

orbit. The supersonic variable-sweep F-111 could emulate the track of 

such a design from Mach 2 and 50,000 feet down to landing, and its 

sophisticated navigation system and two-crew-member layout enabled 

a flight-test engineer/navigator to undertake terminal area navigation. 

The result of these tests demonstrated conclusively that a trained crew 

could fly unpowered instrument approaches from Mach 2 and 50,000 

feet down to a precise runway landing, even at night, an important 

confidence-building milestone on the path to the development of practical lifting reentry logistical spacecraft such as the Shuttle.65

Notice that the landing-pattern simulators discussed above did not 

duplicate the handling qualities of the simulated airplane, only the performance and landing trajectory. Early in the Space Shuttle program, management decided to create a Shuttle Training Aircraft (STA). A Grumman 

G II was selected as the host airplane. Modifications were made to this 

unique airplane to not only duplicate the orbiter’s handling qualities (a 

variable-stability airplane), but also to duplicate the landing trajectory 

and the out-of-the-window visibility from the orbiter cockpit. This NASA 

training device represents the ultimate in a complete electronic and 

65. B.L. Schofield, D.F. Richardson, and P.C. Hoag, “Terminal Area Energy Management, 

Approach, and Landing Investigation for Maneuvering Reentry Vehicles using F-111A and NB-52B 

Aircraft,” AFFTC TD-70-2 (1970).
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A Lockheed F-104 flying chase for an X-15 lakebed landing. NASA.

motion-based training simulator. The success of the gliding entries and 

landings of the Space Shuttle orbiter confirm the value of this trainer.

Digital Computer Simulation

The computational mathematical models for the early simulators 

mentioned previously were performed on analog computers. Analog 

computers were capable of solving complex differential equations in 

real time. The digital computers available in the 1950s were mechanical 

units that were extremely slow and not capable of the rapid integration 

that was required for simulation. One difficulty with analog computers was the existence of electronic noise within the equipment, which 

caused the solutions to drift and become inaccurate after several minutes of operation. For short simulation exercises (such as a 10-minute 

X-15 flight) the results were quite acceptable. A second difficulty was 

storing data, such as aerodynamic functions. 

The X-20 Dyna-Soar program mentioned previously posed a challenge to the field of simulation. The shortest flight was to be a oncearound orbital flight with a flight time of over 90 minutes. A large volume 
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The family of 1960s–1970s reentry shapes developed by the Air Force Flight Dynamics 

Laboratory. USAF.

of aerodynamic data needed to be stored covering a very large range of 

Mach numbers and angles of attack. The analog inaccuracy problem was 

tackled by University of Michigan researchers, who revised the standard 

equations of motion so that the reference point for integration was a 300-

mile circular orbit, rather than the starting Earth coordinates at takeoff. 

These equations greatly improved the accuracy of analog simulations of 

orbiting vehicles. As the AFFTC and NASA began to prepare for testing of 

the X-20, an analog simulation was created at Edwards that was used to 

develop test techniques and to train pilots. Comparing the real-time simulation solutions with non-real-time digital solutions showed that the closure after 90 minutes was within about 20,000 feet—probably adequate for 

training, but they still dictated that the mission be broken into segments 

for accurate results. The solution was the creation of a hybrid computer 

simulation that solved the three rotational equations using analog computers but solved the three translational equations at a slower rate using 

digital computers. The hybrid computer equipment was purchased for 

installation at the AFFTC before the X-20 program was canceled in 1963. 

When the system was delivered, it was reprogrammed to represent the 

X-15A-2, a rebuilt variant of the second X-15 intended for possible flight 

to Mach 7, carrying a scramjet aerodynamic test article on a stub ventral 

fin.66 Although quite complex (it necessitated a myriad of analog-to-digital 

and digital-to-analog conversions), this hybrid system was subsequently 

66. Capt. Austin J. Lyons, “AFFTC Experiences with Hybrid Computation in a Real-Time Simulation 

of the X-15A-2,” AFFTC TR-66-44 (1967). 
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used in the AFFTC simulation lab to successfully simulate several other 

airplanes, including the C-5, F-15, and SR-71, as well as the M2-F2 and 

X-24A/B Lifting Bodies and Space Shuttle orbiter. 

The speed of digital computers increased rapidly in the 1970s, and 

soon all real-time simulation was being done with digital equipment. Outof-the-window visual displays also improved dramatically and began to 

be used in conjunction with the cockpit instruments to provide very realistic training for flight crews. One of the last features to be developed in 

the field of visual displays was the accurate representation of the terrain 

surface during the last few feet of descent before touchdown.

Simulation has now become a primary tool for designers, flight-test 

engineers, and pilots during the design, development, and flight-testing 

of new aircraft and spacecraft.

Dynamic Instabilities

There are dangerous situations that can occur because of either a coupling 

of the aerodynamics in different axes or a coupling of the aerodynamics 

with the inertial characteristics of an airplane. Several of these—Chuck 

Yeager’s close call with the X-1A in December 1953 and Milburn Apt’s 

fatal encounter in September 1956—have been mentioned previously.

Inertial Roll Coupling

Inertial roll coupling is the dynamic loss of control of an airplane occurring during a rapid roll maneuver. The phenomenon of inertial roll coupling is directly related to the evolution of aircraft design. At the time 

of the Wrights through much of the interwar years, wingspan greatly 

exceeded fuselage length. As aircraft flight speeds rose, the aspect ratio 

of wings decreased, and the fineness ratio of fuselages rose, so that by the 

end of the Second World War, wingspan and fuselage length were roughly 

equal. In the supersonic era that followed, wingspan reduced dramatically, and fuselage length grew appreciably (think, for example, of an aircraft such as the Lockheed F-104). Such aircraft were highly vulnerable 

to pitch/yaw/roll-coupling when a rapid rolling maneuver was initiated. 

The late NACA–NASA engineer and roll-coupling expert Dick Day 

described inertial roll coupling as “a resonant divergence in pitch or yaw 

when roll rate equals the lower of the pitch or yaw natural frequencies.”67

67. Richard E. Day, “Coupling Dynamics in Aircraft: A Historical Perspective,” SP-532 (1997), p. 1. 
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The existence of inertial roll coupling was first revealed by NACA Langley 

engineer William H. Phillips in 1948, 5 years before it became a dangerous phenomenon.68 Phillips not only described the reason for the potential 

loss of control but also defined the criteria for identifying the boundaries 

of loss of control for different aircraft. During the 1950s, several research 

airplanes and the Century series fighters encountered fairly severe inertial 

coupling problems exactly as predicted by Phillips. These airplanes differed from the earlier prop-driven airplanes by having thin, short wings 

and the mass of the jet engine and fuel concentrated along the fuselage 

longitudinal axis. This resulted in a higher moment of inertia in the pitch 

and yaw axis but a significantly lower inertia in the roll axis. The low roll 

inertia also allowed these airplanes to achieve higher roll rates than their 

predecessors had. The combination allowed the mass along the fuselage 

to be slung outward when the airplane was rolled rapidly, producing an 

unexpected increase in pitching and yawing motion. This divergence in 

pitch or yaw was related to the magnitude of the roll rate and the duration of the roll. If the roll were sustained long enough, the pitch or yaw 

angles would become quite large, and the airplane would tumble out of 

control. In most cases, the yaw axis had the lowest level of static stability, 

so the divergence was observed as a steady increase in sideslip.69

In 1954, after North American Aviation had encountered roll instability with its F-100 aircraft, the Air Force and NAA transferred an F-100A 

to NACA FRC to allow the NACA to explore the problem through flighttesting and identify a fix. The NACA X-3 research airplane was of a configuration much like the modern fighters and was also used by NACA 

FRC to explore the inertial coupling problem. These results essentially 

confirmed Phillips’s earlier predictions and determined that increasing 

the directional stability via larger vertical fin area would mitigate the 

68. William H. Phillips, “Effect of Steady Rolling on Longitudinal and Directional Stability, NACA 

TN-627 (1948).

69. Joseph Weil, Ordway B. Gates, Jr., Richard D. Banner, and Albert E. Kuhl, “Flight Experience of Inertia Coupling in Rolling Maneuvers,” RM H55WEIL (1955); HSFS, “Flight Experience 

With Two High-Speed Airplanes Having Violent Lateral-Longitudinal Coupling in Aileron Rolls,” RM 

H55A13 (1955); Hubert M. Drake and Wendell H. Stillwell, “Behavior of the Bell X-1A Research 

Airplane During Exploratory Flights at Mach Numbers Near 2.0 and at Extreme Altitudes,” RM 

H55G25 (1955); Hubert M. Drake, Thomas W. Finch, and James R. Peele, “Flight Measurements 

of Directional Stability to a Mach Number of 1.48 for an Airplane Tested with Three Different Vertical Tail Configurations,” RM H55G26 (1955); Walter C. Williams and William H. Phillips, “Some 

Recent Research on the Handling Qualities of Airplanes,” RM H55L29a (1956). 
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problem. The Century series fighters were all reconfigured to reduce 

their susceptibility to inertial coupling. The vertical tail size was increased 

for the F-100C and D airplanes.70 All F-104s were retrofitted with a ventral fin on the lower aft fuselage, which increased their directional stability by 10 to 15 percent. The F-104B, and later models, also had a larger 

vertical fin and rudder. The F-102 and F-105 received a larger vertical 

tails than their predecessors (the YF-102 and YF-105) did, and the Mach 

2+ F-106 had a larger vertical tail than the F-102 had. Control limiting 

and placards against continuous rolls (more than 720 degrees of bank) 

were instituted to ensure safe operation. The X-15 was also susceptible 

to inertial coupling, and its roll divergence tendencies could be demonstrated on the X-15 simulator. Since high roll rates were not necessary 

for the high-speed, high-altitude mission of the airplane, the pilots were 

instructed to avoid high roll rates, and, fortunately, no inertial coupling 

problems occurred during its flight-testing.

Flight Control Coupling

Flight control coupling is a slow loss of control of an airplane because 

of a unique combination of static stability and control effectiveness. Day 

described control coupling—the second mode of dynamic coupling—as 

“ a coupling of static yaw and roll stability and control moments which 

can produce untrimmability, control reversal, or pilot-induced oscillation (PIO).”71 So-called “adverse yaw” is a common phenomenon associated with control of an aircraft equipped with ailerons. The down-going 

aileron creates an increase in lift and drag for one wing, while the upgoing aileron creates a decrease in lift and drag for the opposite wing. 

The change in lift causes the airplane to roll toward the up-going aileron. The change in drag, however, results in the nose of the airplane 

swinging away from the direction of the roll (adverse yaw). If the airplane exhibits strong dihedral effect (roll produced by sideslip, a quality 

more pronounced in a swept wing design), the sideslip produced by the 

aileron deflections will tend to detract from the commanded roll. In the 

extreme case, with high dihedral effect and strong adverse yaw, the roll 

can actually reverse, and the airplane will roll in the opposite direction to 

that commanded by the pilot—as sometimes happened with the Boeing 

70. Robert G. Hoey and Capt. Milburn G. Apt, “F-100C Phase IV Stability and Control Test” 

AFFTC TR-56-25, Oct. 1956, pp 8, 144, 145. 

71. Day, “Coupling Dynamics in Aircraft,” p. 1. 
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B-47, though by aeroelastic twisting of a wing because of air loads. If the 

pilot responds by adding more aileron deflection, the roll reversal and 

sideslip will increase, and the airplane could go out of control.

As discussed previously, the most dramatic incident of control coupling occurred during the last flight of the X-2 rocket-powered research 

airplane in September 1956. The dihedral effect for the X-2 was quite 

strong because of the influence of wing sweep rather than the existence 

of actual wing dihedral. Dihedral effect because of wing sweep is nonexistent at zero-lift but increases proportionally as the angle of attack of 

the wing increases. After the rocket burned out, which occurred at the 

end of a ballistic, zero-lift trajectory, the pilot started a gradual turn by 

applying aileron. He also increased the angle of attack slightly to facilitate the turn, and the airplane entered a region of roll reversal. The sideslip increased until the airplane went out of control, tumbling violently. 

The data from this accident were fully recovered, and the maneuver was 

analyzed extensively by the NACA, resulting in a better understanding 

of the control-coupling phenomenon. The concept of a control parameter was subsequently created by the NACA and introduced to the industry. This was a simple equation that predicted the boundary conditions 

for aileron reversal based on four stability derivatives. When the yawing moment due to sideslip divided by the yawing moment due to aileron is equal to the rolling moment due to sideslip divided by the rolling 

moment due to aileron, the airplane remains in balance and aileron 

deflection will not cause the airplane to roll in either direction.72

Dutch Roll Coupling

Dutch roll coupling is another case of a dynamic loss of control of an 

airplane because of an unusual combination of lateral-directional static 

stability characteristics. Dutch roll coupling is a more subtle but nevertheless potentially violent motion, one that (again quoting Day) is a 

“dynamic lateral–directional stability of the stability axis. This coupling 

of body axis yaw and roll moments with sideslip can produce lateral–

directional instability or PIO.”73 A typical airplane design includes “static 

directional stability” produced by a vertical fin, and a small amount of 

“dihedral effect” (roll produced by sideslip). Dihedral effect is created 

by designing the wing with actual dihedral (wingtips higher than the 

72. In engineering shorthand, Cn/Cna=Cl/Cla.

73. Day, “Coupling Dynamics in Aircraft,” p. 1.
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wing root), wing sweep (wingtips aft of the wing root), or some combination of the two. Generally static directional stability and normal dihedral effect are both stabilizing and both contribute to a stable Dutch roll 

mode (first named for the lateral-directional motions of smooth-bottom 

Dutch coastal craft, which tend to roll and yaw in disturbed seas). When 

the interactive effects of other surfaces of an airplane are introduced, 

there can be potential regions of the flight envelope where these two 

contributions to Dutch roll stability are not stabilizing (i.e., regions of 

negative static directional stability or negative dihedral effect). In these 

regions, if the negative effect is smaller than the positive influence of the 

other, then the airplane will exhibit an oscillatory roll-yaw motion. (If 

both effects are negative, the airplane will show a static divergence in 

both the roll and yaw axes.) All aircraft that are statically stable exhibit 

some amount of Dutch roll motion. Most are well damped, and the Dutch 

roll only becomes apparent in turbulent conditions.

The Douglas DC-3 airliner (equivalent to the military C-47 airlifter) 

had a persistent Dutch roll that could be discerned by passengers watching the wingtips as they described a slow horizontal “figure eight” with 

respect to the horizon.

Even the dart-like X-15 manifested Dutch roll characteristics. The 

very large vertical tail configuration of the X-15 was established by the 

need to control the airplane near engine burnout if the rocket engine 

was misaligned, a “lesson learned” from tests of earlier rocket-powered 

aircraft such as the X-1, X-2, and D-558-2. This led to a large symmetrical vertical tail with large rudder surfaces both above and below the 

airplane centerline. (The rocket engine mechanics and engineers at 

Edwards later devised a method for accurately aligning the engine, so 

that the large rudder control surfaces were no longer needed.) The X-15 

simulator accurately predicted a strange Dutch roll characteristic in the 

Mach 3–4 region at angles of attack above 8 degrees with the roll and yaw 

dampers off. This Dutch roll mode was oscillatory and stable without 

pilot inputs but would rapidly diverge into an uncontrollable pilotinduced-oscillation when pilot control inputs were introduced.

During wind tunnel tests after the airplane was constructed, it was 

discovered that the lower segment of the vertical tail, which was operating in a high compression flow field at hypersonic speeds, was highly 

effective at reentry angles of attack. The resulting rolling motions produced by the lower fin and rudder were contributing a large negative 

dihedral effect. Fortunately, this destabilizing influence was not enough 
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to overpower the high directional stability produced by the very large 

vertical tail area, so the Dutch roll mode remained oscillatory and stable. 

The airplane motions associated with this stable oscillation were completely foreign to the test pilots, however. Whereas a normal Dutch roll 

is described as “like a marble rolling inside a barrel,” NASA test pilot Joe 

Walker described the X-15 Dutch roll as “like a marble rolling on the outside of the barrel” because the phase relationship between rolling and 

yawing were reversed. Normal pilot aileron inputs to maintain the wings 

level were out of phase and actually drove the oscillation to larger magnitudes rather quickly. The roll damper, operating at high gain, was fairly 

effective at damping the oscillation, thus minimizing the pilot’s need to 

actively control the motion when the roll damper was on.74

Because the X-15 roll damper was a single string system (fail-safe), a 

roll damper failure above about 200,000 feet altitude would have caused 

the entry to be uncontrollable by the pilot. The X-15 envelope expansion 

to altitudes above 200,000 feet was delayed until this problem could be 

resolved. The flight control team proposed installing a backup roll damper, 

while members of the aerodynamic team proposed removing the lower 

ventral rudder. Removing the rudder was expected to reduce the directional stability but also would cause the dihedral effect to be stable, thus 

the overall Dutch roll stability would be more like a normal airplane. The 

Air Force-NASA team pursued both options. Installation of the backup roll 

damper allowed the altitude envelope to be expanded to the design value 

of 250,000 feet. The removal of the lower rudder, however, solved the PIO 

problem completely, and all subsequent flights, after the initial ventral-off 

demonstration flights, were conducted without the lower rudder.75

The incident described above was unique to the X-15 configuration, but the analysis and resolution of the problem is instructive in that 

it offers a prudent cautioning to designers and engineers to avoid designs 

that exhibit negative dihedral effect.76

Configuration Influence upon Stall and Departure Behavior

Another maneuver that can lead to loss of control is a stall. An aircraft 

“stalls” when the wing’s angle of attack exceeds a critical angle beyond 

74. Stillwell, X-15 Research Results, pp. 51–52; Thompson, At the Edge of Space, pp. 200–202. 

75. Personal experience as a flight planning engineer during the X-15 initial envelope expansion tests.

76. Robert G. Hoey, “Correlation of X-15 Simulation Experience with Flight Test Results, AGARD 

Report 530 (1966). 
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which the wing can no longer generate the lift necessary to support the 

airplane. A typical stall consists of some pre-stall warning buffet as the 

flow over the wing begins to break down, followed by stall onset, usually accompanied by an uncommanded nose-down pitching rotation 

of the aircraft, as gravity takes over and the airplane naturally tries to 

regain lost airspeed. The loss of control for a normal stall is quite brief 

and can usually be overcome, or prevented, by proper control application at the time of pre-stall warning. There are design features of some 

aircraft that result in quite different stall characteristics. Stalls may be a 

straightforward wings-level gentle drop (typically leading to a swift and 

smooth recovery), or sharply abrupt, or an unsymmetrical wing drop 

leading to a spin entry. The latter can be quite hazardous.

High-performance T-tail aircraft are particularly vulnerable to 

abnormal stall effects. Lockheed’s sleek F-104 Starfighter incorporated 

a T-tail operating behind a short, stubby, and extremely thin wing. 

As the wing approached the critical stall angle, the wing tip vortexes 

impinged on the horizontal tail creating an abrupt nose-up pitching 

moment, commonly referred to as a “pitch-up.” The pitch-up placed 

the airplane in an uncontrollable flight environment: either a highly 

oscillatory spin or a deep stall (a stable condition where the airplane 

remains locked in a high angle of attack vertical descent). To prevent 

inadvertent pitch-ups, the aircraft was equipped with a “stick shaker,” 

and a “stick kicker.” The stick shaker created an artificial vibration of 

the stick, simulating stall buffet, as the airplane approached a stall. The 

stick kicker applied a sharp nose-down command to the horizontal tail 

when the airplane reached the critical condition for an impending pitchup. A similar situation developed for the McDonnell F-101 Voodoo (also 

a T-tail behind a short, stubby wing). Stick shakers and kickers were 

quite successful in allowing these airplanes to operate safely throughout their operational lifespan. Overall, however, the T-tail layout was 

largely discredited for high-performance fighter and attack aircraft, 

the most successful postwar fighters being those with low-placed horizontal tails. Such a configuration, typified by the F-100, F-101, F-105, 

F-5, F-14, F-15, F-16, F/A-18, F-22, F-35, and a host of foreign aircraft, 

is now a design standard for tailed transonic and supersonic military 

aircraft. It was a direct outgrowth of the extensive testing the NACA 

did in the late 1940s and early 1950s on such aircraft as the D-558-2, 

the North American F-86, and the Bell X-5, all of which, to greater or 

lesser extents, suffered from pitch-up.
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The advent of the swept wing induced its own challenges. In 1935, 

German aerodynamicist Adolf Busemann discovered that aircraft could 

operate at higher speeds, and closer to the speed of sound (Mach 1), by 

using swept wings. By the end of the Second World War, American NACA 

researcher Robert T. Jones of Langley Memorial Aeronautical Laboratory 

had independently discovered its benefits as well. The swept wing subsequently transformed postwar military and civil aircraft design, but 

it was not without its own quite serious problems. The airflow over a 

swept wing tends to move aft and outboard, toward the tip. This results 

in the wingtip stalling before the rest of the wing. Because the wingtip 

is aft of the wing root, the loss of lift at the tip causes an uncommanded 

nose-rise as the airplane approaches a stall. This nose-rise is similar to a 

pitch-up but not nearly as abrupt. It can be controlled by the pilot, and 

for most swept wing airplanes there are no control system features specifically to correct nose-rise problems. Understanding the manifestations 

of swept wing stall and swept wing pitch-up commanded a great deal of 

NACA and Air Force interest in the early years of the jet age, for reasons 

of both safety and combat effectiveness. Much of the NACA’s research 

program on its three swept wing Douglas D-558-2 Skyrockets involved 

examination of these problems. Research included analysis of a variety 

of technological “fixes,” such as sawtooth leading edge extensions, wing 

fences, and fixed and retracting slots. Afterward, various combinations 

of flaps, flow direction fences, wing twist, and other design features 

have been used to overcome the tip-stall characteristic in modern swept 

wing airplanes, which, of course, include most commercial airliners.77

Aerothermodynamics: Meeting the Heating Challenge

The prediction of structural heating on airplanes flying at hypersonic 

speeds preceded the actual capability to attain these speeds in controlled flight. There were dire predictions of airplanes burning up when 

they encountered the “thermal thicket,” similar to the dire predictions 

that preceded flight through the sound barrier. Aerodynamic heating is 

created by friction of an object moving at very high speed through the 

atmosphere. Temperatures on the order of 200 degrees Fahrenheit (°F) 

are generated at Mach 2 (the speed of an F-104 Starfighter of the mid1950s), 600 °F at Mach 3 (that of a 1960s SR-71 Blackbird), and 1,200 

77. See, for example, Document 25 in J.D. Hunley, ed., Toward Mach 2: The Douglas D-558 

Program, SP-4222 (Washington, DC: NASA, 1999), pp. 101–103.
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°F at Mach 6 (typical of the X-15). Reentry from orbital speeds (Mach 

26—the entry velocity of the Space Shuttle orbiter) will generate temperatures of around 2,400 °F. Airplanes or spacecraft that fly in, or 

reenter, the atmosphere above Mach 2 must be designed to withstand 

not only aerodynamic forces associated with high Mach number but 

also the high temperatures associated with aerodynamic heating. The 

advent of blunt body reentry theory radically transformed the mental 

image of the spacecraft, from a “pointy” rocket to one having a far more 

bluff and rounded body. Conceived by H. Julian Allen with the assistance of Alfred Eggers of the then-NACA Ames Aeronautical Laboratory 

(now NASA Ames Research Center), blunt-body design postulated using 

a blunt reentry shape to form a strong “detached” shock wave that could 

act to relieve up to 90 percent of the thermal load experienced by a body 

entering Earth’s atmosphere from space.78 Such a technical approach was 

first applied to missile warhead design and the first crewed spacecraft, 

both Soviet and American. But blunt bodies, for all their commendable 

thermodynamic characteristics, likewise have high drag and poor entry 

down-range and cross-range predictability. Tailored higher L/D lifting body 

and blended wing-body shapes (such as those pioneered by the Air Force 

Flight Dynamics Laboratory), while offering far better aerodynamic and 

cross-range performance and predictability, pose far greater cooling challenges. So, too, do concepts for hypersonic air-breathing vehicles. These 

diverse requirements have stimulated the design and development of several potential solutions for thermal protection of a vehicle. For purposes 

of discussion, these concepts are addressed as heat sink structures, ablation, hot structures, active cooling, and advanced ceramic protection.

Heat-Sink Structures

Prior to the X-15 flight-test program, there were several theories predicting the amount of friction heat that would transfer to the surface 

of a winged aircraft, with substantial differences in the theories. Wind 

tunnels, ballistic ranges, and high-temperature facilities such as arc-jets 

were unable to adequately duplicate the flight environment necessitating 

78. H. Julian Allen and Alfred J. Eggers, Jr., “A Study of the Motion and Aerodynamic Heating of 

Ballistic Missiles Entering the Earth’s Atmosphere at High Supersonic Speeds,” NACA Technical 

Report 1381 (1958) [this widely distributed report was preceded by a more restricted limited-issue 

classified report for Government and industry earlier]; see also Hartman, Adventures in Research, pp. 

215–218.
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full-scale flight test to determine which theory was correct. The X-15 

was that test aircraft. The design needed to be robust in order to survive 

the worst-case heating predictions if the theories proved to be correct.

A heat-sink structure was selected as the safest and simplest method 

for providing thermal protection. Inconel X, a nickel alloy normally used 

for jet engine exhaust pipes, was selected as the primary structural material. It maintained adequate structural strength to about 1,200 °F. The 

design proceeded by first defining the size of each structural member 

based on the air loads anticipated during entry, then increasing the size 

of each member to absorb the expected heat load that would occur during the short exposure time of an X-15 flight.

As with most of the first missile and aircraft explorations, early hypersonic flights in the X-15 showed that none of the prediction methods was 

completely accurate, although each method showed some validity in a certain Mach range. In general, the measured heat transfer was less than predicted. Thus, one of the most significant flight-test results from the X-15 

program was development of more accurate prediction methods based 

upon real-world data for the thermal protection of future hypersonic and 

entry vehicles.79 The majority of aerodynamic heating issues that required 

attention during the X-15 flight-test program were associated with localized heating: that is, unexpected hot spots that required modification. 

Some typical examples included loss of cockpit pressurization because of 

a burned canopy seal (resolved by installing a protective shield in front of 

the canopy gap), cockpit glass cracked because of deformation of the glass 

retainer ring (resolved by increasing the clearance around the glass), wing 

skin buckling behind the slot in the leading edge expansion joint (resolved 

by installing a thin cover over the expansion joint), thermal expansion of 

the fuselage triggering nose gear door deployment with resulting damage 

to internal instrumentation (resolved by increasing the slack in the deployment cable), and buckling of skin on side tunnel fairings because of large 

temperature difference between outer skin and liquid oxygen (LOX) tank 

(resolved by adding expansion joints along the side tunnels).

Most of these issues were discovered and resolved fairly easily since 

the flight envelope was expanded gradually on successive flights with 

small increases in Mach number on each flight. Had the airplane been 

exposed to the design entry environment on its very first flight, the 

79. Jenkins, X-15: Extending the Frontiers of Flight, passim.
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combined results of these local heating problems would probably have 

been catastrophic. 

Ablation Cooling

Another potential method for disbursing heat during high-speed flight 

was the application of an “ablation” material to the outer surface of the 

structure. An ablator is a material that is applied to the outside of a vehicle that burns or chars when exposed to high temperature, thus carrying away much of the associated heat and hot gases. Ablators are quite 

efficient for short duration, one-time entries such as an intercontinental ballistic missile (ICBM) nose cone. Ablators were also used on the 

early crewed orbiting capsules (Mercury, Gemini, and Apollo), which 

used ballistic or semiballistic entry trajectories with relatively short peak 

heating exposure times. They seemed to offer special promise for lifting 

bodies, with developers hoping to build classes of aluminum-structured 

spacecraft that could have a cheap, refurbishable ablative coating reapplied after each flight. Indeed, on April 19, 1967, the Air Force did fly 

and recover one such subscale experimental vehicle, the Mach 27 Martin 

SV-5D (X-23) Precision Recovery Including Maneuvering Entry (PRIME) 

lofted over the Pacific Test Range by a modified Atlas ballistic missile.80

But for all their merits, ablators are hardly a panacea. Subsonic and 

transonic testing of several rocket-powered aluminum lifting bodies at 

NASA’s Flight Research Center showed that this class of vehicle could 

be landed; however, later analysis indicated that the rough surface of 

an exposed ablator would probably have reduced the lift and increased 

the drag so that successful landings would have been questionable.81 

Flight-test experience with the X-15 confirmed such conclusions. 

When the decision was made to rebuild the second X-15 after a crash 

landing, it seemed a perfect opportunity to demonstrate the potential 

of ablative coatings as a means of furnishing refurbishable thermal protection to hypersonic aircraft and spacecraft. The X-15A-2 was designed 

to reach Mach 7, absorbing the additional heat load it would experience 

via MA-25S, a thin Martin-developed silica ablative coating. Coating the 

aircraft with the MA-25S proved surprisingly time-consuming, as did 

the refurbishment between flights.

80. Joel W. Powell and Ed Hengeveld, “ASSET and PRIME: Gliding Re-Entry Test Vehicles,” Journal 

of the British Interplanetary Society, vol. 36 (1983), pp. 369–376.

81. Personal inspection of the SV-5D (X-23) heat shield following vehicle recovery.
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During a flight to Mach 6.7 by Maj. William J. “Pete” Knight, unanticipated heating actions severely damaged the aircraft, melting a scramjet 

boilerplate test module off the airplane and burning holes in the external skin. Though Knight landed safely—a great tribute to his piloting 

skills—the X-15A-2 was in no condition to fly without major repairs. 

Although the ablator did provide the added protection needed for most 

of the airplane, the tedious process of applying it and the operational 

problems associated with repairing and protecting the soft coating were 

quite time-consuming and impracticable for an operational military or 

civilian system.82 The postentry ablated surface also increased the drag 

of the airplane by about the same percentage that was observed on the 

PRIME vehicle. Clearly the X-15A-2’s record flight emphasized, as NASA 

engineer John V. Becker subsequently wrote, “the need for maximum 

attention to aerothermodynamic detail in design and preflight testing.”83 

The “lifting body” concept evolved as a means of using ablative protection for entries of wingless, but landable, vehicles. As a result of the 

X-15 and lifting body testing by NASA, an ablative coating has not been 

seriously considered for any subsequent reusable lifting entry vehicle. 

Hot Structure Approaches

Another option for thermal protection during entry was the use of 

exotic, high-temperature materials for the external surface that could reradiate the heat back into space. This concept was proposed for the 

X-20 Dyna-Soar program, and the vehicle was well under construction 

at the time of cancellation.84 In parallel with the X-20 program, the Air 

Force Flight Dynamics Laboratory developed a small radiative-cooled 

hot structure vehicle (essentially the first 4 feet of the X-20 Dyna Soar’s 

nose), called the McDonnell Aerothermodynamic/elastic Structural 

Systems Environmental Tests (ASSET). The ASSET design used the 

same materials and thermal protection concepts as the X-20 and first 

flew in September 1963, 3 months before cancellation of the DynaSoar. The fourth ASSET vehicle successfully completed a Mach 18.4 

entry from 202,000 feet in 1965. Postflight examination indicated it 

82. Johnny G. Armstrong, “Flight Planning and Conduct of the X-15A-2 Envelope Expansion Program”, FTC TD-69-4, July 1969.

83. John V. Becker, “The X-15 Program in Retrospect,” Raumfahrtforschung (Mar.–Apr. 1969).

84. Robert Godwin, ed., Dyna-Soar Hypersonic Strategic Weapons System (Burlington, ON, Canada: Apogee Books, 2003) has an excellent compilation of contemporary Dyna-Soar documents.
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survived the entry well, although the operational problems and manufacturing methods for these exotic materials were expensive and 

time-consuming. Since that time, joint NASA-Air Force-Navy-industry 

developmental programs such as the X-30 National Aero-Space Plane 

(NASP) effort of the late 1980s to early 1990s have advanced materials and fabrication technologies that, in due course, may be applied to 

future hypersonic systems.85

Lightweight Ceramic Tiles

Ceramic tiles, of the kind used in a blast furnace or fireplace to insulate 

the surrounding structure from the extreme temperatures, were far too 

heavy to be considered for use on a flight vehicle. The concept of a lightweight ceramic tile for thermal protection was conceived by Lockheed 

and developed into operational use by NASA Ames Research Center, 

NASA Johnson Space Center, and Rockwell International for use on 

the Space Shuttle orbiter, first flown into orbit in April 1981. The resulting tiles and ceramic blankets provided exceptionally light and efficient 

thermal protection for the orbiter without altering the external shape. 

Although highly efficient for thermal protection, the tiles were—and are—

quite fragile and time-consuming to repair and maintain. The Shuttle 

program experienced considerable delays prior to its first flight because 

of bonding, breaking, and other installation issues. (Unlike the X-15 gradual envelope expansion program, the Shuttle orbiter was exposed to its 

full operational flight envelope on its very first orbital flight and entry, 

thus introducing a great deal of analysis and caution during flight preparation.) Subsequent Shuttle history confirmed the high-maintenance 

nature of the tiles, and their vulnerability to external damage such as 

ice or insulation shedding from the super-cold external propellant tank. 

Even with these limitations, however, they do constitute the most promising technology for future lifting entry vehicles.86 

85. M.H. Shirk, “ASSET: Aerothermoelastic Vehicles (AEV) Results and Conclusions,” AFFDL 65FD1197 (1965); “ASSET Final Briefing,” AFFDL 65FD-850 (1965).

86. Paul Cooper and Paul F. Holloway, “The Shuttle Tile Story,” Astronautics & Aeronautics, vol. 19, 

no. 1 (Jan. 1981), pp. 24–34; Robert G. Hoey, et al., “Evaluation of the Space Shuttle Orbiter First 

Orbital Flight: Final Report,” AFFTC TR-81-21 (1981); NASA LRC, “Shuttle Performance: Lessons 

Learned,” NASA Conference Publication 2283 (1983); and Robert G. Hoey, et al., “Flight Test 

Results from the Entry and Landing of the Space Shuttle Orbiter for the First Twelve Orbital Flights: 

Final Report,” AFFTC TR-85-11 (1985).
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Active Cooling Approaches

There are other proposed methods for protecting vehicles from high 

temperature while flying at high speed or during reentry. Several active 

cooling concepts have been proposed where liquid is circulated through a 

hot area, then through a radiator to dissipate the heat. These concepts are 

quite complex and the risk is very high: failure of an active cooling system 

could result in loss of a hypersonic vehicle within a few seconds. None has 

been demonstrated in flight. Although work is continuing on active cooling 

concepts, their application will probably not be realized for many years. 

As we look ahead to the future of aviation, it is easy to merely assess 

the current fleet of successful aircraft or spacecraft, and decide on what 

improvements we can provide, without considering the history and evolution that produced these vehicles. The danger is that some of the past problems will reappear unless the design and test communities are aware of their 

history. This paper has attempted to summarize some of the problems that 

have been encountered, and resolved, during the technology explosion in 

aviation that has occurred over the last 60 years. The manner in which the 

problems were discovered, the methods used to determine causes, and the 

final resolution or correction that was implemented have been presented. 

Hopefully these brief summaries of historical events will stimulate further 

research by our younger engineers and historians into the various subjects 

covered, and to that end, the following works are particularly relevant. 
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Recommended Additional Readings

Reports, Papers, Articles, and Presentations: 

Johnny G. Armstrong, “Flight Planning and Conduct of the X-15A-2 

Envelope Expansion Program,” AFFTC TD-69-4 (1969).

Theodore G. Ayers and James B. Hallissy, “Historical Background and 

Design Evolution of the Transonic Aircraft Technology Supercritical 

Wing,” NASA TM-81356 (1981).

John V. Becker, “The X-15 Project,” Astronautics and Aeronautics, vol. 

2, no. 2 (Feb. 1964).

John V. Becker, “The X-15 Program in Retrospect,” Raumfahrtforschung 

(Mar.–Apr. 1969).

C.C. Clark and C.H. Woodling, “Centrifuge Simulation of the X-15 

Research Aircraft,” Naval Air Development Center NADC MA-5916 

(1959).

Dwain A. Deets and Kenneth J. Szalai, “Design and Flight Experience 

with a Digital Fly-By-Wire Control System in an F-8 Airplane,” 

NATO Advisory Group for Aeronautical Research and Development, 

Conference Paper AGARD-CP-137 (1974).

Robert G. Hoey, “Horizontal Landing Techniques for Hypersonic 

Vehicles,” NATO Advisory Group for Aeronautical Research and 

Development, AGARD Report No. 428 (1963).

Robert G. Hoey and Richard E. Day, “Mission Planning and Operational 

Procedures for the X-15 Airplane,” NASA TN D-1158 (1962).

Robert G. Hoey, et al., “Flight Test Results from the Entry and Landing 

of the Space Shuttle Orbiter for the First Twelve Orbital Flights,” 

AFFTC TR-85-11 (1985).

E.C. Holloman, “Piloting Performance During the Boost of the X-15 

Airplane to High-Altitude,” NASA TN-D-2289 (1964).
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C. Hwang, E. Jonson, G. Mills, T. Noll, and M. Farmer, “Wind Tunnel 

Test of a Fighter Aircraft Wing/Store Flutter Suppression System: 

An International Effort,” NATO Advisory Group for Aeronautical 

Research and Development, AGARD R-689 (1980).

Kenneth W. Iliff and L.W. Taylor, Jr., “Determination of Stability 

Derivatives from Flight Data Using a Newton-Raphson Minimization 

Technique,” NASA TN-D-6579 (1972).

Kenneth W. Iliff, “Aircraft Parameter Estimation,” AIAA Paper 1987-

0623 (1987).

E.E. Kordes and R.B. Noll, “Flight Flutter Results for Flat Rectangular 

Panels,” NASA TN-D-1058 (1962).

E.E. Kordes, R.D. Reed, and A.L. Dawdy, “Structural Heating Experiences 

on the X-15 Airplane,” NASA TM-X-711 (1962).

Andrew M. Lizotte and Michael J. Allen, “Twist Model Development and 

Results From the Active Aeroelastic Wing F/A-18 Aircraft,” NASA 

TM-2005-212861 (2005).

NASA, “Flight Test Results Pertaining to the Space Shuttlecraft,” NASA 

TM-X-2101 (1970).

R.B. Noll, C.R. Jarvis, C. Pembo, W.P. Lock, and B.J. Scott, “Aerodynamic 

and Control-System Contributions to the X-15 Airplane LandingGear Loads,” NASA TN-D-2090 (1963).

Weneth D. Painter and George J. Sitterle, “Ground and Flight Test 

Methods for Determining Limit Cycle and Structural Resonance 

Characteristics of Aircraft Stability Augmentation Systems,” NASA 

TN-D-6867 (1972).

NASA, “Shuttle Performance: Lessons Learned,” NASA CP-2283, pt. 

1–2, (1983).
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Paul Pellicano, Joseph Krumenacker, and David Van Hoy, “X-29 High 

Angle-of-Attack Flight Test Procedures, Results, and Lessons 

Learned,” Society of Flight Test Engineers 21st Annual Symposium 

(1990).

F.S. Petersen, H.A. Rediess, and J. Weil, “Lateral-Directional Control 

Characteristics of the X-15 Airplane,” NASA TM-X-726 (1962).

Paul W. Phillips and Stephen B. Smith, “AFTI/F-111 Mission Adaptive 

Wing (MAW) Automatic Flight Control System Modes Lift and Drag 

Characteristics,” AFFTC TR-89-03 (1989).

William H. Phillips, “Effect of Steady Rolling on Longitudinal and 

Directional Stability,” NACA TN-627 (1948).

B.L. Schofield, D.F. Richardson, and P.C. Hoag, “Terminal Area Energy 

Management, Approach, and Landing Investigation for Maneuvering 

Reentry Vehicles using F-111A and NB-52B Aircraft,” AFFTC TD-70-2 

(1970).

L.W. Taylor, Jr., “Analysis of a Pilot-Airplane Lateral Instability 

Experienced with the X-15 Airplane,” NASA TN-D-1059 (1961).

L.W. Taylor, Jr., and E.J. Adkins, “Adaptive Flight Control Systems—Pro 

and Con,” NASA TM-X-56008 (1964).

L.W. Taylor, Jr., and G.B. Merrick, “X-15 Airplane Stability Augmentation 

System,” NASA TN-D-1157 (1962).

L.W. Taylor, Jr., and J.W. Smith, “An Analysis of the Limit-Cycle 

and Structural-Resonance Characteristics of the X-15 Stability 

Augmentation System,” NASA TN-D-4287 (1967).

Robert A. Tremant, “Operational Experience and Characteristics 

of the X-15 Flight Control System,” NASA TN-D-1402 (1962). 
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Books and Monographs:

Joseph R. Chambers, Partners in Freedom: Contributions of the Langley 

Research Center to U.S. Military Aircraft of the 1990s, SP-2000-4519 

(Washington, DC: NASA, 2000).

Richard E. Day, Coupling Dynamics in Aircraft: A Historical Perspective,

SP-532 (1997).

Dennis R. Jenkins, X-15: Extending the Frontiers of Flight, SP-2007-562 

(Washington, DC: NASA, 2007).

Ben R. Rich and Leo Janos, Skunk Works: A Personal Memoir of My Years 

of Lockheed (Boston: Little, Brown, and Co., 1994).

Milton O. Thompson, At the Edge of Space: The X-15 Flight Program 

(Washington, DC: Smithsonian Institution Press, 1992).

Milton O. Thompson with J.D. Hunley, Flight Research: Problems 

Encountered and What They Should Teach Us, SP-2000-4522 

(Washington, DC: NASA, 2000).
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426

Langley CFD Shuttle study, showing the hypersonic flow pattern as it executes atmospheric entry 

from orbit. NASA.
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NASA and the Evolution 

of Computational Fluid 

Dynamics

John D. Anderson, Jr. 

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The expanding capabilities of the computer readily led to its increasing 

application to the aerospace sciences. NACA–NASA researchers were 

quick to realize how the computer could supplement traditional test methodologies, such as the wind tunnel and structural test rig. Out of this came 

a series of studies leading to the evolution of computer codes used to 

undertake computational fluid dynamics and structural predictive studies. 

Those codes, refined over the last quarter century and available to the 

public, are embodied in many current aircraft and spacecraft systems.

T

HE VISITOR TO THE SMITHSONIAN INSTITUTION’S National Air and 

Space Museum (NASM) in Washington, DC, who takes the east 

escalator to the second floor, turns left into the Beyond the Limits 

exhibit gallery, and then turns left again into the gallery’s main bay is 

suddenly confronted by three long equations with a bunch of squiggly 

symbols neatly painted on the wall. These are the Navier-Stokes equations, and the NASM (to this author’s knowledge) is the world’s only 

museum displaying them so prominently. These are not some introductory equations drawn for a first course in algebra, with simple symbols 

like a + b = c. Rather, these are “partial derivatives” strung together from 

the depths of university-level differential calculus. What are the NavierStokes equations, why are they in a gallery devoted to the history of the 

computer as applied to flight vehicles, and what do they have to do with 

the National Aeronautics and Space Administration (which, by the way, 

dominates the artifacts and technical content exhibited in this gallery)?

The answers to all these questions have to do with computational 

fluid dynamics (CFD) and the pivotal role played by the National 

Aeronautics and Space Administration (NASA) in the development of 

CFD over the past 50 years. The role played by CFD in the study and 

understanding of fluid dynamics in general and in aerospace engineering 
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in particular has grown from a fledgling research activity in the 1960s 

to a powerful “third” dimension in the profession, an equal partner with 

pure experiment and pure theory. Today it is used to help design airplanes, study the aerodynamics of automobiles, enhance wind tunnel 

testing, develop global weather models, and predict the tracts of hurricanes, to name just a few. New jet engines are developed with an extensive use of CFD to model flows and combustion processes, and even the 

flow field in the reciprocating engine of the average family automobile is 

laid bare for engineers to examine and study using the techniques of CFD.

The history of the development of computational fluid dynamics 

is an exciting and provocative story. In the whole spectrum of the history of technology, CFD is still very young, but its importance today 

and in the future is of the first magnitude. This essay offers a capsule 

history of the development of theoretical fluid dynamics, tracing how 

the Navier-Stokes equations came about, discussing just what they are 

and what they mean, and examining their importance and what they 

have to do with the evolution of computational fluid dynamics. It then 

discusses what CFD means to NASA—and what NASA means to CFD. 

Of course, many other players have been active in CFD, in universities, 

other Government laboratories, and in industry, and some of their work 

will be noted here. But NASA has been the major engine that powered 

the rise of CFD for the solution of what were otherwise unsolvable problems in the fields of fluid dynamics and aerodynamics.

The Evolution of Fluid Dynamics from da Vinci to Navier-Stokes

Fluid flow has fascinated humans since antiquity. The Phoenicians 

and Greeks built ships that glided over the water, creating bow waves 

and leaving turbulent wakes behind. Leonardo da Vinci made detailed 

sketches of the complex flow fields over objects in a flowing stream, showing even the smallest vortexes created in the flow. He observed that the 

force exerted by the water flow over the bodies was proportional to the 

cross-sectional area of the bodies. But nobody at that time had a clue 

about the physical laws that governed such flows. This prompted some 

substantive experimental fluid dynamics in the 17th and 18th centuries. 

In the early 1600s, Galileo observed from the falling of bodies through 

the air that the resistance force (drag) on the body was proportional to 

the air density. In 1673, the French scientist Edme Mariotte published 

the first experiments that proved the important fact that the aerodynamic 

force on an object in a flow varied as the square of the flow velocity, not 
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directly with the velocity itself as believed by da Vinci and Galileo before 

him.1 Seventeen years later, Dutch scientist Christiaan Huygens published the same result from his experiments. Clearly, by this time, fluid 

dynamics was of intense interest, yet the only way to learn about it was 

by experiment, that is, empiricism.2

This situation began to change with the onset of the scientific revolution in the 17th century, spearheaded by the theoretical work of 

British polymath Isaac Newton. Newton was interested in the flow of 

fluids, devoting the whole Book II of his Principia to the subject of fluid 

dynamics. He conjured up a theoretical picture of fluid flow as a stream 

of particles in straight-line rectilinear motion that, upon impact with 

an object, instantly changed their motion to follow the surface of the 

object. This picture of fluid flow proved totally wrong, as Newton himself suspected, and it led to Newton’s “sine-squared law” for the force on 

a object immersed in a flow, which famously misled many early aeronautical pioneers. But if quantitatively incorrect, it was nevertheless the 

first to theoretically attempt an explanation of why the aerodynamic 

force varied directly with the square of the flow velocity.3

Newton, through his second law contributed indirectly to the breakthroughs in theoretical fluid dynamics that occurred in the 18th century. 

Newton’s second law states that the force exerted on a moving object is 

directly proportional to the time rate of change of momentum of that 

object. (It is more commonly known as “force equals mass time acceleration,” but this is not found in the Principia). Applying Newton’s second law to an infinitesimally small fluid element moving as part of a 

1. John D. Anderson, Jr., A History of Aerodynamics, (Cambridge: Cambridge University Press, 

1997), pp. 23, 25, 31, 35; Edme Mariotte, Traite de la percussion ou choc des corps (Paris: 

Academy of Sciences, 1673). 

2. For further references in the history of fluid dynamics, see R. Giacomelli and E. Pistolesi, “Historical Sketch,” in William F. Durand, ed., Aerodynamic Theory, vol. 1 (Berlin: Julius Springer Verlag, 

1934); Theodore von Kármán, Aerodynamics (Cornell: Cornell University Press, 1954); G.A. 

Tokaty, A History and Philosophy of Fluid Mechanics (Henley, UK: G.T. Foulis and Co., Ltd., 1971); 

Olivier Darrigol, Worlds of Flow: A History of Hydrodynamics from the Bernoullis to Prandtl (Oxford: 

Oxford University Press, 2005); and Michael Eckert, The Dawn of Fluid Dynamics: A Discipline 

Between Science and Technology (Weinheim, Germany: Wiley-VCH Verlag GmbH & Co. KGaA, 

2006). All trace the evolution of fluid mechanics and aerodynamics in great detail, with excellent 

references to primary sources and other works.

3. Isaac Newton, Mathematical Principles of Natural Philosophy (Chicago: Encyclopedia Britannica, Inc., 1952 ed. of a 1687 work), pp. 159–267. 
NASA’s Contributions to Aeronautics

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fluid flow that is actually a continuum material, Leonhard Euler constructed an equation for the motion of the fluid as dictated by Newton’s 

second law. Euler, arguably the greatest scientist and mathematician of 

the 18th century, modeled a fluid as a continuous collection of infinitesimally small fluid elements moving with the flow, where each fluid element 

can continually change its size and shape as it moves with the flow, but, 

at the same time, all the fluid elements taken as a whole constitute an 

overall picture of the flow as a continuum. That was somewhat in contrast to the individual and distinct particles in Newton’s impact theory 

model mentioned previously. To his infinitesimally small fluid element, 

Euler applied Newton’s second law in a form that used differential calculus, leading to a differential equation relating the variation of velocity and pressure throughout the flow. This equation, simply labeled the 

“momentum equation,” came to be known simply as Euler’s equation. 

In the 18th century, it constituted a bombshell in launching the field of 

theoretical fluid dynamics and was to become a pivotal equation in CFD 

in the 20th century, a testament to Euler’s insight and its application. 

There is a second fundamental principle that underlies all of fluid 

dynamics, namely that mass is conserved. Euler applied this principle 

also to his model of an infinitesimally small moving fluid element, constructing another differential equation labeled the “continuity equation.” These two equations, the continuity equation and the momentum 

equation, were published in 1753, considered one of his finest works. 

Moreover, these two equations, 200 years later, were to become the physical foundations of the early work in computational fluid dynamics.4

After Euler’s publication, for the next century all serious efforts to 

theoretically calculate the details of a fluid flow centered on efforts to 

solve these Euler equations. There were two problems, however. The 

first was mathematical: Euler’s equations are nonlinear partial differential equations. In general, nonlinear partial differential equations are 

not easy to solve. (Indeed, to this day there exists no general analytical 

solution to the Euler equations.) When faced with the need to solve a 

practical problem, such as the airflow over an airplane wing, in most 

cases an exact solution of the Euler equations is unachievable. Only by 

simplifying the fluid dynamic problem and allowing certain terms in the 

4. As explicated by Leonhard Euler in his Principles of the Motion of Fluids (1752), General 

Principles of the State of Equilibrium of Fluids (1753), and General Principles of the Motion of Fluids

(1755). 
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equations to be either dropped or modified in such a fashion to make 

the equations linear rather than nonlinear can these equations be solved 

in a useful manner. But a penalty usually must be paid for this simplification because in the process at least some of the physical or geometrical accuracy of the flow is lost. 

The second problem is physical: when applying Newton’s second law 

to his moving fluid element, Euler did not account for the effects of friction in the flow, that is, the force due to the frictional shear stresses rubbing on the surfaces of the fluid element as it moves in the flow. Some 

fluid dynamic problems are reasonably characterized by ignoring the 

effects of friction, but the 18th and 19th century theoretical fluid dynamicists were not sure, and they always worried about what role friction 

plays in a flow. However, a myriad of other problems are dominated 

by the effect of friction in the flow, and such problems could not even 

be addressed by applying the Euler equations. This physical problem 

was exacerbated by controversy as to what happens to the flow moving 

along a solid surface. We know today that the effect of friction between 

a fluid flow and a solid surface (such as the surface of an airplane wing) 

is to cause the flow velocity right at the surface to be zero (relative to 

the surface). This is called the no-slip condition in modern terminology, 

and in aerodynamic theory, it represents a “boundary condition” that 

must be accounted for in conjunction with the solution of the governing flow equations. The no-slip condition is fully understood in modern 

fluid dynamics, but it was by no means clear to 19th century scientists. 

The debate over whether there was a finite relative velocity between a 

solid surface and the flow immediately adjacent to the surface continued into the 2nd decade of the 20th century.5

 In short, the world of theoretical fluid dynamics in the 18th and 19th centuries was hopelessly 

cast adrift from many desired practical applications.

The second problem, that of properly accounting for the effects of 

friction in the flow, was dealt with by two mathematicians in the middle 

19th century, France’s Claude-Louis-Marie-Henri Navier, and Britain’s 

Sir George Gabriel Stokes. Navier, an instructor at the famed École nationale des ponts et chaussées, changed the pedagogical style of teaching 

civil engineering from one based mainly on cut-and-try empiricism to 

a program emphasizing physics and mathematical analysis. In 1822, he 

5. Anderson, History of Aerodynamics, p. 89.
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gave a paper to the Academy of Sciences that contained the first accurate representation of the effects of friction in the general partial differential momentum equation for fluid flow.6

 Although Navier’s equations 

were in the correct form, his theoretical reasoning was greatly flawed, 

and it was almost a fluke that he arrived at the correct terms. Moreover, 

he did not fully understand the physical significance of what he had 

derived. Later, quite independently from Navier, Stokes, a professor at 

Cambridge who occupied the Lucasian Chair at Cambridge University 

(the same chair Newton had occupied a century and a half earlier) took 

up the derivation of the momentum equation including the effects of 

friction. He began with the concept of internal shear stress caused by 

friction in the fluid and derived the governing momentum equation much 

like it would be derived today in a fluid dynamics class, publishing it 

in 1845.7 Working independently, then, Navier and Stokes derived the 

basic equations that describe fluid flows and contain terms to account 

for friction. They remain today the fundamental equations that fluid 

dynamicists employ for analyzing frictional flows.

Finally, in addition to the continuity and momentum equations, 

a third fundamental physical principle is required for any flow that 

involves high speeds and in which the density of the flow changes from 

one point to another. This is the principle of conservation of energy, 

which holds that energy cannot be created or destroyed; it can only 

change its form. The origin of this principle in the form of the first law of 

thermo-dynamics is found in the history of the development of thermodynamics in the late 19th century. When applied to a moving fluid 

element in Euler’s model, and including frictional dissipation and 

heat transfer by thermal conduction, this principle leads to the energy 

equation for fluid flow.

So there it is, the origin of the three Navier-Stokes equations exhibited so prominently at the National Air and Space Museum. They are 

horribly nonlinear partial differential equations. They are also fully coupled together because the variables of pressure, density, and velocity that 

appear in these equations are all dependent on each other. Obtaining a 

6. C.L.M.H. Navier, “Mémoire sur les du mouvement des fluids,” Mémoires de l’Academie Royale 

des Sciences, No. 6 (1823), pp. 389–416. 

7. G.G. Stokes, “On the Theories of the Internal Friction of Fluids in Motion, and of the Equilibrium 

and Motion of Elastic Solids,” Transactions of the Cambridge Philosophical Society, vol. 8, no. 22 

(1845), pp. 287–342.
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general analytical solution of the Navier-Stokes equations is much more 

daunting than the problem of obtaining a general analytical solution of 

the Euler equations, for they are far more complex. There is today no 

general analytical solution of the Navier-Stokes equations (as is likewise 

true in the case of the Euler equations). Yet almost all of modern computational fluid dynamics is based on the Navier-Stokes equations, and 

all of the modern solutions of the Navier-Stokes equations are based on 

computational fluid dynamics.

Computational Fluid Dynamics: What It Is, What It Does

What constitutes computational fluid dynamics? The basic equations 

of fluid dynamics, the Navier-Stokes equations, are expressions of three 

fundamental principles: (1) mass is conserved (the continuity equation), 

(2) Newton’s second law (the momentum equation), and (3) the energy 

equation (the first law of thermodynamics). Moreover, these equations 

in their most general form are either partial differential equations (as 

we have discussed) or integral equations (an alternate form we have not 

discussed involving integrals from calculus).

The partial differential equations are those exhibited at the NASM. 

Computational fluid dynamics is the art and science of replacing the 

partial derivatives (or integrals, as the case may be) in these equations 

with discrete algebraic forms, which in turn are solved to obtain numbers for the flow-field values (pressure, density, velocity, etc.) at discrete 

points in time and or space.8 At these selected points in the flow, called 

grid points, each of the derivatives in each of the equations are simply 

replaced with numbers that are advanced in time or space to obtain a 

solution for the flow. In this fashion, the partial differential equations 

are replaced by a large number of algebraic equations, which can then 

be solved simultaneously for the flow variables at all the grid points.

The end product of the CFD process is thus a collection of numbers, 

in contrast to a closed-form analytical solution (equations). However, 

in the long run, the objective of most engineering analyses, closed-form 

or otherwise, is a quantitative description of the problem: that is, numbers. Along these lines, in 1856, the famous British scientist James Clerk 

Maxwell wrote: “All the mathematical sciences are founded on relations 

between physical laws and laws of numbers, so that the aim of exact 

8. This discussion is elaborated upon in the author’s Computational Fluid Dynamics: The Basics With 

Applications (New York: McGraw-Hill, 1995).
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science is to reduce the problems of nature to the determination of quantities 

by operations with numbers.”9 Well over a century later, it is worth noting 

how well Maxwell captured the essence of CFD: operations with numbers.

Note that computational fluid dynamics results in solutions for the 

flow only at the distinct points in the flow called grid points, which were 

identified earlier. In a CFD solution, grid points are either initially distributed throughout the flow and/or generated during the course of the 

solution (called an “adaptive grid”). This is in theoretical contrast with 

a closed-form analytical solution for the flow, where the solution is in 

the form of equations that allow the calculation of the flow variables at 

any point of one’s choosing, that is, an analytical solution is like a continuous answer spread over the whole flow field. Closed-form analytical solutions may be likened to a traditionalist Dutch master’s painting 

consisting of continuous brush strokes, while a CFD solution is akin to a 

French pointillist consisting of multicolored dots made with a brush tip.

Generating a grid is an essential part of the art of CFD. The spacing 

between grid points and the geometric ways in which they are arrayed is 

critical to obtaining an accurate numerical CFD solution. Poor grids almost 

always ensure poor CFD solutions. Though good grids do not guarantee 

good CFD solutions, they are essential for useful solutions. Grid generation is a discipline all by itself, a subspecialty of CFD. And grid generation 

can become very labor-intensive—for some flows over complex threedimensional configurations, it may take months to generate a proper grid.

To summarize, the Navier-Stokes equations, the governing equations 

of fluid dynamics, have been in existence for more than 160 years, their 

creation a triumph of derivative insight. But few knew how to analytically solve them except for a few simple cases. Because of their complexity, they thus could not serve as a practical widely employed tool in the 

engineer’s arsenal. It took the invention of the computer to make that possible. And because it did so, it likewise permitted the advent of computational fluid dynamics. So how did the idea of numerical solutions to the 

Navier-Stokes equations evolve?

The Concept of Finite Differences Enters the Mathematical Scene

The earliest concrete idea of how to simulate a partial derivative with an 

algebraic difference quotient was the brainchild of L.F. Richardson in 

9. Ibid., p. 3.
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1910.10 He was the first to introduce the numerical solution of partial differential equations by replacing each derivative in the equations with an 

algebraic expression involving the values of the unknown dependent variables in the immediate neighborhood of a point and then solving simultaneously the resulting massive system of algebraic equations at all grid 

points. Richardson named this approach a “finite-difference solution,” a 

name that has come down without change since 1910. Richardson did 

not attempt to solve the Navier-Stokes equations, however. He chose a 

problem reasonably described by a simpler partial differential equation, 

Laplace’s equation, which in mathematical speak is a linear partial differential equation and which the mathematicians classify as an elliptic partial differential equation.11 He set up a numerical approach that 

is still used today for the solution of elliptic partial differential equations called a relaxation method, wherein a sweep is taken throughout 

the whole grid and new values of the dependent variables are calculated 

from the old values at neighboring grid points, and then the sweep is 

repeated over and over until the new values at each grid point converges 

to the old value from the previous sweep, i.e., the numbers “relax” eventually to the correct solution.

In 1928, Richard Courant, K.O. Friedrichs, and Hans Lewy published “On the Partial Difference Equations of Mathematical Physics,” 

a paper many consider as marking the real beginning of modern finite 

difference solutions; “Problems involving the classical linear partial differential equations of mathematical physics can be reduced to algebraic 

ones of a very much simpler structure,” they wrote, “by replacing the 

differentials by difference quotients on some (say rectilinear) mesh.”12

Courant, Friedrichs, and Lewy introduced the idea of “marching solutions,” whereby a spatial marching solution starts at one end of the flow 

and literally marches the finite-difference solution step by step from one 

10. L.F. Richardson, “The Approximate Arithmetical Solution by Finite Differences of Physical 

Problems Involving Differential Equations, With Application to the Stresses in a Masonry Dam,” 

Philosophical Transactions of the Royal Society of London, ser. A, vol. 210 (1910), pp. 307–357. 

11. Many partial differential equations fall within one of the following categories: elliptic equations, 

parabolic equations, and hyperbolic equations, differentiated by their mathematical (and graphical) behavior. The setting-up of a given numerical solution of these equations depends critically on 

whether the equation in question is elliptic, parabolic, or hyperbolic.

12. Richard Courant, K.O. Friedrichs, and Hans Lewy, “Über die Partiellen Differenzengleichungen 

der Mathematischen Physik,” Mathematische Annalen, vol. 100, (1928), pp. 32–74. All three 

emigrated to the U.S.
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end to the other end of the flow. A time marching solution starts with 

the all the flow variables at each grid point at some instant in time and 

marches the finite-difference solution at all the grid points in steps of 

time to some later value of time. These marching solutions can only be 

carried out for parabolic or hyperbolic partial differential equations, 

not for elliptic equations.

Courant, Friedrichs, and Lewy highlighted another important aspect 

of numerical solutions of partial differential equations. Anyone attempting numerical solutions of this nature quickly finds out that the numbers 

being calculated begin to look funny, make no sense, oscillate wildly, 

and finally result in some impossible operation such as dividing by zero 

or taking the square root of a negative number. When this happens, the 

solution has blown up, i.e., it becomes no solution at all. This is not a 

ramification of the physics, but rather, a peculiarity of the numerical 

processes. Courant, Friedrichs, and Lewy studied the stability aspects 

of numerical solutions and discovered some essential criteria to maintain stability in the numerical calculations. Today, this stability criterion 

is referred to as the “CFL criterion” in honor of the three who identified 

it. Without it, many attempted CFD solutions would end in frustration.

So by 1928, the academic foundations of finite difference solutions 

of partial differential equations were in place. The Navier-Stokes equations finally stood on the edge of being solved, albeit numerically. But 

who had the time to carry out the literally millions of calculations that 

are required to step through the solution? For all practical purposes, it 

was an impossible task, one beyond human endurance. Then came the 

electronic revolution and, with it, the digital computer.

The Critical Tool: Emergent High-Speed Electronic Digital Computing

During the Second World War, J. Presper Eckert and John Mauchly at 

the University of Pennsylvania’s Moore School of Electrical Engineering 

designed and built the ENIAC, an electronic calculator that inaugurated 

the era of digital computing in the United States. By 1951, they had 

turned this expensive and fragile instrument into a product that was manufactured and sold, a computer they called the UNIVAC, which stands 

for Universal Automatic Computer. The National Advisory Committee 

for Aeronautics (NACA) was quick to realize the potential of a high-speed 

computer for the calculation of fluid dynamic problems. After all, the 

NACA was in the business of aerodynamics and after 40 years of trying 

to solve the equations of motion by simplified analysis, it recognized 
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the breakthrough supplied by the computer to solve these equations 

numerically on a potentially practical basis. In 1954, Remington Rand 

delivered an ERA 1103 digital computer intended for scientific and 

engineering calculations to the NACA Ames Aeronautical Laboratory at 

Sunnyvale, CA. This was a state-of-the-art computer that was the first 

to employ a magnetic core in place of vacuum tubes for memory. The 

ERA 1103 used binary arithmetic, a 36-bit word length, and operated 

on all the bits of a word at a time. One year later, Ames acquired its first 

stored-program electronic computer, an IBM 650. In 1958, the 650 was 

replaced by an IBM 704, which in turn was replaced with an IBM 7090 

mainframe in 1961.13

The IBM 7090 had enough storage and enough speed to allow 

the first generation of practical CFD solutions to be carried out. By 

1963, four additional index registers were added to the 7090, making 

it the IBM 7094. This computer became the workhorse for the CFD of 

the 1960s and early 1970s, not just at Ames, but throughout the aerodynamics community; the author cut his teeth solving dissertation on 

an IBM 7094 at the Ohio State University in 1966. The calculation speed 

of a digital computer is measured in its number of floating point operations per second (FLOPS). The IBM 7094 could do 100,000 FLOPS, 

making it about the fastest computer available in the 1960s. With this 

number of FLOPS, it was possible to carry out for the first time detailed 

flow-field calculations around a body moving at hypersonic speeds, one 

of the major activities within the newly formed NASA that drove both 

computer and algorithm development for CFD. The IBM 7094 was a 

“mainframe” computer, a large electronic machine that usually filled 

a room with equipment. The users would write their programs (usually in the FORTRAN language) as a series of logically constructed line 

statements that would be punched on cards, and the decks of punched 

cards (sometimes occupying many boxes for just one program) would 

be fed into a reader that would read the punches and tell the computer 

what calculations to make. The output from the calculations would be 

printed on large sheets and returned to the user. One program at a time 

was fed into the computer, the so-called “batch” operation. The user 

would submit his or her batch to the computer desk and then return 

hours or days later to pick up the printed output. As cumbersome as it 

13. Paul E. Ceruzzi, Beyond the Limits: Flight Enters the Computer Age (Cambridge: MIT Press, 

1989), p. 15.
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may appear today, the batch operation worked. The field of CFD was 

launched with such batch operations on mainframe computers like the 

IBM 7094. And NASA Ames was a spearhead of such activities. Indeed, 

because of the synergism between CFD and the computers on which it 

worked, the demands on the central IBM installation at Ames grew at a 

compounded rate of over 100 percent per year in the 1960s.

With these computers, it became practical to set up CFD solutions 

of the Euler equations for two-dimensional flows. These solutions could 

be carried out with a relatively small number of grid points in the flow, 

typically 10,000 to 100,000 points, and still have computer run times 

on the order of hours. Users of CFD in the 1960s were happy to have 

this capability, and the three primary NASA Research Centers—Langley, 

Ames, and Lewis (now Glenn)—made major strides in the numerical 

analysis of many types of flows, especially in the transonic and hypersonic regimes. The practical calculation of inviscid (that is, frictionless), 

three-dimensional flows and especially any type of high Reynolds number flows was beyond the computer capabilities at that time.

This situation changed markedly when the supercomputer came 

on the scene in the 1970s. NASA Ames acquired the Illiac IV advanced 

parallel-processing machine. Designed at the University of Illinois, this 

was an early and controversial supercomputer, one bridging both older 

and newer computer architectures and processor approaches. Ames 

quickly followed with the installation of an IBM 360 time-sharing computer. These machines provided the capability to make CFD calculations 

with over 1 million grid points in the flow field with a computational 

speed of more than 106 FLOPS. NASA installed similar machines at 

the Langley and Lewis Research Centers. On these machines, NASA 

researchers made the first meaningful three-dimensional inviscid flowfield calculations and significant two-dimensional high Reynolds number calculations. Supercomputers became the engine that propelled CFD 

into the forefront of aerospace design as well as research. Bigger and 

better supercomputers, such as the pioneering Cray-1 and its successor, the Cray X-MP, allowed grids of tens of millions of grid points to be 

used in a flow-field calculation with speeds beginning to approach the 

hallowed goal of gigaflops (109 floating point operations per second). 

Such machines made it possible to carry out numerical solutions of 

the Navier-Stokes equations for three-dimensional fairly high Reynolds 

number viscous flows. The first three-dimensional Navier-Stokes solutions of the complete flow field around a complete airplane at angle of 
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attack came on the scene in the 1980s, enabled by these supercomputers. Subsonic, transonic, supersonic, and hypersonic flow solutions covered the whole flight regime. Again, the major drivers for these solutions 

were the aerospace research and development problems tackled by NASA 

engineers and scientists. This headlong development of supercomputers has continued unabated. The holy grail of CFD researchers in the 

1990s was the teraflop machine (1012 FLOPS); today, it is the petaflop 

(1015 FLOPS) machine. Indeed, recently the U.S. Energy Department has 

contracted with IBM to build a 20-petaflop machine in 2012 for calculations involving the safety and reliability of the Nation’s aging nuclear 

arsenal.14 Such a machine will aid the CFD practitioner’s quest for the 

ultimate flow-field calculations—direct numerical simulation (DNS) of 

turbulent flows, an area of particularly interest to NASA researchers.

Some Seminal Solutions and Applications

We have discussed the historical evolution of the governing flow equations, the first essential element of CFD. We then discussed the evolution of the high-speed digital computer, the second essential element of 

CFD. We now come to the crux of this article, the actual CFD flow-field 

solutions, their evolution, and their importance. Computational fluid 

dynamics has grown exponentially in the past four decades, rendering any selective examination of applications problematical. This case 

study examines four applications that have driven the development of 

CFD to its present place of prominence: the supersonic blunt body problem, transonic airfoils and wings, Navier-Stokes solutions, and hypersonic vehicles.

The Supersonic Blunt Body Problem

On November 1, 1952, the United States detonated a 10.4-megaton 

hydrogen test device on Eniwetok Atoll in the Marshall Islands, the 

first implementation of physicist Edward Teller’s concept for a “super 

bomb” and a major milestone toward the development of the American 

hydrogen bomb. With it came the need for a new entry vehicle beyond 

the long-range strategic bomber, namely the intercontinental ballistic 

missile (ICBM). This vehicle would be launched by a rocket booster, go 

into a suborbital trajectory in space, and then enter Earth’s atmosphere 

14. “Super Supercomputers,” Aviation Week (Feb. 16, 2009).
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at hypersonic speeds near orbital velocity. This was a brand-new flight 

regime, and the design of the entry vehicle was dominated by an emerging 

design consideration: aerodynamic heating. Knowledge of the existence 

of aerodynamic heating was not new. Indeed, in 1876, Lord Rayleigh 

published a paper in which he noted that the compression process that 

creates a high stagnation pressure on a high-velocity body also results in 

a correspondingly large increase in temperature. In particular, he commented on the flow-field characteristic of a meteor entering Earth’s atmosphere, noting: “The resistance to a meteor moving at speeds comparable 

with 20 miles per second must be enormous, as also the rise of temperature due to the compression of the air. In fact it seems quite unnecessary to appeal to friction in order to explain the phenomena of light and 

heat attending the entrance of a meteor into the earth’s atmosphere.”15

We note that 20 miles per second is a Mach number greater than 100. 

Thus, the concept of aerodynamic heating on very high-speed bodies 

dates back before the 20th century. However, it was not until the middle of the 20th century that aerodynamic heating suddenly became a 

showstopper in the design of high-speed vehicles, initiated by the pressing need to design the nose cones of ICBMs.

In 1952, conventional wisdom dictated that the shape of a missile’s 

nose cone should be a slender, sharp-nosed configuration. This was a 

natural extension of good supersonic design in which the supersonic 

body should be thin and slender with a sharp nose, all designed to reduce 

the strength of the shock wave at the nose and therefore reduce the 

supersonic wave drag. (Among airplanes, the Douglas X-3 Stiletto and 

the Lockheed F-104A Starfighter constituted perfect exemplars of good 

supersonic vehicle design, with long slender fuselage, sharp noses, and 

very thin low aspect ratio [that is, stubby] wings having extremely sharp 

leading edges. This is all to reduce the strength of the shock waves on 

the vehicle. The X-3 and F-104 were the first jet airplanes designed for 

flight at Mach 2, hence their design was driven by the desire to reduce 

wave drag.) With this tradition in mind, early thinking of ICBM nose 

cones for hypersonic flight was more of the same, only more so. On 

the other hand, early calculations showed that the aerodynamic heating to such slender bodies would be enormous. This conventional wisdom was turned on its head in 1951 because of an epiphany by Harry 

15. J.W. Strutt (Lord Rayleigh), “On the Resistance of Fluids,” Philosophical Magazine, ser. 57 

(1876), pp. 430–441.
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Julian Allen (“Harvey” Allen to his friends because of Allen’s delight in 

the rabbit character named Harvey, played by Jimmy Stewart in the 

movie of the same name). Allen was at that time the Chief of the HighSpeed Research Division at the NACA Ames Research Laboratory. One 

day, Harvey Allen walked into the office and simply stated that hypersonic bodies should “look like cannonballs.”

His reasoning was so fundamental and straightforward that it is 

worth noting here. Imagine a vehicle coming in from space and entering the atmosphere. At the edge of the atmosphere the vehicle velocity 

is high, hence it has a lot of kinetic energy (one-half the product of its 

mass and velocity squared). Also, because it is so far above the surface 

of Earth (the outer edge of the atmosphere is about 400,000 feet), it has 

a lot of potential energy (its mass times its distance from Earth times 

the acceleration of gravity). At the outer edge of the atmosphere, the 

vehicle simply has a lot of energy. By the time it impacts the surface of 

Earth, its velocity is zero and its height is zero—no kinetic or potential 

energy remains. Where has all the energy gone? The answer is the only 

two places it could: the air itself and the body. To reduce aerodynamic 

heating to the body, you want more of this energy to go into the air and 

less into the body. Now imagine two bodies of opposite shapes, a very 

blunt body (like a cannonball) and a very slender body (like a needle), 

both coming into the atmosphere at hypersonic speeds. In front of the 

blunt body, there will be a very strong bow shock wave detached from 

the surface with a very high gas temperature behind the strong shock 

(typically about 8,000 kelvins). Hence the air is massively heated by the 

strong shock wave. A lot of energy goes into the air, and therefore, only 

a moderate amount of energy goes into the body. In contrast, in front 

of the slender body there will be a much weaker attached shock wave 

with more moderate gas temperatures behind the shock. Hence the air 

is only moderately heated, and a massive amount of energy is left to 

go into the body. As a result, a blunt body shape will reduce the aerodynamic heating in comparison to a slender body. Indeed, if a slender 

body would be used, the heating would melt and blunt the nose anyway. 

This was Allen’s thinking. It led to the use of blunt noses on all modern 

hypersonic vehicles, and it stands as one of the most important aerodynamic contributions of the NACA over its history.

When Allen introduced his blunt body concept in the early 1950s, 

there were no theoretical solutions of the flow over a blunt body moving at supersonic or hypersonic speeds. In the flow behind the strong 
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curved bow shock wave, the flow behind the almost vertical portion of the 

shock near the centerline is subsonic, and that behind the weaker, more 

inclined part of the shock wave further above the centerline is supersonic. There were no pure theoretical solutions to this flow. Numerical 

solutions of this flow were tried in the 1950s, but all without success. 

Whatever technique worked in the subsonic region of the flow fell apart in 

the supersonic region, and whatever technique worked in the supersonic 

region of the flow fell apart in the subsonic region. This was a potential 

disaster, because the United States was locked in a grim struggle with 

the Soviet Union to field and employ intercontinental and intermediate-range ballistic missiles, and the design of new missile nose cones 

desperately needed solutions of the flow over the body were the United 

States to ever successfully field a strategic missile arsenal.

On the scene now crept CFD. A small ray of hope came from one 

of the NACA’s and later NASA’s most respected theoreticians, Milton O. 

Van Dyke. Spurred by the importance of solving the supersonic blunt 

body problem, Van Dyke developed an early numerical solution for the 

blunt body flow field using an inverse approach: take a curved shock 

wave of given shape, calculate the flow behind the shock, and solve for 

the shape of the body that would generate the assumed shock shape. 

In turn, the flow over a blunt body of given shape could be approached 

by repetitive applications of this inverse solution, eventually converging to the shape of interest. If critical, it was nevertheless a potentially 

tedious task that could have consumed thousands of hours by hand 

calculation, but by using the early IBM computers at Ames, Van Dyke 

was able to obtain the first reliable numerical solution of the supersonic blunt body flow field, publishing his pioneering work in the first 

NASA Technical Report issued after the establishment of the Agency.16

Van Dyke’s solution constituted the first important and practical use of 

CFD but was not without limitations. Although the first major advancement toward the solution of the supersonic blunt body problem, it was 

only half a loaf. His procedure worked well in the subsonic region of the 

flow field, but it could penetrate only a small distance into the supersonic region before blowing up. A uniform solution of the whole flow 

field, including both the subsonic and supersonic regions, was still not 

obtainable. The supersonic blunt body problem rode into the decade of 

16. Milton O. Van Dyke, “Supersonic Flow Past a Family of Blunt Axisymmetric Bodies,” NASA 

Technical Report R-1 (1959).
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the 1960s as daunting as it ever was. Then came the breakthrough, which 

was both conceptual and numerical.

First the conceptual breakthrough: at this time the flow was being 

calculated as a steady flow using the Euler equations, i.e., the flow was 

assumed to be inviscid (frictionless). For this flow, the governing partial differential equations of continuity, momentum, and energy (the 

Euler equations) exhibited one type of mathematical behavior (called 

elliptic behavior) in the subsonic region of the flow and a completely 

different type of mathematical behavior (called hyperbolic behavior) 

in the supersonic region of the flow. The equations themselves remain 

identical in these two regions, but the actual behavior of the mathematical solutions is different. (This is no real surprise because the physical behavior of the flow is certainly different between a subsonic and 

a supersonic flow.) This change in the mathematical characteristics of 

the equations was the root cause of all the problems in obtaining a solution to the supersonic blunt body problem. Hence, any numerical solution appropriate for the elliptic (subsonic) region simply was ill-posed 

in the supersonic region, and any numerical solution appropriate for 

the hyperbolic (supersonic) region was ill-posed in the subsonic region. 

Hence, no unified solutions for the whole flow field could be obtained. 

Then, in the middle 1960s, the following idea surfaced: the Euler equations written for an unsteady flow (carrying along the time derivatives 

in the equations) were completely hyperbolic with respect to time no 

matter whether the flow were locally subsonic or supersonic. Why not 

solve the blunt body flow field by first arbitrarily assuming flow-field 

properties at all the grid points, calling this the initial flow field at time 

zero, and then solving the unsteady Euler equations in steps of time, 

obtaining new flow-field values at each new step in time? The problem 

is properly posed because the unsteady equations are hyperbolic with 

respect to time throughout the whole flow field. After continuing this 

process over a large number of time steps, eventually the changes in 

the flow properties from one time step to the next grow smaller, and if 

one goes out to a sufficiently large number of time steps, the flow converges to the steady-state solution. It is this steady-state solution that 

is desired. The time-marching process is simply a means to the end of 

obtaining the solution.17

17. In the old days of obtaining the computer results printed out on paper, I would tell my students to 

tear off the last page, keep it, and throw the rest out, because the last page contained the answer.
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The numerical breakthrough was the implementation of this timemarching approach by means of CFD. Indeed, this process can only be 

carried out in a practical fashion on a high-speed computer using CFD 

techniques. The time-marching approach revolutionized CFD. Today, 

this approach is used for the solution of a whole host of different flow 

problems, but it got its start with the supersonic blunt body problem. The 

first practical implementation of the time-marching idea to the supersonic blunt body was carried out by Gino Moretti and Mike Abbett in 

1966.18 Their work transformed the field of CFD. The supersonic blunt 

body problem in the 1950s and 1960s was worked on by platoons of 

researchers leading to hundreds of research papers at an untold number of conferences, and it cost millions of dollars. Today, because of the 

implementation of the time-marching approach by Moretti and Abbett 

using a finite-difference CFD solution, the blunt body solution is readily carried out in many Government and university aerodynamic laboratories, and is a staple of those aerospace companies concerned with 

supersonic and hypersonic flight. Indeed, this approach is so straightforward that I have assigned the solution of the supersonic blunt body 

problem as a homework problem in a graduate course in CFD. What better testimonial of the power of CFD! A problem that used to be unsolvable and for which much time and money was expended to obtain its 

solution is now reduced to being a “teachable moment” in a graduate 

engineering course.

CFD and Transonic Airfoils

The analysis of transonic flows suffers from the same problems as those 

for the supersonic blunt body discussed above. Just considering the 

flow to be inviscid, the governing Euler equations are highly nonlinear 

for both transonic and hypersonic flows. From the numerical point of 

view, both flow fields are mixed regions of locally subsonic and supersonic flows. Thus, the numerical solution of transonic flows originally 

encountered the same problem as that for the supersonic blunt body 

problem: whatever worked in the subsonic region did not work in the 

supersonic region, and vice versa. Ultimately, this problem was solved 

from two points of view. Historically, the first truly successful CFD solution for the inviscid transonic flow over an airfoil was carried out in 

18. G. Moretti and M. Abbett, “A Time-Dependent Computational Method for Blunt Body Flows,” 

AIAA Journal, vol. 4, no. 12 (1966), pp. 2136–2141. 
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1971 by Earll Murman and Julian Cole of Boeing Scientific Research 

Laboratories, whose collaborative research began at the urging of Arnold 

“Bud” Goldburg, then Chief Scientist of Boeing.19 They treated a simplified version of the Euler equations called the small-perturbation velocity potential equation. This limited their solutions to the flows over thin 

airfoils at small angles of attack. Nevertheless, Murman and Cole introduced the concept of writing the finite differences in the equations such 

that they reached in both the upstream and downstream directions when 

in the subsonic region, but they reached in only the upwind direction in 

the supersonic regions. This is motivated by the physical process that in 

subsonic flow disturbances propagate in all directions but in a supersonic 

flow disturbances propagate only in the downstream direction. Thus it 

is proper to form the finite differences in the supersonic region such that 

they take only information from the upstream side of the grid point.

Today, this approach in modern CFD is called “upwinding” and is 

part of many modern algorithms in use for all kinds of flows. In 1971, this 

idea was groundbreaking, and it allowed Murman and Cole to obtain the 

first successful numerical solutions of the transonic flow over a body. In 

addition to the restriction of thin airfoils at small angles of attack, however, their use of the small perturbation velocity potential equation also 

limited their solutions to isentropic flows. This meant that, although their 

solution captured the semblance of a shock wave in the flow, the location and flow changes across a shock wave were not accurate. Because 

many transonic flows involve shock waves embedded in the flow, this 

was definitely a bit of a problem. The solution to this problem involved 

the numerical treatment of the Euler equations, which, as we have discussed early in this article, accurately pertain to any inviscid flow, not 

just one with small perturbations and free of shocks.

The finest in such CFD solutions were developed by Antony Jameson, 

then a professor at Princeton University (and now at Stanford), whose 

work was heavily sponsored by the NASA Langley Research Laboratory. 

Using the concept of time marching in combination with a Runge-Kutta 

time integration of the unsteady equations, Jameson constructed a series 

of outstanding transonic airfoil codes under the general code name of 

the FLO codes. These codes entered standard use in many aircraft companies and laboratories. Once again, NASA had been responsible for a 

19. Earll Murman and Julian D. Cole, “Calculation of Plane Steady Transonic flows,” AIAA Journal, 

vol. 9, no. 1 (Jan. 1971), pp 114–121.
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major advancement in CFD, helping to develop transonic flow codes 

that advanced the design of many airfoil shapes used today on modern 

commercial jet transports.20

Navier-Stokes CFD Solutions

As described earlier in this article, the Navier-Stokes equations are the 

full equations that govern a viscous flow. Solutions of the Navier-Stokes 

equations are the ultimate in fluid dynamics. To date, no general analytical solutions of these highly nonlinear equations have been obtained. 

Yet they are the equations that reflect the real world of fluid dynamics. 

The only way to obtain useful solutions for the Navier-Stokes equations 

is by means of CFD. And even here such solutions have been slow in 

coming. The problem has been the very fine grids that are necessary to 

define certain regions of a viscous flow (in boundary layers, shear layers, 

separated flows, etc.), thus demanding huge numbers of grid point in the 

flow field. Practical solutions of the Navier-Stokes equations had to wait 

for supercomputers such as the Cray X-MP and Cyber 205 to come on 

the scene. NASA became a recognized and emulated leader in CFD solutions of Navier-Stokes equations, its professionalism evident by its having established the Institute for Computer Applications in Science and 

Engineering (ICASE) at Langley Research Center, though other Centers 

as well, particularly Ames, shared this interest in burgeoning CFD. 21

In particular, NASA researcher Robert MacCormack was responsible 

for the development of a Navier-Stokes CFD code that, by far, became 

the most popular and most widely used Navier-Stokes CFD algorithm 

in the last quarter of the 20th century. MacCormack, an applied mathematician at NASA Ames (and now a professor at Stanford), conceived 

a straightforward algorithm for the solution of the Navier-Stokes equations, simply identified everywhere as “MacCormack’s method.”

To understand the significance of MacCormack’s method, one must 

understand the concept of numerical accuracy. Whenever the derivatives 

in a partial differential equation are replaced by algebraic difference 

20. A. Jameson, W. Schmidt, and E. Turkel, “Numerical Solution of the Euler Equations by Finite 

Volume Methods Using Runge-Kutta Time-Stepping Schemes,” AIAA Paper 81-1259 (1981); see 

also A. Jameson, “Successes and Challenges in Computation Aerodynamics,” AIAA Paper 87-

1184 (1987).

21. Eli Turkel, “Algorithms for the Euler and Navier-Stokes equations for supercomputers,” NASA 

CR-172543 (1985), p. 1. 
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quotients, there is always a truncation error that introduces a degree of 

inaccuracy in the numerical calculations. The simplest finite differences, 

usually involving only two distinct grid points in their formulation, are 

identified as “first-order” accurate (the least accurate formulation). The 

next step up, using a more sophisticated finite difference reaching to 

three grid points, is identified as second-order accurate. For the numerical solution of most fluid flow problems, first-order accuracy is not 

sufficient; not only is the accuracy compromised, but such algorithms 

frequently blow up on the computer. (The author’s experience, however, 

has shown that second-order accuracy is usually sufficient for many 

types of flows.) On the other hand, some of the early second-order algorithms required a large computation effort to obtain this second-order 

accuracy, requiring many pages of paper to write the algorithm and a 

lot of computations to execute the solution. MacCormack developed a 

predictor-corrector two-step scheme that was second-order accurate but 

required much less effort to program and many fewer calculations to 

execute. He introduced this scheme in an imaginative paper on hypervelocity impact cratering published in 1969.22

MacCormack’s method broke open the field of Navier-Stokes solutions, allowing calculation of myriad viscous flow problems, beginning 

in the 1970s and continuing to the present time, as was as well (in this 

author’s opinion) the most “graduate-student friendly” CFD scheme 

in existence. Many graduate students have cut their CFD teeth on this 

method and have been able to solve many viscous flow problems that 

otherwise could not have attempted. Today, MacCormack’s method has 

been supplanted by several very sophisticated modern CFD algorithms, 

but even so, MacCormack’s method goes down in history as one of NASA’s 

finest contributions to the aeronautical sciences.

Three-Dimensional Flows and Hypersonic Vehicles

Three-dimensional flow-field calculation was, for decades, a frustrating impossibility. I recall colleagues in the 1960s who would have 

sold their children (at least they said) to be able to calculate threedimensional flow fields. The number of grid points required for such calculations simply exceeded the capability of any computer at that time. 

With the advent of supercomputers, however, the practical calculation 

22. R.W. MacCormack, “The Effect of Viscosity in Hypervelocity Impact Cratering,” AIAA Paper 

69-354 (1969).
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Steady-state three-dimensional shock wave shape, showing the CFD-generated Mach 16.25 

flow around a Space Shuttle-like body at a reentry angle of attack of over 39 degrees. From 

author’s collection.

of three-dimensional flow fields became realizable. Once again, NASA 

researchers led the way. The first truly three-dimensional flow calculation of real importance was carried out by K.J. Weilmuenster in 1983 at 

the NASA Langley Research Center. He calculated the inviscid flow over 

a Shuttle-like body at angle of attack, including the shape and location 

of the three-dimensional bow shock wave. This was no small feat at the 

time, and it proved to the CFD community that the time had come for 

such three-dimensional calculations.23

This was followed by an even more spectacular success. In 1986, 

using the predictor-corrector method conceived by NASA Ames Research 

Center’s MacCormack, Joseph S. Shang and S.J. Scherr of the Air Force 

Flight Dynamics Laboratory (AFFDL) published the first Navier-Stokes 

calculation of the flow field around a complete airplane. The airplane 

23. K.J. Weilmuenster, “High Angle of Attack Inviscid Flow Calculations of Shuttle-Like Vehicles with 

Comparisons to Flight Data,” AIAA Paper No. 83-1798 (1983). For Weilmuenster’s earlier work, 

see his “Solution of a large hydrodynamic problem using the STAR-100 computer,” NASA TM-X7394 (1976).
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X-24C computed surface streamlines. From author’s collection.

was the “X-24C,” a proposed (though never completed) rocket-powered 

Mach 6+ hypersonic test vehicle conceived by the AFFDL, and the calculation was made for flow conditions at Mach 5.95. The mesh system consisted of 475,200 grid points throughout the flow field, and the explicit 

time-marching procedure took days of computational time on a Cray 

computer. But it was the first such calculation and a genuine watershed 

in the advancement of computational fluid dynamics.24

Note that both of these pioneering three-dimensional calculations 

were carried out for hypersonic vehicles, once again underscoring the 

importance of hypersonic aerodynamics as a major driving force behind 

the development of computational fluid dynamics and of the leading role 

played by NASA in driving the whole field of hypersonics.25

24. J.S. Shang and S.J. Scherr, “Navier-Stokes Solution for a Complete Re-entry Configuration,” 

Journal of Aircraft, vol. 23, no. 12 (1986), pp. 881–888.

25. See P. Perrier, “Industrial Methodologies for the Design of Hypersonic Vehicles,” and Richard D. 

Neumann, “Defining the Aerothermodynamic Methodology,” in J.J. Bertin, et al., Hypersonics, vol. 

1: Defining the Hypersonic Environment (Boston: Birkhäuser Boston, Inc., 1989), pp. 93–160.
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USAF–NASA cooperative CFD Shuttle study. NASA.

Direct Numerical Simulation

What of the future of CFD? Most flows of practical interest are turbulent flows. Turbulence is still one of the few unsolved problems in classical physics. In the calculation of turbulent flows, we therefore have 

to model the effect of turbulence. Any turbulence model involves some 

empirical data, and all models are inaccurate to some greater or lesser 

degree. The uncertainty in turbulence models is the reason for much 

uncertainty in the calculation of turbulent flows in computational fluid 

dynamics. This will continue for years to come. There is, however, an 

approach that requires no turbulence modeling. Nature creates a turbulent flow using the same fundamental principles that are embodied 

in the Navier-Stokes equations. Indeed, turbulence on its most detailed 

scale is simply a flow field developed by nature. If one can put enough 

grid points in the flow, then a Navier-Stokes solution will calculate all the 

detailed turbulence without the need for any type of model. This is called 

direct numerical simulation. The key is “enough grid points,” which 

even for the simplest flow over a flat plate requires millions of points. 
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Once again, NASA researchers have been leading the way. Calculations 

made at NASA Ames for flow over a flat plate have required over 10 million grid points taking hundreds of hours on supercomputers, an indication of what would be required to calculate the whole flow field around 

a complete airplane using direct numerical simulation. But this is the 

future, perhaps, indeed, as far as three decades away. By that time, the 

computational power of computers will have undoubtedly continued 

to increase many-fold, and, as well, NASA will be continuing to play a 

leading role in advancing CFD, even as it is today and has in the past.

Some Important NASA CFD Computer Codes

Not only has NASA played a strong role in the development of new CFD 

algorithms, it has delivered these contributions to the technical public in the form of highly developed computer codes for the user. In the 

context of this survey, it would be remiss not to underscore the importance of these codes, three in particular, which this author (and his students) have used as numerical tools for carrying out research: LAURA, 

OVERFLOW, and CFL3D.

The LAURA code was developed principally by Dr. Peter Gnoffo at 

the NASA Langley Research Laboratory.26 This code solves the threedimensional Euler or Navier-Stokes equations for high-speed supersonic and hypersonic flow fields. It is particularly noteworthy because 

26. The author has had the privilege of looking over Peter Gnoffo’s shoulder for two decades as 

he progressively developed LAURA. Some examples of work carried out with the LAURA code are 

given in Robert B. Greendyke and Peter A. Gnoffo “Convective and Radiative Heating for Vehicle 

Return from the Moon and Mars,” NASA TM-110185 (1995); Kenneth Sutton and Peter A. Gnoffo, 

“Multi-Component Diffusion with Application to Computational Aerothermodynamics,” AIAA Paper 

98-2575 (1998); William L. Kleb, et al., “Collaborative Software Development in Support of 

Fast Adaptive AeroSpace Tools (FAAST),” AIAA Paper 2003-3978 (2003); Peter A. Gnoffo and 

Jeffrey A. White, “Computational Aerothermodynamic Simulation Issues on Unstructured Grids,” 

AIAA Paper 2004-2371 (2004); Peter A. Gnoffo, “Simulation of Stagnation Region Heating in 

Hypersonic Flow on Tetrahedral Grids,” AIAA Paper 2007-3960 (2007); Karen L. Bibb, et al., 

“Parallel, Gradient-Based Anisotropic Mesh Adaption for Re-Entry Vehicle Configurations,” AIAA 

Paper 2006-3579 (2007); Christopher O. Johnson, et al.,“The Influence on Radiative Heating 

for Earth Entry,” NASA Document ID 20080023455 (2008); Richard A. Thompson and Peter A. 

Gnoffo, “Implementation of a Blowing Boundary Condition in the LAURA Code,” NASA Document 

ID 20080008560 (2008); Peter A. Gnoffo, “Multi-Dimensional, Inviscid Flux Reconstruction for 

Simulation of Hypersonic Heating on Tetrahedral Grids,” AIAA Paper 2009-0599 (2009); and 

Gnoffo, Peter A., et al., “Implementation of Radiation, Ablation, and Free Energy Minimization 

Modules for Coupled Simulations of Hypersonic Flow,” AIAA Paper 2009-1399 (2009).
NASA’s Contributions to Aeronautics

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it deals with very detailed nonequilibrium and equilibrium chemically 

reacting flows pertaining to hypersonic reentry vehicles in Earth’s and 

foreign planetary atmospheres. Some applications involve flow-field 

temperatures so high that radiation becomes a dominant physical feature. The LAURA program readily handles radiative gas dynamics, and, 

to this author’s knowledge, it is the only existing standard code to do so. 

The LAURA code has been used for the design and analysis of all NASA 

entry bodies in recent experience and is the most powerful and useful 

code in existence for high-temperature flow fields.

Of particular use for computing lower speed subsonic and transonic 

flows is OVERFLOW. This code was developed in the early 1990s by 

Pieter Burning and Dennis Jesperson as a collaborative effort between 

NASA Johnson Space Center and the NASA Ames Research Center. It 

solves the compressible three-dimensional Reynolds-averaged NavierStokes equations by means of a time-marching algorithm. OVERFLOW 

is widely used for the calculation of three-dimensional subsonic and transonic flows, and it proved particularly valuable for computing subsonic 

viscous flows over airfoils in a recent graduate study of innovative new 

airfoil shapes for high lift undertaken at the University of Maryland at 

College Park’s Department of Aerospace Engineering.27

In the mid-1980s, Dr. Jim Thomas and his colleagues at the NASA 

Langley Research Center recognized the need for a code that contained 

the latest advancements in CFD methodology being developed by the 

applied mathematics community. Out of their interest sprang CFL3D, one 

of the earliest (yet still most powerful) CFD codes developed by NASA.28

27. Judy Conlon, “OVERFLOW Code Empowers Computational Fluid Dynamics,” NASA Insights, 

vol. 5 (Apr. 1998).

28. For more on CFL3D and its applications, see J.L. Thomas, and R.W. Walters, “Upwind 

Relaxation Algorithms for the Navier-Stokes Equations,” AIAA Paper 85-1501-CP (1985); J.L. 

Thomas, B. van Leer, and R. W. Walters, “Implicit Flux-Split Schemes for the Euler Equations,” AIAA 

Paper 85-1680 (1985); W. K. Anderson, J. L. Thomas, and B. van Leer, “Comparison of Finite 

Volume Flux Vector Splitting for the Euler Equations,” AIAA Journal, vol. 24, no. 9 (Sept. 1986), 

pp. 1453–1460; J.L. Thomas, et al., “High-Speed Inlet Flows,” Symposium on Advances and 

Applications in CFD, Winter Annual Meeting of ASME, Chicago, IL, Nov. 1988; W.B. Compton, 

III, J.L. Thomas, W.K. Abeyounis, and M.L. Mason, “Transonic Navier-Stokes Solutions of ThreeDimensional Afterbody Flows,” NASA TM-4111 (1989); J.L. Thomas, “An Implicit Multigrid Scheme 

for Hypersonic Strong-Interaction Flowfields,” Comm. Applied Numerical Methods, vol. 8 (1992), 

pp. 683–693; and J.L. Thomas, “Reynolds Number Effects on Supersonic Asymmetrical Flows over 

a Cone,” Journal of Aircraft, vol. 30 no. 4, (Apr. 1993), pp. 488–495.
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This code is applicable across the whole flight spectrum, from low-speed 

subsonic flow to hypersonic flow. Not only does it handle steady flows, 

but it calculates time-accurate unsteady flows as well. Much effort was 

invested in the development of detailed grids so that it readily handles 

flows over complex three-dimensional bodies. An appreciation of the 

power, usefulness, and widespread acceptance of CFL3D can be gained 

by noting that it is used by over 100 researchers in 22 companies, 13 

universities, NASA, and the military services.

LAURA, OVERFLOW, and CFL3D are just three of the CFD codes 

NASA researchers have generated. Most importantly, because they are the 

product of taxpayer-supported research, all are readily available, free of 

charge, to the general public, making NASA unique among other organizations working in the field of CFD. NASA’s commitment to making scientific and technical information of the highest quality available to the 

public—a legacy of its predecessor, the National Advisory Committee for 

Aeronautics—has influenced its approach to CFD code development and 

may be counted one of the Agency’s most valuable contributions to the 

whole discipline of computational fluid dynamics. When students and 

professional practitioners alike need viable computer codes for complex 

fluid dynamic applications, they have ready access to such codes and 

the extremely competent individuals who develop them. This is perhaps 

the highest accolade one can pronounce upon NASA’s computational 

fluid dynamics efforts.

In closing, a proper history of CFD would require a lengthy book and a 

greater perspective of the past: something yet impossible, for the history of 

this rather young discipline is still evolving. The challenge is akin to what 

one might have expected trying to write a history of the balloon in the early 

1800s, or a history of flight in 1914. In this case, I have tried to share my 

perspective in an accessible format, based in part on my own experiences 

and on my familiarity with the work of many colleagues, especially those 

within NASA. I have had to leave out so many others and so much great 

work in CFD just to tell a short story in a limited amount of pages that 

I feel compelled to apologize to those many others that I have not mentioned. To them I would say that their absence from this case certainly does 

not mean their contributions were any less important. But this has been 

an effort to paint a broad-stroke picture, and, like any such picture, it is 

somewhat subjective. My best wishes go out to all those researchers, present and future, who have and will continue to make computational fluid 

dynamics a vital, essential, and lasting tool for the study of fluid dynamics.
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Recommended Additional Readings

Reports, Papers, Articles, and Presentations: 

Anon., “The Fluidity of Thought and Vision: 40 Years of CFD at Iowa 

State,” Innovate [Magazine of the Iowa State University College of 

Engineering] (fall 2007).

K. Anderson, J.L. Thomas, and B. van Leer, “Comparison of Finite Volume 

Flux Vector Splitting for the Euler Equations,” AIAA Journal, vol. 24, 

no. 9 (Sept. 1986), pp. 1453–1460.

Karen L. Bibb, et al., “Parallel, Gradient-Based Anisotropic Mesh Adaption 

for Re-Entry Vehicle Configurations,” AIAA Paper 2006-3579 (2007).

W.B. Compton, III, J.L. Thomas, W.K. Abeyounis, and M.L. Mason, 

“Transonic Navier-Stokes Solutions of Three-Dimensional Afterbody 

Flows,” NASA TM-4111 (1989).

Judy Conlon, “OVERFLOW Code Empowers Computational Fluid 

Dynamics,” NASA Insights, vol. 5 (Apr. 1998).

Peter A. Gnoffo, “Simulation of Stagnation Region Heating in Hypersonic 

Flow on Tetrahedral Grids,” AIAA Paper 2007-3960 (2007).

Peter A. Gnoffo, “Multi-Dimensional, Inviscid Flux Reconstruction for 

Simulation of Hypersonic Heating on Tetrahedral Grids,” AIAA Paper 

2009-0599 (2009).

Peter A. Gnoffo and Jeffrey A. White, “Computational Aerothermodynamic 

Simulation Issues on Unstructured Grids,” AIAA Paper 2004-2371 

(2004).

Peter A. Gnoffo, et al., “Implementation of Radiation, Ablation, and Free 

Energy Minimization Modules for Coupled Simulations of Hypersonic 

Flow,” AIAA Paper 2009-1399 (2009).

Robert B. Greendyke and Peter A. Gnoffo, “Convective and Radiative 

Heating for Vehicle Return from the Moon and Mars,” NASA 

TM-110185 (1995).
Case 7 | NASA and the Evolution of Computational Fluid Dynamics

455

7

William L. Kleb, et al., “Collaborative Software Development in Support 

of Fast Adaptive AeroSpace Tools (FAAST),” AIAA Paper 2003-3978 

(2003).

Antony Jameson, “Successes and Challenges in Computation 

Aerodynamics,” AIAA Paper 87-1184 (1987).

Antony Jameson, “Computational Aerodynamics for Aircraft Design,” 

Science, vol. 245 (July 28, 1989), pp. 361–371.

A. Jameson, W. Schmidt, and E. Turkel, “Numerical Solution of the Euler 

Equations by Finite Volume Methods Using Runge-Kutta TimeStepping Schemes,” AIAA Paper 81-1259 (1981).

Christopher O. Johnson, et al., “The Influence on Radiative Heating for 

Earth Entry,” NASA Document ID 20080023455 (2008).

R.W. MacCormack, “The Effect of Viscosity in Hypervelocity Impact 

Cratering,” AIAA Paper 69-354 (1969).

G. Moretti and M. Abbett, “A Time-Dependent Computational Method 

for Blunt Body Flows,” AIAA Journal, vol. 4, no. 12 (1966), pp. 

2136–2141.

E. Murman and J.D. Cole, “Calculation of Plane Steady Transonic flows,” 

AIAA Journal, vol. 9 no. 1, (Jan. 1971), pp 114–121.

C.L.M.H. Navier, “Mémoire sur les du mouvement des fluids,” Mémoires 

de l’Academie Royale des Sciences, no. 6 (1823), pp. 389–416.

J.S. Shang and S.J. Scherr, “Navier-Stokes Solution for a Complete 

Re-entry Configuration,” Journal of Aircraft, vol. 23, no. 12 (1986), 

pp. 881–888.

G.G. Stokes, “On the Theories of the Internal Friction of Fluids in Motion, 

and of the Equilibrium and Motion of Elastic Solids,” Transactions of 

the Cambridge Philosophical Society, vol. 8, no. 22 (1845), p. 287–342.
NASA’s Contributions to Aeronautics

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7

Kenneth Sutton and Peter A. Gnoffo, “Multi-Component Diffusion with 

Application to Computational Aerothermodynamics,” AIAA Paper 

98-2575 (1998).

J.L. Thomas, “An Implicit Multigrid Scheme for Hypersonic StrongInteraction Flowfields,” Comm. Applied Numerical Methods, vol. 8 

(1992), pp. 683–693.

J.L. Thomas, “Reynolds Number Effects on Supersonic Asymmetrical 

Flows over a Cone,” Journal of Aircraft, vol. 30, no. 4 (Apr. 1993), 

pp. 488–495.

J.L. Thomas and R.W. Walters, “Upwind Relaxation Algorithms for the 

Navier-Stokes Equations,” AIAA Paper 85-1501-CP (1985).

J.L. Thomas, B. van Leer, and R.W. Walters, “Implicit Flux-Split Schemes 

for the Euler Equations,” AIAA Paper 85-1680 (1985).

J.L. Thomas, et al., “High-Speed Inlet Flows,” Symposium on Advances 

and Applications in CFD, Winter Annual Meeting of ASME, Chicago, 

IL, Nov. 1988.

Richard A. Thompson and Peter A. Gnoffo, “Implementation of a Blowing 

Boundary Condition in the LAURA Code,” NASA Document ID 

20080008560 (2008).

Eli Turkel, “Algorithms for the Euler and Navier-Stokes equations for 

supercomputers,” NASA CR-172543 (1985).

Milton O. Van Dyke, “Supersonic Flow Past a Family of Blunt 

Axisymmetric Bodies,” NASA Technical Report R-1 (1959).

K.J. Weilmuenster, “Solution of a large hydrodynamic problem using 

the STAR-100 computer,” NASA TM-X-7394 (1976).

K.J. Weilmuenster, “High Angle of Attack Inviscid Flow Calculations 

of Shuttle-Like Vehicles with Comparisons to Flight Data,” AIAA 

Paper 83-1798 (1983).

7
Case 7 | NASA and the Evolution of Computational Fluid Dynamics

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7 7

Books and Monographs:

John D. Anderson, Jr., A History of Aerodynamics (Cambridge: Cambridge 

University Press, 1997).

John D. Anderson, Jr., Computational Fluid Dynamics: The Basics With 

Applications (New York: McGraw-Hill, 1995).

J.J. Bertin, et al., Hypersonics, vol. 1: Defining the Hypersonic Environment 

(Boston: Birkhäuser Boston, Inc., 1989).

J. Blazek, Computational Fluid Dynamics: Principles and Applications

(Oxford: Elsevier, 2001 ed.)

Paul E. Ceruzzi, Beyond the Limits: Flight Enters the Computer Age 

(Cambridge: MIT Press, 1989), p. 15.

Olivier Darrigol, Worlds of Flow: A History of Hydrodynamics from the 

Bernoullis to Prandtl (Oxford: Oxford University Press, 2005).

William F. Durand, ed., Aerodynamic Theory, vol. 1 (Berlin: Julius Springer 

Verlag, 1934).

Michael Eckert, The Dawn of Fluid Dynamics: A Discipline Between Science 

and Technology (Weinheim, Germany: Wiley-VCH Verlag GmbH & 

Co. KGaA, 2006). 

Alfred Gessow and Dana J. Morris, A Survey of Computational Aerodynamics 

in the United States, NASA SP-394 (Washington, DC: NASA, 1977).

Charles Hirsch, Numerical Computation of Internal and External Flows: 

The Fundamentals of Computational Fluid Dynamics (Oxford: Elsevier, 

2007 ed.).

Theodore von Kármán, Aerodynamics (Cornell: Cornell University Press, 

1954).

Culbert B. Laney, Computational Gas Dynamics (Cambridge: Cambridge 

University Press, 1998 ed.).
NASA’s Contributions to Aeronautics

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7

Harvard Lomax, Thomas H. Pulliam, and David W. Zingg, Fundamentals 

of Computational Fluid Dynamics (Berlin: Springer-Verlag, 2001).

John C. Tannehill, Dale A. Anderson, and Richard H. Pletcher, 

Computational Fluid Dynamics and Heat Transfer (London: Taylor 

and Francis, 1997 ed.).

G.A. Tokaty, A History and Philosophy of Fluid Mechanics (Henley, UK: 

G.T. Foulis and Co., Ltd., 1971).

Jiyuan Tu, Guan Heng Yeoh, and Chaoqun Liu, Computational Fluid 

Dynamics: A Practical Approach (Oxford: Elsevier, 2008).

Z.U.A. Warsi, Fluid Dynamics: Theoretical and Computational Approaches

(Boca Raton, FL: CRC Press, 1998 ed.).

John Wendt, ed., Computational Fluid Dynamics: An Introduction (Berlin: 

Springer-Verlag GmbH, 2008 ed.).

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NASTRAN model of the X29A aircraft. NASA.
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NASA and Computational 

Structural Analysis

David C. Aronstein

CASE

8

NASA research has been pivotal in its support of computational analytical methods for structural analysis and design, particularly through 

the NASTRAN program. NASA Centers have evolved structural analysis programs tailored to their own needs, such as assessing hightemperature aerothermodynamic structural loading for high-performance 

aircraft. NASA-developed structural tools have been adopted throughout the aerospace industry and are available on the Agency Web site.

T

HE FIELD OF COMPUTER METHODS in structural analysis, and the 

contributions of the National Aeronautics and Space Administration 

(NASA) to it, is wide-ranging. Nearly every NASA Center has a structural analysis group in some form. These groups conduct research and 

assist industry in grappling with a broad spectrum of problems. This 

paper is an attempt to show both aspects: the origins, evolution, and 

application of NASA Structural Analysis System (NASTRAN), and the 

variety and depth of other NASA activities and contributions to the field 

of computational structural methods.

In general terms, the goal of structural analysis is to establish that a 

product has the required strength and stiffness—structural integrity—

to perform its function throughout its intended life. Its strength must 

exceed the loads to which the product is subjected, by some safety margin, the value of which depends on the application.

With aircraft, loads derive from level flight, maneuvering flight, gusts, 

landings, engine thrust and torque, vibration, temperature and pressure 

differences, and other sources. Load cases may be specified by regulatory agency, by the customer, and/or by the company practice and experience. Many of the loads depend on the weight of the aircraft, and the 

weight in turn depends on the design of the structure. This makes the 

structural design process iterative. Because of this, and also because a 

large fraction of an aircraft’s weight is not actually accounted for by primary structure, initial weight estimates are usually based on experience 
NASA’s Contributions to Aeronautics

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rather than on a detailed buildup of structural material. A sizing process must be performed to reconcile the predicted empty weight and its 

relationship to the assumed maximum gross weight, with the required 

payload, fuel, and mission performance.1

After the sizing process has converged, the initial design is documented in the form of a three-view drawing with supporting data. From 

there, the process is approximately as follows:

• 

• 

• 

• 

The weights group generates an initial estimate of the 

weights of the major airframe components.

The loads group analyzes the vehicle at the defined 

condition(s) to determine forces, bending moments, 

etc., in the major components and interfaces.

The structures group defines the primary load paths 

and sizes the primary structural members to provide 

the required strength.

Secondary load paths, etc., are defined to the required 

level of detail.

Process details vary between different organizations, but at some 

point, the structural definition reaches a level of maturity to enable a check 

of the initial weight estimate. Then the whole design may be iterated, 

if required. Iteration may also be driven by maturing requirements 

or by evolution in other aspects of the design, e.g., aerodynamics, 

propulsion, etc.

Structural Analysis Prior to Computers

Basic principles of structural analysis—static equilibrium, trusses, 

and beam theory—were known long before computers, or airplanes, 

existed. Bridges, towers and other buildings, and ships were designed 

1. There are many topics touching on structural analysis and loads prediction, which the author has 

not covered. Materials science is excluded, for this paper considers methods for predicting the macroscopic characteristics of a structure, not analyzing the microscopic properties of the materials from 

which it is built. So too are computer-aided design, computational fluid dynamics, multidisciplinary 

optimization, structural test and data analysis techniques, and nondestructive inspection/evaluation, 

aside from passing reference. Though this paper focuses on NASA activities and contributions, there 

is no intent to minimize the importance of contributions from the aircraft industry, the technical software 

industry, universities, and other research organizations.
Case 8 | NASA and Computational Structural Analysis

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by a combination of experience and some amount of analysis—more 

so as designs became larger and more ambitious during and after the 

Industrial Revolution.

With airplanes came much greater emphasis on weight minimization. Massive overdesign was no longer an acceptable means to achieve 

structural integrity. More rigorous analysis and structural sizing was 

required. Simplifications allowed the analysis of primary members 

under simple loading conditions:

• 

• 

• 

• 

• 

Slender beams: axial load, shear, bending, torsion.

Trusses: members carry axial load only, joined to other 

such members at ends.

Simple shells: pressure loading.

Semi-monocoque (skin and stringer) structures: shear 

flow, etc.

Superposition of loading conditions.

With these simplifications, primary structural members could be 

sized appropriately to the expected loads. In the days of wood, wire, 

and fabric, many aircraft structures could be analyzed as trusses: externally braced biplane wings; fuselage structures consisting of longerons, 

uprights, and cross braces, with diagonal braces or wires carrying torsion; landing gears; and engine mounts. As early as the First World War 

and in the 1920s, researchers were working to cover every required aspect 

of the problem: general analysis methods, analysis of wings, horizontal 

and vertical tails, gust loads, test methods, etc. The National Advisory 

Committee for Aeronautics (NACA) contributed significantly to the building of this early body of methodology.2

Structures with redundancy—multiple structural members capable 

of sharing one or more loading components—may be desirable for safety, 

2. A.F. Zahm and L.H. Crook, “Airplane Stress Analysis,” Report No. 82, Aerodynamical Laboratory, 

Bureau of Construction and Repair, U.S. Navy (1918); Roy G. Miller, “Torsion of wing trusses at diving 

speeds,” NACA TR-104, Langley Research Center (1921); F.H. Norton and D.L. Bacon, “The pressure 

distribution over the horizontal tail surfaces of an airplane II,” NACA TR-119, Langley Research Center 

(1921); Norton and W.G. Brown, “Pressure distribution over the rudder and fin of an airplane in 

flight,” NACA TR-149, Langley Research Center (1923); J.C. Hunsaker and E.B. Wilson, “Report on 

Behavior of Aeroplanes in Gusts,” Massachusetts Institute of Technology (1915), (prepared by MIT and 

predating the establishment of the NACA but listed and marked as NACA-TR-1); E.P. Warner, “Static 

Testing and Proposed Standard Specifications,” NACA TN-6, Langley Field, VA (July 1920), p. 1.
NASA’s Contributions to Aeronautics

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but they posed new problems for analysis. Redundant structures cannot 

be analyzed by force equilibrium alone. A conservative simplification, 

often practiced in the early days of aviation, was to analyze the structure with redundant members missing. A more precise solution would 

require the consideration of displacements and “compatibility” conditions: members that are connected to one another must deform in such 

a manner that they move together at the point of connection. Analysis 

was feasible but time-consuming. Large-scale solutions to redundant 

(“statically indeterminate”) structure problems would become practical 

with the aid of computers. Until then, more simplifications were made, 

and specific types of solutions—very useful ones—were developed.

While these analysis methods were being developed, there was a 

lot of airplane building going on without very much analysis at all. In 

the “golden age of aviation,” many airplanes were built in garages or at 

small companies that lacked the resources for extensive analysis. “In 

many cases people who flew the airplanes were the same people who carried out the analysis and design. They also owned the company. There 

was very little of what we now call structural analysis. Engineers were 

brought in and paid—not to design the aircraft—but to certify that the 

aircraft met certain safety requirements.”3

Through the 1930s, as aircraft structures began to be formed out of 

aluminum, the semi-monocoque or skin-and-stringer structure became 

prevalent, and analysis methods were developed to suit. “In the 1930s, 

’40s, and ’50s, techniques were being developed to analyze specific structural components, such as wing boxes and shear panels, with combined 

bending, torsion, and shear loads and with stiffeners on the skins.”4

A number of exact solutions to the differential equations for stress 

and strain in a structural member were known, but these generally exist 

only for very simple geometric shapes and very limited sets of loading 

conditions and boundary conditions. Exact solutions were of little practical value to the aircraft designer or stress analyst. Instead, “free body 

diagrams” were used to analyze structures at selected locations, or “stations.” The structure was considered to be cut by a theoretical plane at 

the station of interest. All loads, applied and inertial, on the portion of 

the aircraft outboard of the cut had to be borne (reacted) by the structure at the cut.

3. W. Jefferson Stroud, e-mail message to author, Mar. 29, 2009.

4. Stroud, e-mail to author.
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In principle, this allowed the stress at any point in the structure 

to be analyzed—given the time to make an arbitrarily large number of 

these theoretical cuts through the aircraft. In practice, free body diagrams were used to analyze the structure at key locations—selected 

fuselage stations, the root, and selected stations of wings and tail surfaces. Structural members were left constant, or tapered appropriately, 

according to experience and judgment, between the analyzed sections. 

For major projects such as airliners or bombers, the analysis would be 

more thorough, and consequently, major design organizations had rooms 

full of people whose jobs were to perform the required calculations.

The NACA also utilized this brute-force approach to large calculations, and the people who performed the calculations—overwhelmingly women—were called “computers.” Annie J. Easley, who worked at 

the NASA Lewis (now Glenn) Research Center starting in 1955, recalls:

. . . we were called computers until we started to get the 

machines, and then we were changed over to either math technicians or mathematicians. . . . The engineers and the scientists are working away in their labs and their test cells, and 

they come up with problems that need mathematical computation. At that time, they would bring that portion to the computers, and our equipment then were the huge calculators, 

where you’d put in some numbers and it would clonk, clonk, 

clonk out some answers, and you would record them by hand. 

Could add, subtract, multiply, and divide. That was pretty 

much what those big machines, those big desktop machines, 

could do. If we needed to find a logarithm or an exponential, 

we then pulled out the tables.5

After World War II, with jet engines pushing aircraft into ever more 

demanding flight regimes, the analytical community sought to keep up. 

The NACA continued to improve the methodologies for calculating loads 

on various parts of an aircraft, and some of the reports generated during 

that time are still used by industry practitioners today. NACA Technical 

Report (TR) 1007, for horizontal tail loads in pitch maneuvers, is a good 

5. Interview of Annie J. Easley by Sandra Johnson, Cleveland, OH, Aug. 21, 2001, pp .2, 8, in 

NASA Oral History Project, on NASA Glenn Research Center History Office Web site at http://

grchistory.grc.nasa.gov/index.cfm, accessed Apr. 29, 2009.
NASA’s Contributions to Aeronautics

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example, although it does not cover all of the conditions required by 

recent airworthiness regulations.6

For structural analysis, energy methods and matrix methods began 

to receive more attention. Energy methods work as follows: one first 

expresses the deflection of a member as a set of assumed shape functions, each multiplied by an (initially unknown) coefficient; expresses the 

total strain energy in terms of these unknown coefficients; and finally, 

finds the values of the coefficients that minimize the strain energy. If the 

shape functions, from which the solution is built, satisfy the boundary 

conditions of the problem, then so does the final solution.

Energy methods were not new. The concept of energy minimization 

was introduced by Lord Rayleigh in the late 19th century and extended by 

Walter Ritz in two papers of 1908 and 1909.7 Rayleigh and Ritz were particularly concerned with vibrations. Carlo Alberto Castigliano, an Italian 

engineer, published a dissertation in 1873 that included two important 

theorems for applying energy principles to forces and static displacements in structures.8

 However, in the early works, the shape functions 

were continuous over the domain of interest. The idea of breaking up 

(discretizing) a complex structure into many simple elements for numerical solution would lead to the concept of finite elements, but for this to 

be useful, computing technology needed to mature.

The Advent of Direct Analog Computers

The first computers were analog computers. Direct analog computers are 

networks of physical components (most commonly, electrical components: 

resistors, capacitors, inductances, and transformers) whose behavior is governed by the same equations as some system of interest that is being modeled. Direct analog computers were used in the 1950s and 1960s to solve 

problems in structural analysis, heat transfer, fluid flow, and other fields.

The method of analysis and the needs that were driving the move from 

classical idealizations such as slender-beam theory toward computational 

6. Henry Pearson, William McGowan, and James Donegan, “Horizontal Tail Loads in Maneuvering Flight,” NACA TR-1007, Langley Aeronautical Laboratory, VA (1950); Interview of Gonzalo 

Mendoza and Zachary Hazen by author, Wichita, KS, Apr. 6, 2009.

7. Arthur W. Leissa, “The Historical Basis of the Rayleigh and Ritz Methods,” in 7th International 

Symposium on Vibrations of Continuous Systems, Zakopane, Poland, July 2009.

8. “Carlo Alberto Castigliano,” at http://www-groups.dcs.st-and.ac.uk/~history/Biographies/

Castigliano.html, accessed May 22, 2009.
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Representation of structural elements by analog circuits. NASA.

methods are well stated in the following passage, from an NACA-sponsored 

paper by Stanley Benscoter and Richard MacNeal (subsequently a 

cofounder of the MacNeal Schwendler Corporation [MSC] and member 

of the NASTRAN development team):

The theory is expressed entirely in terms of first-order difference equations in order that analogous electrical circuits can 

be readily designed and solutions obtained on the Caltech analog computer. . . . In the process of designing thin supersonic 

wings for minimum weight it is found that a convenient construction with aluminum alloy consists of a rather thick skin 

with closely spaced spars and no stringers. Such a wing deflects 

in the manner of a plate rather than as a beam. Internal stress 

distributions may be considerably different from those given 

by beam theory.9

Their implementation of analog circuitry for bending loads is illustrated here and serves as an example of the direct analog modeling 

of structures.10

Direct analog computing had its advocates well into the 1960s. 

“For complex problems [direct analog] computers are inherently faster 

than digital machines since they solve the equations for the several 

nodes simultaneously, while the digital machines solve them sequentially. Direct analogs have, moreover, the advantage of visualization; 

9. Stanley U. Benscoter and Richard H. MacNeal, “Equivalent Plate Theory for a Straight Multicell 

Wing,” NACA TN-2786 (1952), p. 1.

10. Benscoter and MacNeal, “Equivalent Plate Theory,” NACA TN-2786, p. 28.
NASA’s Contributions to Aeronautics

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computer setups as well as programming are more closely related 

to the actual problem and are based primarily on physical insight 

rather than on numerical skills.”11

The advantages came at a price, however. It could take weeks, in 

some cases, to set up an analog computer to solve a particular type of 

problem. And there was no way to store a problem to be revisited at a 

later date. These drawbacks may not have seemed so important when 

there was no other recourse available, but they became more and more 

apparent as the programmable digital computer began to mature.

Hybrid direct-analog/digital computers were hypothesized in the 

1960s: essentially a direct analog computer controlled by a digital 

computer capable of storing and executing program instructions. This 

would have overcome some of the drawbacks of direct analog computers.12 However, this possibility was most likely overtaken by the 

rapid progress of digital computers. At the same time these hybrid analog/digital computers were just being thought about, NASTRAN was 

already in development.

A different type of analog computer—the active-element, or indirect, analog—consisted of operational amplifiers that performed arithmetic operations. These solved programmed mathematical equations, 

rather than mimicking a physical system. Several NACA locations—

including Langley, Ames, and the Flight Research Center (now Dryden 

Flight Research Center)—used analog computers of this type for flight 

simulation. Ames installed its first analog computer in 1947.13 The Flight 

Research Center flight simulators used analog computers exclusively from 

1955 to 1964 and in combination with digital computers until 1975.14

This type of analog computer can be thought of as simply a less precise, 

less reliable, and less versatile predecessor to the digital computer.

Digital Computation Triggers Automated Structural Analysis

In 1946, the ENIAC, “commonly accepted as the first successful highspeed electronic digital computer,” became operational at the University 

11. Victor Paschkis and Frederick Ryder, Direct Analog Computers (New York: John Wiley & Sons, 

Inc., 1968), preface.

12. Paschkis and Ryder, p. 383.

13. Glenn Bugos, Atmosphere of Freedom: Sixty Years at the NASA Ames Research Center, NASA 

SP-4314 (Washington, DC: NASA, 2000), pp. 181–182.

14. Gene Waltman, Black Magic and Gremlins, pp. 1–4.
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Simple example of discretized structure and single element. NASA.

of Pennsylvania.15 It took up as much floor space as a medium-sized 

house and had to be “programmed” by physically rearranging its control connections. Many advances followed rapidly: storing instructions 

in memory, conditional control transfer, random access memory, magnetic core memory, and the transistor-circuit element. With these and 

other advances, digital computers progressed from large and ungainly 

experimental devices to programmable, useful, commercially available 

(albeit expensive) machines by the mid-1950s.16

The FORTRAN programming language was also developed in the 

mid-1950s and rapidly gained acceptance in technical communities. This 

was a “high level language,” which allowed programming instructions to 

be written in terms that an engineer or analyst could understand; a compiler handled the translation into “machine language” that the computer 

could understand. International Business Machines (IBM) developed the 

original FORTRAN language and also some of the early practical digital computers. Other early digital computers were produced by Control 

Data Corporation (CDC) and UNIVAC. These developments made it pos15. Jeremy Meyers, “A Brief History of the Computer (b.c. – 1993a.d.),” on Meyers’s Web site at 

http://www.jeremymeyers.com/comp#EDC, accessed Apr. 29, 2009; see also Paul E. Ceruzzi, 

A History of Modern Computing (Cambridge: MIT Press, 1999), pp. 7, 15, 20–21.

16. Meyers, “A Brief History of the Computer (b.c. – 1993a.d.).”
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sible to take the new methods of structural analysis that were emerging 

and implement them in an automated, repeatable manner.

The essence of these new methods was to treat a structure as a 

finite number of discrete elastic elements, rather than as a continuum. 

Reactions (forces and moments) and deflections are only calculated at 

specific points, called “nodes.” Elements connect the nodes. The stress 

and strain fields in the regions between the nodes do not need to be 

solved in the global analysis. They only need to be solved when developing the element-level solution, and once this is done for a particular type 

of element, that element is available as a prepackaged building block. 

Complex shapes and structures can then be built up from the simple 

elements. A simple example—using straight beam elements to model a 

curved beam structure—is illustrated here.

To find, for example, the relationship between the displacements of 

the nodes and the corresponding reactions, one could do the following 

(called the unit displacement method). First, a hypothetical unit displacement of one node in one degree of freedom (d.o.f.) only is assumed. 

This displacement is transposed into the local element coordinate systems of all affected elements. (In the corresponding figure, this would 

entail the relatively simple transformation between global horizontal 

and vertical displacements, and element axial and transverse displacements. The angular displacements would require no transformation, 

except in some cases a sign change.) The predetermined element stiffness matrices are used to find the element-level reactions. The element 

reactions are then translated back into global coordinates and summed 

to give the total structure reactions—to the single hypothetical displacement. This set of global reactions, plus zeroes for all forces unaffected 

by the assumed displacement, constitutes one column in the “stiffness 

matrix.” By repeating the exercise for every degree of freedom of every 

node, the stiffness matrix can be built. Then the reactions to any set of 

nodal displacements may be found by multiplying the stiffness matrix 

by the displacement vector, i.e., the ordered list of displacements. This 

entails difficult bookkeeping but simple math.

It is more common in engineering, however, to have to find 

unknown displacements and stresses from known applied forces. This 

answer is not possible to obtain so directly. (That is, if the process just 

described seems direct to you. If it does, you are probably an engineer. 

If it seems too trivial to have even mentioned, then you are probably 

a mathematician.)
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Instead, after the stiffness matrix is found, it must be inverted to 

obtain the flexibility matrix. The inversion of large matrices is a science in itself. But it can be done, using a computer, if one has time to 

wait. Most of the science lies in improving the efficiency of the process. 

Another important output is the stress distribution throughout the structure. But this problem has already been solved at the element level for 

a hypothetical set of element nodal displacements. Scaling the generic 

stress distribution by the actual displacements, for all elements, yields 

the stress state throughout the structure.

There are, of course, many variations on this theme and many complexities that cannot be addressed here. The important point is that 

we have gone from an insoluble differential equation to a soluble 

matrix arithmetic problem. This, in turn, has enabled a change from 

individual analyses by hand of local portions of a structure to a modeling effort followed by an automated calculation of the stresses and deflections of the entire structure.

Pioneering papers on discretization of structures were published by 

Alexander Hrennikoff in 1941 at the Massachusetts Institute of Technology 

and by Richard Courant in 1943 at the mathematics institute he founded 

at New York University that would later bear his name. These papers did 

not lead to immediate application, in part perhaps because they were 

ahead of the necessary computational technology and in part because 

they were still somewhat theoretical and had not yet developed a wellformed practical implementation. The first example of what we now call 

the finite element method (FEM) is commonly considered to be a paper 

by M.J. Turner (Boeing), R.W. Clough (University of California at Berkeley, 

Civil Engineering Department), H.C. Martin (University of Washington, 

Aeronautical Engineering Department), and L.J. Topp in 1956.17 This paper 

presented a method for plane stress problems, using triangular elements. 

John Argyris at the University of Stuttgart, Germany, also made important early contributions. The term “finite element method” was actually 

coined by Clough in 1960. The Civil Engineering Department at Berkeley 

became a major center of early finite element methods development.18

17. M.J. Turner, R.W. Clough, H.C. Martin, and L.J. Topp , “Stiffness and deflection analysis of 

complex structures,” Journal of the Aeronautical Sciences, vol. 23 (1956), pp. 805–823.

18. John R. Brauer, What Every Engineer Should Know about Finite Element Analysis (New York: 

Marcel Dekker, 1993), pp. 2–3; Edward R. Champion, Jr., and J. Michael Ensminger, Finite Element Analysis with Personal Computers (New York: Marcel Dekker, Inc., 1988), p. 1.
NASA’s Contributions to Aeronautics

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By the mid-1960s, aircraft companies, computing companies, universities, and Government research centers were beginning to explore the 

possibilities—although the method allegedly suffered some initial lack 

of interest in the academic world, because it bypassed elegant mathematical solutions in favor of numerical brute force.19 However, the practical value could not long be ignored. The following insightful comment, 

made by a research team at the University of Denver in 1966 (working 

under NASA sponsorship), sums up the expectation of the period: “It is 

certain that this concept is going to become one of the most important 

tools of engineering in the future as structures become more complex 

and computers more versatile and available.”20

NASA Spawns NASTRAN, Its Greatest Computational Success

The project to develop a general-purpose finite element structural 

analysis system was conceived in the midst of this rapid expansion of 

finite element research in the 1960s. The development, and subsequent 

management, enhancement, and distribution, of the NASA Structural 

Analysis System, or NASTRAN, unquestionably constitutes NASA’s greatest single contribution to computerized structural analysis—and arguably the single most influential contribution to the field from any source. 

NASTRAN is the workhorse of structural analysis: there may be more 

advanced programs in use for certain applications or in certain proprietary or research environments, but NASTRAN is the most capable generalpurpose, generally available, program for structural analysis in existence 

today, even more than 40 years after it was introduced.

Origins of NASTRAN

In the early 1960s, structures researchers from the various NASA 

Centers were gathering annually at Headquarters in Washington, DC, to 

exchange ideas and coordinate their efforts. They began to realize that 

many organizations—NASA Centers and industry—were independently 

developing computer programs to solve similar types of structural problems. There were several drawbacks to this situation. Effort was being 

duplicated needlessly. There was no compatibility of input and output formats, or consistency of naming conventions. The programs 

19. Thomas J. Butler, “Operating in the Age of NASTRAN ,” p. 1.

20. Anita S. West and William F. Hubka, “Matrix Methods and Automation in Structural Engineering,” NASA CR-71230 (1966), p. 12.
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were only as versatile as the developers cared to make them; the inherent versatility of the finite element method was not being exploited. 

More benefit might be achieved by pooling resources and developing 

a truly general-purpose program. Thomas G. Butler of the Goddard 

Space Flight Center (GSFC), who led the team that developed NASTRAN 

between 1965 and 1970, recalled in 1982:

NASA’s Office of Advanced Research and Technology (OART) 

under Dr. Raymond Bisplinghoff sponsored a considerable 

amount of research in the area of flight structures through its 

operating centers. Representatives from the centers who managed research in structures convened annually to exchange 

ideas. I was one of the representatives from Goddard Space 

Flight Center at the meeting in January 1964. . . . Center after 

center described research programs to improve analysis of 

structures. Shells of different kinds were logical for NASA 

to analyze at the time because rockets are shell-like. Each 

research concentrated on a different aspect of shells. Some 

were closed with discontinuous boundaries. Other shells had 

cutouts. Others were noncircular. Others were partial spans of 

less than 360°. This all seemed quite worthwhile if the products of the research resulted in exact closed-form solutions. 

However, all of them were geared toward making some simplifying assumption that made it possible to write a computer 

program to give numerical solutions for their behavior. . . . 

Each of these computer programs required data organization 

different from every other. . . . Each was intended for exploring localized conditions rather than complete shell-like structures, such as a whole rocket. My reaction to these programs 

was that . . . technology was currently available to give engineering solutions to not just localized shells but to whole, 

highly varied structures. The method was finite elements.21

Doug Michel led the meetings at NASA Headquarters. Butler, Harry 

Runyan of Langley Research Center, and probably others proposed that 

NASA develop its own finite element program, if a suitable one could 

21. Butler, “Operating in the Age of NASTRAN,” p. 2.
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not be found already existing. “The group thought this was a good idea, 

and Doug followed up with forming the Ad Hoc Group for Structural 

Analysis, which was headed by Tom Butler of Goddard,” recalled C. 

Thomas Modlin, Jr., who was one of the representatives from what is 

now Johnson Space Center.22 The committee included representatives 

from all of the NASA Centers that had any significant activity in structural analysis methods at the time, plus an adjunct member from the 

U.S. Air Force at Wright-Patterson Air Force Base, as listed in the accompanying table.23

CENTER REPRESENTATIVE(S)

Ames Richard M. Beam and Perry P. Polentz

Flight Research (now Dryden) Richard J. Rosecrans

Goddard Thomas G. Butler (Chair) and Peter A. Smidinger

Jet Propulsion Laboratory Marshall E. Alper and Robert M. Bamford

Langley Herbert J. Cunningham

Lewis William C. Scott and James D. McAleese

Manned Spacecraft (now Johnson) C. Thomas Modlin, Jr., and William W. Renegar

Marshall Robert L. McComas

Wright-Patterson AFB James Johnson (adjunct member)

After visiting several aerospace companies, all of whom were 

“extremely cooperative and candid,” and reviewing the existing methods, the committee recommended to Headquarters that NASA sponsor the development of its own finite element program “to update the 

analytical capability of the whole aerospace community. The program 

should incorporate the best of the state of the arts, which were currently splintered.”24

The effort was launched, under the management of Butler at the 

Goddard Space Flight Center, to define and implement the General 

Purpose Structural Analysis program. Requirements were collected from 

the information brought from the various Centers, from the industry visits, and from a conference on “Matrix Methods in Structural Mechanics” 

22. C. Thomas Modlin, Jr., e-mail message to author, Apr. 23, 2009.

23. Caleb W. McCormick, The NASTRAN Users’ Manual, NASA SP-222 (1970), p. i.

24. Butler, “Age of NASTRAN,” p. 2.
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held at Wright-Patterson Air Force Base.25 Key requirements included 

the following:26

• 

• 

• 

• 

• 

• 

General-purpose. The system must allow different 

analysis types—static, transient, thermal, etc.—to be performed on the same structural model without alteration.

Problem size. At least 2,000 degrees of freedom for static 

and dynamic analyses alike. (Prior state of the art was 

approximately 100 d.o.f. for dynamic mode analysis and 

100 to 600 d.o.f. for static analysis.)

Modular. Parts of the program could be changed without disrupting other parts.

Open-ended. New types of elements, new analysis modules, and new formats could be added.

Maintainable and capable of being updated.

Machine-independent. Capable of operating on IBM 360, 

CDC 6000 Series, and UNIVAC 1108 (the only 3 commercially available computers capable of performing such 

analysis at the time), and future generations of computers.

After an initial design phase, the implementation contract was 

awarded to a team led by Computer Sciences Corporation (CSC), with 

MacNeal Schwendler Corporation and Martin Baltimore as subcontractors. Coding began in July 1966. Dr. Paul R. Peabody was the principal 

architect of the overall system design. Dr. Richard H. MacNeal (MacNeal 

Schwendler) designed the solution structure, taking each type of solution from physics, to math, to programming, assisted by David Harting. 

Keith Redner was the implementation team lead and head programmer, assisted by Steven D. Wall and Richard S. Pyle. Frank J. Douglas 

coded the element routines and wrote the programmer’s manual. Caleb 

W. McCormick was the author of the user’s manual and supervised 

NASTRAN installation and training. Other members of the development 

team included Stanley Kaufman (Martin Baltimore), Thomas L. Clark, 

25. “Matrix Methods in Structural Mechanics,” Air Force Flight Dynamics Laboratory, Wright-Patterson Air Force Base, OH, (proceedings published Nov. 1966, conference date unknown); abstract 

on Defense Technical Information Center Web site at http://oai.dtic.mil/oai/oai?verb=getRecord&

metadataPrefix=html&identifier=AD0646300, accessed May 25, 2009.

26. Butler, “Age of NASTRAN,” pp. 3–4.
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David B. Hall, Carl Hennrich, and Howard Dielmann. The project staff 

at Goddard included Richard D. McConnell, William R. Case, James B. 

Mason, William L. Cook, and Edward F. Puccinelli.27

NASTRAN embodied many technically advanced features that are 

beyond the scope of this paper (and, admittedly, beyond the scope of 

this author’s understanding), which provided the inherent capability to 

handle large problems accurately and efficiently. It was referred to as 

a “system” rather than just a program by its developers, and for good 

reasons. It had its own internal control language, called Digital Matrix 

Abstraction Programming (DMAP), which gave flexibility in the use of its 

different modules. There were 151,000 FORTRAN statements, equating 

to more than 1 million machine language statements. Twelve prepackaged “rigid formats” permitted multiple types of analysis on the same 

structural model, including statics, steady-state frequency response, 

transient response, etc.28

The initial development of NASTRAN was not without setbacks and 

delays, and at introduction it did not have all of the intended capabilities. But the team stayed focused on the essentials, choosing which features to defer until later and which characteristics absolutely had to be 

maintained to keep NASTRAN true to its intent.29 According to Butler: 

“One thing that must be mentioned about the project, that is remarkable, pertains to the spirit that infused it everywhere. Every man thought 

that he was the key man on the whole project. As it turned out, every 

man was key because for the whole to mesh no effort was inconsequential. The marvelous thing was that every man felt it inside. There was a 

feeling of destiny on the project.”30

That the developers adhered to the original principles to make 

NASTRAN modular, open-ended, and general-purpose—with common formats and interfaces among its different routines—proved to be 

more important in the long term than how many elements and analysis 

27. McCormick, The NASTRAN Users’ Manual, p. iii; Butler, “Age of NASTRAN,” pp. 

4–6; Butler also lists many more members and describes their contributions. At time of writing, document is available in the Tenth NASTRAN Users’ Colloquium, NASA CP-2249, on 

the NASA Technical Reports Server at http://ntrs.nasa.gov/archive/nasa/casi.ntrs.nasa.

gov/198.300.04182_198.300.4182.pdf, accessed Aug. 13, 2009.

28. J.P. Raney, D.J. Weidman, and H.M. Adelman, “NASTRAN: Status, Maintenance, and Future 

Development,” in First NASTRAN Users’ Colloquium, NASA Langley Research Center (1971), p. 1.

29. McCormick, NASTRAN User’s Manual, pp. iii–iv.

30. Butler, “Age of NASTRAN,” p. 7.
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capabilities were available at introduction. Preserving the intended 

architecture ensured that the details could be filled in later.

Early Use and Continuing Development of NASTRAN

The first components of NASTRAN became operational at Goddard in 

May 1968. Distribution to other Centers, training, and a debugging period 

followed through 1969 and into 1970.31 With completion of the initial development, “the management of NASTRAN was transferred to the Langley 

Research Center. The NASTRAN Systems Management Office (NSMO) 

was established in the Structures Division at Langley October 4, 1970.”32

Initial public release followed just 1 month later, in November 1970.

 NSMO responsibilities included:33

• 

• 

• 

• 

• 

Centralized program development (advisory committees).

Coordinating user experiences (bimonthly NASTRAN 

Bulletin and annual Users’ Colloquia).

System maintenance (error correction and essential 

improvements).

Development and addition of new capability.

NASTRAN-focused research and development (R&D).

The actual distribution of NASTRAN to the public was handled by 

the Computer Software Management and Information Center (COSMIC), 

NASA’s clearinghouse for software distribution (which is described 

in a subsequent section of this paper). The price at initial release was 

$1,700, “which covers reproducing and supplying the necessary system 

tapes and documentation.”34 Documentation was published in four volumes, each with a distinct purpose: one for users, one for programmers 

who would be involved in maintenance and subsequent development, a 

theory manual, and finally a volume of demonstration problems. (The 

900-page user’s manual could be obtained from COSMIC for $10, if 

purchased separately from the program itself. The author assumes that 

the other volumes were similarly priced.)35

31. Ibid., pp. 2–3.

32. Raney, Weidman, and Adelman, 1971, in First NASTRAN Users’ Colloquium, p. 1.

33. Ibid., p. 2.

34. Foreword to the First User’s Colloquium, 1971.

35. McCormick, The NASTRAN User’s Manual, pp. vii–viii.
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NASTRAN user community profile in 1974. NASA.

Things were happening quickly. Within the first year after public 

release, NASTRAN was installed on over 60 machines across the United 

States. There were urgent needs requiring immediate attention. “When 

NSMO was established in October 1970, there existed a dire need for 

maintenance of the NASTRAN system. With the cooperation of Goddard 

Space Flight Center, an interim maintenance contract was negotiated 

with Computer Sciences Corporation through a contract in effect at 

GSFC. This contract provided for the essential function of error correction until a contract for full time maintenance could be negotiated 

through an open competition. The interim maintenance activity was 

restricted to the correction of over 75 errors reported to the NSMO, 

together with all associated documentation changes. New thermal bending and hydroelastic elements previously developed by the MacNealSchwendler Corporation under contract to GSFC were also installed. 

Levels 13 and 14 were created for government testing and evaluation. 

The next version of NASTRAN to be released to the public . . . will 

be built upon the results of this interim maintenance activity and will 

be designated Level 15,” according to a status report to the user 

community in 1971.36

36. Raney, Weidman, and Adelman, “NASTRAN: Status, Maintence, and Future Development,” in 

First NASTRAN Users’ Colloquium, NASA Langley Research Center (1971).
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In June 1971, the contract for full-time maintenance was awarded 

to MacNeal Schwendler Corporation, which then opened an office near 

Langley. A bug reporting and correction system was established. Bell 

Aerospace Company received a contract to develop new elements and 

a thermal analysis capability. Other efforts were underway to improve 

efficiency and execution time. A prioritized list of future upgrades was 

started, with input from all of the NASA Centers. However, for the time 

being, the pace of adding new capability would be limited by the need 

to also keep up with essential maintenance.37

By 1975, NASTRAN was installed on 269 computers. The estimated 

composition of the user community (based on a survey taken by the 

NSMO) is illustrated here.

By this time, the NSMO was feeling the pressure of trying to keep 

up with maintenance, bug fixes, and requested upgrades from a large 

and rapidly growing user community. There was also a need to keep 

up with changing hardware technology. Measures under consideration 

included improvements to the Error Correction Information System 

(ECIS); more user involvement in the development of improvements 

(although this would also require effort to enforce coding standards 

and interface requirements, and to develop procedures for verification 

and implementation); and a price increase to help support the NSMO’s 

maintenance costs and also possibly recoup some of the NASTRAN 

development costs. COSMIC eventually changed its terms for all software distribution to help offset the costs of maintenance.

An annual NASTRAN Users’ Colloquium was initiated, the first 

of which occurred approximately 1 year after initial public release. 

Each Colloquium usually began with an overview from the NSMO on 

NASTRAN status, including usage trends, what to expect in the next 

release, and planned changes in NASTRAN management or terms of 

distribution. Other papers covered experiences and lessons learned in 

deployment, installation, and training; technical presentations on new 

types of elements or new solution capabilities that had recently been, 

were being, or could be, implemented; evaluation and comparison of 

NASTRAN with test data or other analysis methods; and user experiences 

and applications. (The early NASTRAN Users’ Colloquia proceedings 

were available from the National Technical Information Service for $6.)

37. Ibid.
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The first Colloquia were held at Langley and hosted by the NSMO 

staff. As the routine became more established and the user community 

grew, the Colloquia were moved to different Centers and cochaired, usually by the current NSMO Manager and a representative from the hosting Center. There were 21 Users’ Colloquia, at which 429 papers were 

presented. The breakdown of papers by contributing organization is 

shown here.(Note: collaborative papers are counted under each contributing organization, so the sum of the subtotals exceeds the overall total.)

ORGANIZATIONS PRESENTING PAPERS AT NASTRAN USERS’ COLLOQUIA

TOTAL PAPERS 429

NASA SUBTOTAL: 91

Goddard 33

Langley 35

Other NASA 23

INDUSTRY SUBTOTAL: 274

Computer and software companies 104

Aircraft and spacecraft industry 116

Nonaerospace industry 54

UNIVERSITIES: 26

OTHER GOVERNMENT SUBTOTAL: 91

Air Force 10

Army 15

Navy 61

National Laboratories 5

Computing companies were typically involved in theory, modeling 

technique, resolution of operational issues, and capability improvements 

(sometimes on contracts to NASA or other agencies), but also collaborated with “user” organizations assisting with NASTRAN application 

to problems of interest. All participants were actively involved in the 

improvement of NASTRAN, as well as its application.

Major aircraft companies—Boeing, General Dynamics, Grumman, 

Lockheed, McDonnell-Douglas, Northrop, Rockwell, and Vought—were 

frequent participants, presenting a total of 70 papers. Smaller aerospace 

companies also began to use NASTRAN. Gates Learjet modeled the 

Lear 35/36 wing as a test case in 1976 and then used NASTRAN in the 
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design phase of the Lear 28/29 and Lear 55 business jets.38 Beechcraft 

used NASTRAN in the design of the Super King Air 200 twin turboprop 

and the T-34C Mentor military trainer.39 Dynamic Engineering, Inc., 

(DEI) began using NASTRAN in the design and analysis of wind tunnel 

models in the 1980s.40

Nonaerospace applications appeared almost immediately. By 1972, 

NASTRAN was being used in the automotive industry, in architectural 

engineering, by the Department of Transportation, and by the Atomic 

Energy Commission. The NSMO had “received expressions of interest 

in NASTRAN from firms in nearly every West European country, Japan, 

and Israel.”41 That same year, “NASTRAN was chosen as the principal 

analytical tool” in the design and construction of the 40-story Illinois 

Center Plaza Hotel building.42

Other nonaerospace applications included:

• 

• 

• 

• 

• 

• 

• 

• 

• 

Nuclear power plants.

Automotive industry, including tires as well as primary 

structure.

Ships and submarines.

Munitions.

Acoustic and electromagnetic applications.

Chemical processing plants.

Steam turbines and gas turbines.

Marine structures.

Electronic circuitry.

B.F. Goodrich, General Motors, Tennessee Eastman, and Texas 

Instruments were common presenters at the Colloquia. Frequent 

Government participants, apart from the NASA Centers, included 

the David Taylor Naval Ship Research & Development Center, 

the Naval Underwater Systems Cener, the U.S. Army Armament 

38. 5th NASTRAN Users’ Colloquium, 1976, pp. 331–352; 8th, 1979, pp. 11–32; 9th, 1980, 

pp. 201–223; 11th, 1983, pp. 226–248; and 13th, 1985, pp. 320–340.

39. “Aircraft Design Analysis,” in Spinoff, NASA, 1979, p. 63.

40. 15th NASTRAN Users’ Colloquium, 1987, pp. 166–183.

41. Rainey and Weidman, “NASTRAN: A Progress Report,” 2nd NASTRAN Users’ Colloquium

(1972), p. 1.

42. 2nd NASTRAN Users’ Colloquium, p. 421.
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Research & Development Command, and several U.S. Army arsenals 

and laboratories.43

Technical improvements, too numerous to describe them all, were 

continually being made. At introduction (Level 12), NASTRAN offered 

linear static and dynamic analysis. There were two main classes of new 

capability: analysis routines and structural elements. Developments were 

often tried out on an experimental basis by users and reported on at the 

Colloquia before being incorporated into standard NASTRAN. Evaluations 

and further improvements to the capability would typically follow. In addition, of course, there were bug fixes and operational improvements. A few 

key developments are identified below. Where dates are given, they represent the initial introduction of a capability into standard NASTRAN, 

not to imply full maturity:

• 

• 

• 

• 

• 

• 

• 

Thermal analysis: initial capability introduced at Level 

15 (1973).

Pre- and post-processing: continuous.

Performance improvements: continuous.

Adaptation to new platforms and operating systems: continuous. (The earliest mention the author has found of 

NASTRAN running on a PC is 1992.44)

New elements: continuous. Level 15 included a dummy 

structural element to facilitate user experimentation.

Substructuring: the decomposition of a larger model into 

smaller models that could be constructed, manipulated, 

and/or analyzed independently. It was identified as an 

important need when NASTRAN was first introduced. 

Initial substructuring capability was introduced at Level 

15 in 1973.

Aeroelastics and flutter: studies were conducted in the 

early 1970s. Initial capability was introduced in Level 16 

by 1976. NASTRAN aeroelastic, flutter, and gust load analysis uses a doublet-lattice aerodynamic method, which 

approximates the wing as an initially flat surface for the 

aerodynamic calculation (does not include camber or 

43. Additional information is provided in the appendixes.

44. Harry G. Schaefer, “Implementation of Mixed Formulation Elements in PC/NASTRAN,” in 21st 

NASTRAN Users’ Colloquium (1992), pp. 1–7.
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thickness). The calculation is much simpler than fullfledged computational fluid dynamics (CFD) analysis but 

neglects many real flow effects as well as configuration 

geometry details. Accuracy is provided by using correction factors to match the static characteristics of the doublet-lattice model to higher fidelity data from flight test, 

wind tunnel test, and/or CFD. One of the classic references 

on correcting lifting surface predictions is a paper by J.P. 

Giesing, T.P. Kalman, and W.P. Rodden of McDonnellDouglas, on contract to NASA, in 1976.45

• 

• 

Automated design and analysis: automated fully stressed 

design was introduced in Level 16 (1976). Design automation is a much broader field than this, however, and most 

attempts to further automate design, analysis, and/or optimization have taken the form of applications outside of, 

and interfacing with, NASTRAN. In many industries, automated design has become routine; in others, the status of 

automated design remains largely experimental, primarily 

because of the inherent complexity of design problems.46

Nonlinear problems: geometric and/or material. 

Geometric nonlinearity is introduced, for example, when 

displacements are large enough to change the geometric configuration of the structure in significant ways. 

Material nonlinearity occurs when local stresses exceed 

the linear elastic limit. Applications of nonlinear analysis include engine hot section parts experiencing regions 

of local plasticity, automotive and aircraft crash simulation, and lightweight space structures that may experience large elastic deformations—to name a few. Studies 

and experimental implementations were made during the 1970s. There are many different classes of nonlinear problems encompassed in this category, requiring a variety of solutions, many of which were added to 

standard NASTRAN through the 1980s.

45. J.P. Giesing, T.P. Kalman, and W.P. Rodden, “Correction Factor Techniques for Improving Aerodynamic Prediction Methods,” NASA CR-144967 (1976).

46. Interview of Thomas Christy and John Splichal (aircraft industry) by author, Wichita, KS, May 

12, 2009. 
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Moving to Diversify and Commercialize NASTRAN

All of the improvements described above took time to implement. 

However, many of the using organizations had their own priorities. Several 

organizations therefore developed their own versions of NASTRAN 

for internal use, including IBM, Rockwell, and the David Taylor Naval 

Ship Research & Development Center, not to mention the different 

NASA Centers. These organizations sometimes contracted with software development companies to make enhancements to their internal 

versions. Thus, there developed several centers of expertise forging the 

way forward on somewhat separate paths, but sharing experiences with 

each other at the Users’ Colloquia and other venues. The NSMO did 

not take responsibility for maintenance of these disparate versions but 

did consider capabilities developed elsewhere for inclusion in the standard NASTRAN, with appropriate review. This was possible because 

of the modular structure of NASTRAN to accept new solutions or new 

elements with little or no disruption to anything else, and it allowed the 

NSMO’s standard NASTRAN to keep up, somewhat, with developments 

being made by various users.

The first commercial version was announced by MacNeal Schwendler 

Corporation in 1971.47 Others followed. SPERRY/NASTRAN was marketed 

by Sperry Support Services in Europe and by Nippon Univac Kaisha, 

Ltd., (NUK) in Japan from 1974 to at least 1986. (Sperry was also the 

UNIVAC parent company—producer of one of the three computers that 

could run NASTRAN when it was first created.) SPERRY/NASTRAN was 

developed from COSMIC NASTRAN Level 15.5.48

At the 10th NASTRAN Users’ Colloquium in 1982, the following commercial versions were identified:49

• 

• 

• 

• 

• 

UAI/NASTRAN (Universal Analytics).

UNIVAC NASTRAN (Sperry).

DTNSRDC NASTRAN (David Taylor Naval Ship Research 

& Development Center).

MARC NASTRAN (Marc Analysis & Research).

NKF NASTRAN (NKF Engineering Associates).

47. “MSC Software Highlights” on MSC Web site at http://www.mscsoftware.com/about/

history.cfm?Q=135&Z=339&Y=346, accessed Feb. 12, 2009.

48. 14th User Colloquium, 1986, p. 1.

49. Butler, “Age of NASTRAN,” p. 7.
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In spite of this proliferation, common input and output formats 

were generally maintained. In 1982, Thomas Butler compared COSMIC 

NASTRAN with the “competition,” which included at that time, in addition to the commercial NASTRAN versions, the following programs: ASKA, 

STRUDL, DET VERITAS NORSKE, STARDYNE, MARC, SPAR, ANSYS, 

SAP, ATLAS, EASE, and SUPERB. He noted that: “during the period in 

which NASTRAN maintenance decisions emphasized the intensive debugging of existing capability in preference to adding new capabilities and conveniences, the competitive programs strove hard to excel in something that 

NASTRAN didn’t. They succeeded. They added some elastic elements, e.g.:

• 

• 

• 

• 

• 

Bending plate with offsets, tapered thickness, and a 10:1 

aspect ratio.

Pipe elbow element.

Tapered beam element.

Membrane without excessive stiffness.

High-order hexagon.

“These new elements make it a bit easier to do some analyses in the 

category of mechanical structures, listed above.”

In addition to new elements, some of the commercial codes added 

such capabilities as dynamic reduction to assist in condensing a largeorder model to a smaller-order one prior to conducting eigenvalue analysis, nonlinear analysis, and packaged tutorials on FEM analysis.

Viewed from the standpoint of the tools that an analyst needs 

. . . NASTRAN Level 17.7 can accommodate him with 95 percent of those tools. . . . The effect of delaying the addition of 

new capability during this ‘scrubbing up’ period is to temporarily lose the ability to serve the analyst in about 5 percent of his 

work with the tools that he needs. In the meantime NASTRAN 

has achieved a good state of health due to the caring efforts of 

P.R. Pamidi [of CSC] and Bob Brugh [of COSMIC].50

The commitment to maintaining the integrity of NASTRAN, rather 

than adding capability at an unsustainable pace, paid off in the long run.

50. Ibid., p. 8.
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Computerized Structural Analysis and Research (CSAR) introduced a version of NASTRAN in 1988.51 However, the trend from the 

1980s through the 1990s was toward consolidation of commercial 

sources for NASTRAN. In 1999, MSC acquired Universal Analytics and 

CSAR. The Federal Trade Commission (FTC) alleged that by these acquisitions, MSC had “eliminated its only advanced NASTRAN competitors.” 

In 2002, MSC agreed to divest at least one copy of its software, with 

source code, to restore competition.52

At the time of this writing, there are several versions of NASTRAN commercially available. Commercial versions have almost completely superseded 

NASA’s own version, although it is still available through the Open Channel 

Foundation (as discussed elsewhere in this paper). Even NASA now uses 

commercial versions of NASTRAN, in addition to other commercial and 

in-house structural analysis programs, when they meet a particular need.53

If one had to sum up the reasons for NASTRAN’s extraordinary history, it might be: ripe opportunity, followed by excellent execution. Finite 

elements were on the cusp. The concepts, and the technology to carry 

them out, were just emerging. The 1960s were the decade in which the 

course of the technology would be determined—splintered, or integrated—

not that every single activity could possibly be brought under one roof. 

But, if a single centerpiece of finite element analysis was going to emerge, 

to serve as a standard and reference point for everything else, it had to 

happen in that decade, before the technical community took off running 

in a myriad of different directions.

In execution, the project was characterized by focus, passion, establishment of rules, and adherence to those rules, all coming together under 

an organization that was dedicated to getting its product out rather 

than hoarding it. Even with these ingredients, successfully producing a 

general-purpose computer program, able to adapt through more than 40 

51. “Stresses and Strains,” Spinoff, 1998.

52. Federal Trade Commission news release, “MSC Software Settles FTC Charges by Divesting 

Nastran [sic] Software,” Aug. 14, 2002, at http://www.ftc.gov/opa/2002/08/mscsoftware.

shtm, accessed May 14, 2009.

53. John P. Gyekenyesi, “SCARE: A Post-Processor Program to MSC/NASTRAN,” abstract from NASA 

Technical Reports Server at http://ntrs.nasa.gov; Karen F. Bartos and Michael A. Ernst, “Evaluation 

of MARC for the Analysis of Rotating Composite Blades,” NASA TM-4423 (1993), p. 1; and Jim 

Loughlin, “FEM and Multiphysics Applications at NASA/GSFC,” NASA Goddard Space Flight Center, 

in 2004 ANSYS Conference, May 24, 2004, Greenbelt, MD; view graph presentation on NASA 

Technical Report Server at http://ntrs.nasa.gov, accessed May 12, 2009.
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years of changing hardware and software technology, was remarkable. 

Staying true to the guiding principles (general-purpose, modular, openended, etc.), even as difficult decisions had to be made and there was 

not time to develop every intended capability, was a crucial quality of the 

development team. In contrast, a team that gets sloppy under time pressure would not have produced a program such lasting value. NASTRAN 

may be one of the closest approaches ever achieved to 100-percent successful technology transition. Not every structural analyst uses NASTRAN, 

but certainly every modern structural analyst knows about it. Those who 

think they need it have access to copious information about it and multiple sources from which they can get it in various forms.

This state of affairs exists in part because of the remarkable nature 

of the product, and in part because of the priority that NASA places on 

the transition of its technology to other sectors. In preparation to address 

the other half of this paper—those many accomplishments that, though 

lesser than NASTRAN, also push the frontier forward in incremental 

steps—we now move to a discussion of those activities in which NASA 

engages for the specific purpose of accomplishing technology transition.

Dissemination and Distribution: NASA Establishes COSMIC

Transitioning technology to U.S. industry, universities, and other 

Government agencies is part of NASA’s charter under the 1958 Space Act. 

Some such transfer happens “naturally” through conferences, journal publications, collaborative research, and other interactions among the technical 

community. However, NASA also has established specific, structured technology transfer activities. The NASA Technology Utilization (TU) program 

was established in 1963. The names of the program’s components and activities have changed over time but have generally included the following:54

• Publications.

• 

• 

• 

• 

Tech briefs: notification of promising technologies.

Technology Utilization compilations.

Small-business announcements.

Technical Support Packages: more detailed 

54. F. Douglas Johnson and Martin Kokus, “NASA Technology Utilization Program: A Summary of Cost 

Benefit Studies,” NASA CR-201936 (1977), pp. i, 3. See also Johnson, Emily Miller, Nancy Gunderson, Panayes Gatseos, Charles F. Mourning, Thomas Basinger, and Kokus, “NASA Tech Brief Program: 

A Cost Benefit Evaluation,” NASA CR-201938 (1977); and Spinoff, 1991, pp. 132–143.
NASA’s Contributions to Aeronautics

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 information about a specific technology, 

 provided on request.

• 

• 

• 

• 

Industrial Applications Centers: university-based services 

to assist potential technology users in searching NASA 

scientific and technical information.

Technology Utilization Officers (TUOs) at the NASA 

Centers to assist interested parties in identifying and 

understanding opportunities for technology transfer.

An Applications Engineering process for developing and 

commercializing specific technologies, once interest has 

been established.

The Computer Software Management and Information 

Center—a university-based center making NASA software and documentation available to industrial clients.

To expedite and enhance its technology transfer mandate, NASA 

established a Computer Software Management and Information Center 

at the University of Georgia at Athens. Within the first few years of the 

Technology Utilization program, it became apparent that many of the 

“technology products” being generated by NASA were computer programs. NASA therefore started COSMIC in 1966 to provide the services 

necessary to manage and distribute computer programs. These services 

included screening programs and documentation for quality and usability; announcing available programs to potential users; and storing, copying, and distributing the programs and documentation. In addition, as the 

collection grew, it was necessary to ensure that each new program added 

capability and was not duplicative with programs already being offered.55

After the first year of operation, COSMIC published a catalog of 113 

programs that were fully operational and documented. Another 11 programs with incomplete documentation and 7 programs missing subroutines were also offered for customers who would find the data useful even 

in an incomplete state. Monthly supplements to the catalog added approximately 20 programs per month.56 By 1971, COSMIC had distributed over 

2,500 software packages and had approximately 900 computer programs 

55. Joseph M. Carlson, “NASA Technology Transfer: The Computer Software Dissemination Program,” in NASTRAN: Users’ Experiences, Langley Research Center (1971), pp. 551–552.

56. Computer Software Management and Information Center (COSMIC), A Directory of Computer 

Programs Available from COSMIC, vol. 1, (Athens, GA: University of Georgia, 1967).
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available. New additions were published in a quarterly Computer Programs 

Abstracts Journal. The collection expanded to include software developed 

by other Government agencies besides NASA. The Department of Defense 

(DOD) joined the effort in 1968, adding DOD-funded computer programs 

that were suitable for public release to the collection.57 In 1981, there were 

1,600 programs available.58 Programs were also withdrawn, because of 

obsolescence or other reasons. During the early 1990s, the collection consisted of slightly more than 1,000 programs.59 NASTRAN, when released 

publicly in 1970, was distributed through COSMIC, as were most of the 

other computer programs mentioned throughout this paper.

Customers included U.S. industry, universities, and other Government 

agencies. Customers received source code and documentation, and unlimited rights to copy the programs for their own use. Initially, the cost to the 

user was just the cost of the media on which the software and documentation were delivered. Basic program cost in 1967 was $75, furnished on 

cards (2,000 or less) or tape. Card decks exceeding 2,000 cards were priced 

on a case-by-case basis. Documentation could also be purchased separately, 

to help the user determine if the software itself was applicable to his or 

her needs. Documentation prices ranged from $1.50 for 25 pages or less, 

to $15 for 300 pages or more.60 Purchase terms eventually changed to a 

lease/license format, and prices were increased somewhat to help defray 

the costs of developing and maintaining the programs. Nevertheless, the 

pricing was still much less than that of commercial software. A costbenefit study, conducted in 1977 and covering the period from 

1971–1976, noted that the operation of COSMIC during that period had 

only cost $1.7 million, against an estimated $43.5 million in benefit provided to users. During that period, there were 21,000 requests for computer 

program documentation, leading to 1,200 computer programs delivered.61

COSMIC operations continued through the 1990s. In 2001, custody 

of the COSMIC collection was transferred to a new organization, Open 

57. Carlson, “NASA Technology Transfer,” pp. 551–552.

58. COSMIC, “COSMIC: 1981 Catalog of Computer Programs,” NASA CR-163916 (1981), 

abstract, on NASA Technical Reports Server at http://ntrs.nasa.gov, accessed Mar. 16, 2009.

59. COSMIC, “Monthly Progress Report,” NASA CR-195809 (1994), pp. 1–3; Innovation, vol. 

1, no. 6 (Nov./Dec. 1993), at http://ipp.nasa.gov/innovation/Innovation16/

SoftwareWarehouse.html, accessed Feb. 11, 2009.

60. COSMIC, A Directory of Computer Programs Available from COSMIC, vol. 1 (Athens, GA: 

University of Georgia, 1967).

61. Johnson and Kokus, “Summary of Cost-Benefit Studies.”
NASA’s Contributions to Aeronautics

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Channel Software (OCS). OCS and a related nonprofit organization, 

the Open Channel Foundation, were started in 1999 at the University 

of Chicago. Originally established to provide dissemination of universitydeveloped software, this effort, like COSMIC, soon grew to include software 

from a broader range of academic and research institutions. The agreement 

to transfer the COSMIC collection to OCS was made through the Robert C. 

Byrd National Technology Transfer Center (NTTC), which itself was established in 1989 and had been working with NASA and other Government 

agencies to facilitate beneficial technology transfer and partnerships.62

Although COSMIC is no longer active, NASA continues to make new 

technical software available to universities, other research centers, and 

U.S. companies that can benefit from its use. User agreements are made 

directly with the Centers where the software is developed. User interfaces and documentation are typically not as polished as commercial 

software, but the level of technology is often ahead of anything commercially available, and excellent support is usually available from the 

research group that has developed the software. The monetary cost is 

minimal or, in many cases, zero.63 Joint development agreements may be 

made if a potential user desires enhancements and is willing to participate in their development. Whether through COSMIC or by other means, 

most of the computer programs discussed in the following sections have 

been made available to U.S. industry and other interested users.

NASA Centers and Their Computational Structural Research

To gain a sense of the types of computational structures projects undertaken by NASA and the contributions of individual Centers to the Agency’s 

efforts, it is necessary to examine briefly the computational structures 

analysis activities undertaken at each Center, reviewing representative 

projects, computer programs, and instances of technology transfer to 

industry—aircraft and otherwise. Projects included the development of 

62. Open Channel Foundation, “About the Open Channel Foundation and Open Channel Software” at http://www.openchannelfoundation.org, accessed Feb. 10, 2009; “NTTC and Open 

Channel Software Launch NASA Software Applications Collection as Space Agency Celebrates 

Birthday,” Oct. 10, 2001; “Open Channel Publishes the NASA COSMIC Collection,” at http://

www.openchannelfoundation.org/cosmic, Open Channel Foundation, accessed Feb. 12, 2009. 

A list of NASA-developed programs for finite element analysis is reported in Appendix D. As of this 

writing, the programs and related tasks are available through the Open Channel Foundation Web 

site. All except GEOFEST were originally released through COSMIC.

63. Interview of Chittur Venkatasubban by author, Wichita, KS, May 14, 2009.
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new computer programs, the enhancement of existing programs, integration of programs to provide new capabilities, and, in some cases, just 

the development of methods to apply existing computer programs to new 

types of problems. The unique missions of the different Centers certainly 

influenced the research, but many valuable developments came from collaborative efforts between Centers and applying tools developed at one 

Center to the problems being worked at another.64

Ames Research Center

The Ames Research Center—with research responsibilities within aerodynamics, aeronautical and space vehicle studies, reentry and thermal 

protection systems, simulation, biomedical research, human factors, 

nanotechnology, and information technology—is one of the world’s premier aerospace research establishments. It was the second NACA laboratory, established in 1939 as war loomed in Europe. The Center was built 

initially to provide for expansion of wind tunnel facilities beyond the 

space and power generation capacity available at Langley. Accordingly, 

in the computer age, Ames became a major center for computational 

fluid dynamics methods development.65 Ames also developed a large 

and active structures effort, with approximately 50 to 100 researchers 

involved in the structural disciplines at any given time.66 Areas of research 

include structural dynamics, hypersonic flight and reentry, rotorcraft, 

and multidisciplinary design/analysis/optimization. These last two are 

discussed briefly below.

64. Jennifer Ross-Nazzal, e-mail to author, Apr. 10, 2009; NASA, 9th NASTRAN Users’ Colloquium; 

Two Centers are not discussed in great detail because they undertake little work in computational structural analysis: the Kennedy Space Center and the Stennis Space Center. With responsibility for spacecraft launch operations, vehicle preparation, and integration, Kennedy maintains a minimal structural 

engineering staff, which is concerned primarily with supporting the Center’s launch equipment. Thus, this 

Center is not active in structural analysis methods research and development, though it did host the 9th 

NASTRAN Users’ Colloquium. The primary mission of Stennis is to test a very specific type of hardware, 

namely the ground-testing of rocket engines. More recently, it has added research in remote sensing. But 

with the exception of a small number of studies related to the operation of the Center’s own test equipment, the Stennis Space Center’s involvement in research and development in computational methods 

is minimal. For more information, see John D. Anderson, Jr., “NASA and the Evolution of Computational 

Fluid Dynamics,” a companion essay in this volume.

65. For more information, see Anderson, “NASA and the Evolution of Computational Fluid 

Dynamics.”

66. Interview of John Gallman by author, Wichita, KS, Apr. 4, 2009.
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In the early 1970s, a joint NASA–U.S. Army rotorcraft program led to a 

significant amount of rotorcraft flight research at Ames. “The flight research 

activity initially concentrated on control and handling issues. . . . Later on, 

rotor aerodynamics, acoustics, vibration, loads, advanced concepts, and 

human factors research would be included as important elements in the 

joint program activity.”67 As is typically the case, this effort impacted the 

direction of analytical work as well in rotor aeroelastics, aeroservoelastics, 

acoustics, rotor-body coupling, rotor air loads prediction, etc. For example, 

a “comprehensive analytical model” completed in 1980 combined structural, inertial, and aerodynamic models to calculate rotor performance, 

loads, noise, vibration, gust response, flight dynamics, handling qualities, 

and aeroelastic stability of rotorcraft.68 Other efforts were less comprehensive and produced specialized methods for treating various aspects of the 

rotorcraft problem, such as blade aeroelasticity.69 The General Rotorcraft 

Aeromechanical Stability Program (GRASP) combined finite elements with 

concepts used in spacecraft multibody dynamics problems, treating the 

helicopter as a structure with flexible, rotating substructures.70

Rotorcraft analysis has to be multidisciplinary, because of the 

many types of coupling that are active. Fixed wing aircraft have not 

always been treated with a multidisciplinary perspective, but the multidisciplinary analysis and optimization of aircraft is a growing field and 

one in which Ames has made many valuable contributions. The Advanced 

Concepts Branch, not directly associated with Structures & Loads but 

responsible for multidisciplinary vehicle design and optimization studies, has performed and/or sponsored much of this work.

A general-purpose optimization program, CONMIN, was developed jointly by Ames and by the U.S. Army Air Mobility Research & 

67. Paul F. Borchers, James A. Franklin, and Jay W. Fletcher, Flight Research at Ames: Fifty-Seven 

Years of Development and Validation of Aeronautical Technology, 1940–1997, NASA SP-3300 

(Washington, DC: NASA, 1998), p. 67.

68. Wayne Johnson, “A Comprehensive Analytical Model of Rotorcraft Aerodynamics and Dynamics, Part III: Program Manual,” NASA TM-81184/AVRADCOM TR-80-A-7, Ames Research Center 

and U.S. Army Aviation Research and Development Command, June 1980, p. vii.

69. F.K. Straub, K.B. Sangha, and B. Panda, “Advance finite element modeling of rotor blade 

aeroelasticity,” American Helicopter Society Journal, vol. 39, no. 2 (Apr. 1994), pp. 56–68.

70. A. Stewart Hopkins and Peter Likins, “Analysis of structures with rotating, flexible substructures,” 

AIAA Paper 87-0951 (1987); D.H. Hodges, A.S. Hopkins, D.L. Kunz, and H.E. Hinnant, “Introduction to GRASP: General Rotorcraft Aeromechanical Stability Program—A Modern Approach to 

Rotorcraft Modeling,” 42nd American Helicopter Society, Washington, DC (1986).
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Development Laboratory in 197371 and had been used extensively by 

NASA Centers and contractors through the 1990s. Garret Vanderplaats 

was the principal developer. Because it is a generic mathematical function minimization program, it can in principle drive any design/analysis 

process toward an optimum. CONMIN has been coupled with many different types of analysis programs, including NASTRAN.72

Aircraft Synthesis (ACSYNT) was an early example of a multidisciplinary aircraft sizing and conceptual design code. Like many early 

(and some current) total-vehicle sizing and synthesis tools, ACSYNT did 

not actually perform structural analysis but instead used empirically 

based equations to estimate the weight of airframe structure. ACSYNT 

was initially released in the 1970s and has been widely used in the aircraft industry and at universities. Collaboration between Ames and the 

Virginia Polytechnic Institute’s CAD Laboratory, to develop a computeraided design (CAD) interface for ACSYNT, eventually led to the commercialization of ACSYNT and the creation of Phoenix Integration, Inc., in 

1995.73 Phoenix Integration is currently a major supplier of analysis integration and multidisciplinary optimization software.

Tools such as ACSYNT are very practical, but it has also been a goal 

at Ames to couple the prediction of aerodynamic forces and loads to 

more rigorous structural design and analysis, which would give more 

insight into the effects of new materials or novel vehicle configurations. 

To this end, a code called ENSAERO was developed, combining finite element structural analysis capability with high-fidelity Euler (inviscid) and 

Navier-Stokes (viscous) aerodynamics solutions. “The code is capable of 

computing unsteady flows on flexible wings with vortical flows,”74 and provisions were made to include control or thermal effects as well. ENSAERO 

was introduced in 1990 and developed and used throughout the 1990s. 

71. Garret N. Vanderplaats, CONMIN: A FORTRAN Program for Constrained Function Minimization, User’s Manual, (Moffett Field, CA: Ames Research Center and U.S. Army Air Mobility R&D 

Laboratory, 1973).

72. Ashish K. Sareen, Daniel P. Schrage, and T.S. Murthy, “Rotorcraft Airframe Structural Optimization for Combined Vibration and Fatigue Constraints,” 47th American Helicopter Society (AHS) 

Forum, Phoenix, AZ (1991).

73. “Aircraft Design Software,” Spinoff, 1997, p. 107; Phoenix Integration, “About Phoenix,” on 

Phoenix Integration Web site at http://www.phoenix-int.com/about/company_profile.php, accessed 

May 11, 2009.

74. G.P. Guruswamy, “ENSAERO: A multidisciplinary Program for Fluid/Structural Interaction Studies of 

Aerospace Vehicles,” Computing Systems in Engineering, vol. 1, nos. 2–4 (1990), pp. 237–256.
NASA’s Contributions to Aeronautics

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In a cooperative project with Virginia Tech and McDonnell-Douglas 

Aerospace, ENSAERO was eventually coupled with NASTRAN to provide 

higher structural fidelity than the relatively limited structural capability 

intrinsic to ENSAERO.75 Guru Guruswamy was the principal developer.

In the late 1990s, Juan Alonso, James Reuther, and Joaquim Martins, 

with other researchers at Ames, applied the adjoint method to the problem of combined aerostructural design optimization. The adjoint method, 

first applied to purely aerodynamic shape optimization in the late 1980s 

by Dr. Antony Jameson, is an approach to optimization that provides 

revolutionary gains in efficiency relative to traditional methods, especially when there are a large number of design variables. It is not an 

exaggeration to say that adjoint methods have revolutionized the art of 

aerodynamic optimization. Technical conferences often contain whole 

sessions on applications of adjoint methods, and several aircraft companies have made practical applications of the technique to the aerodynamic design of aircraft that are now in production.76 Bringing this 

approach to aerostructural optimization is extremely significant.

Dryden Flight Research Center

NASA Dryden has a deserved reputation as a flight research and flighttesting center of excellence. Its personnel had been technically responsible for flight-testing every significant high-performance aircraft since 

the advent of the world’s first supersonic research airplane, the Bell XS-1. 

When this facility first became part of the NACA, as the Muroc Flight Test 

Unit in the late 1940s, there was no overall engineering functional organization. There was a small team attached to each test aircraft, consisting of a project engineer, an engineer, and “computers”—highly skilled 

women mathematicians. There were also three supporting groups: Flight 

Operations (pilots, crew chiefs, and mechanics), Instrumentation, and 

Maintenance. By 1954, however, the High-Speed Flight Station (as it 

was then called) had been organized into four divisions: Research, Flight 

Operations, Instrumentation, and Administrative. The Research division 

included three branches: Stability & Control, Loads, and Performance. 

75. Manoj K. Bhardwaj, “Aeroelastic Analysis of Modern Complex Wings Using ENSAERO and 

NASTRAN,” Progress Report, Virginia Polytechnic Institute and State University (1995); and Rakesh 

K. Kapania, Bhardwaj, Eric Reichenbach, and Guruswamy, “Aeroelastic Analysis of Modern Complex Wings,” AIAA-96-4011 (1996).

76. Author’s experience.
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Shortly thereafter, Instrumentation became Data Systems, to include 

Computing and Simulation (sometimes together, sometimes separately). 

There were changes to the organization, mostly gradual, after that, 

but these essential functions were always present from that time 

forward.77 There are approximately 50 people in the structures, structural 

dynamics, and loads disciplines.78

Analysis efforts at Dryden include establishing safety of flight for the 

aircraft tested there, flight-test and ground-test data analysis, and the 

development and improvement of computational methods for prediction. 

Commercially available codes are used when they meet the need, and inhouse development is undertaken when necessary. Methods development 

has been conducted in the fields of general finite element analysis, reentry problems, fatigue and structural life prediction, structural dynamics 

and flutter, and aeroservoelasticity.

Reentry heating has been an important problem at Dryden since 

the X15 program. Extensive thermal research was conducted during the 

NASA YF-12 flight project, which is discussed in a later section. One very 

significant application of thermal-structural predictive methods was the 

thermal modeling of the Space Shuttle orbiter, using the Lewis-developed 

Structural Performance and Redesign (SPAR) finite element code. Prior 

to first flight, the conditions of the boundary layer on various parts of 

the vehicle in actual reentry conditions were not known. SPAR was used 

to model the temperature distribution in the Shuttle structure, for three 

different cases of aerodynamic heating: laminar boundary layer, turbulent boundary layer, and separated flow. Analysis was based on the space 

transportation system—trajectory 1 (STS-1) flight profile—and results 

were compared with temperature time histories from the first mission. 

The analysis showed that the flight data were best matched under the 

assumption of extensive laminar flow on the lower surface, and partial 

laminar flow on the upper surface. This was one piece of evidence confirming the important realization that laminar boundary layers could exist, 

under conditions of practical interest for hypersonic flight.79

77. Richard P. Hallion and Michael H. Gorn, On the Frontier: Experimental Flight at NASA Dryden 

(Washington, DC: Smithsonian Books, 2002), pp. 344–354.

78. Martin Brenner, e-mail message to author, May 8, 2009.

79. W.L. Ko, R.D. Quinn, L. Gong, L.S. Schuster, and D. Gonzales, “Reentry Heat Transfer Analysis 

of the Space Shuttle Orbiter,” in Howard B. Adelman, ed., NASA Langley Research Center Computational Aspects of Heat Transfer in Structures, NASA CP-2216 (1981), pp. 295–325.
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Dryden has a unique thermal loads laboratory, large enough to house 

an SR-71 or similar-sized aircraft and heat the entire airframe to temperatures representative of high-speed flight conditions. This facility is 

used to calibrate flight instrumentation at expected temperatures and 

also to independently apply thermal and structural loads for the purpose of validating predictive methods or gaining a better understanding 

of the effects of each. It was built during the X15 program in the 1960s 

and is still in use today.

Aeroservoelastics—the interaction of air loads, flexible structures, 

and active control systems—has become increasingly important since 

the late 1970s. As active fly-by-wire control entered widespread use in 

high-performance aircraft, engineers at Dryden worked to integrate control system modeling with finite-element structural analysis and aerodynamic modeling. Structural Analysis Routines (STARS) and other 

programs were developed and improved from the 1980s through the 

present. Recent efforts have addressed the modeling of uncertainty 

and adaptive control.80

At Dryden, much of the technology transfer to industry comes not 

so much from the release of codes developed at Dryden, but from the 

interaction of the contractors who develop the aircraft with the technical groups at Dryden who participate in the analysis and testing. Dryden 

has been involved, for example, in aeroservoelastic analysis of the X-29; 

F15s and F18s in standard and modified configurations (including physical airframe modifications and/or modifications to the control laws); 

High Altitude Long Endurance (HALE) unpiloted vehicles, which have 

their own set of challenges, usually flying at lower speeds but also having longer and more flexible structures than fighter class aircraft; and 

many other aircraft types.

Glenn (Formerly Lewis) Research Center

Glenn is the primary Center for research on all aspects of aircraft and 

spacecraft propulsion, including engine-related structures. The structures area has typically consisted of approximately 50 researchers (not 

80. K.K. Gupta, “STARS—A General-Purpose Finite Element Computer Program for Analysis of 

Engineering Structures,” NASA RP-1129 (1984); Gupta, M.J. Brenner, and L.S. Voelker, “Development of an Integrated Aeroservoelastic Analysis Program and Correlation With Test Data,” NASA 

TP-3120 (1991); and Brenner, e-mail message to author, May 8, 2009.
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counting materials).81 Structures research topics include: structures subjected to thermal loading, dynamic loading, and cyclic loading; spinning 

structures; coupled thermo-fluid-structural problems; structures with 

local plasticity and time-varying properties; probabilistic methods and 

reliability; analysis of practically every part of a turbine engine; Space 

Shuttle Main Engine (SSME) components; propeller and propfan flutter; failed blade containment analysis; and bird impact analysis. Some 

of the impact analysis research has been collaborative with Marshall 

Space Flight Center, which was interested in meteor and space debris 

impact effects on spacecraft.82 Glenn has also collaborated extensively 

with Langley. In 1987, there was a joint Lewis-Langley Workshop on 

Computational Structural Mechanics (CSM) “to encourage a cooperative Langley-Lewis CSM program in which Lewis concentrates on 

engine structures applications, Langley concentrates on airframe and 

space structures applications, and all participants share technology 

of mutual interest.”83

Glenn has been involved in NASTRAN improvements since 

NASTRAN was introduced in 1970 and hosted the sixth NASTRAN Users’ 

Colloquium. Many of the projects at Glenn built supplemental capability for NASTRAN to handle the unique problems of propulsion system 

structural analysis: “The NASA Lewis Research Center has sponsored 

the development of a number of related analytical/computational capabilities for the finite element analysis program, NASTRAN. This development is based on a unified approach to representing and integrating 

the structural, aerodynamic, and aeroelastic aspects of the static and 

dynamic stability and response problems of turbomachines.”84

The aircraft and spacecraft engine industries are naturally the primary customers of Glenn technology. However, no attempt is made here 

to document this technology transfer in detail. Other essays in this volume address advances in propulsion technology and high-temperature 

materials. Instead, attention is given here to those projects at Glenn that 

have advanced the general state of the art in computational structures 

81. Interview of Charles Blankenship by author, Mar. 26, 2009.

82. Search of NASA Technical Reports Server at http://ntrs.nasa.gov, Apr. 20, 2009.

83. Nancy P. Sykes, ed., NASA Workshop on Computational Structural Mechanics, NASA CP10012 (1989).

84. Open Channel Foundation, “AIRLOADS,” on Open Channel Foundation Web site at http://

www.openchannelfoundation.org/projects/AIRLOADS, accessed May 6, 2009.
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methods and that have found other applications in addition to aerospace propulsion. These include SPAR, NESSUS, SCARE/CARE (and 

derivatives), ICAN, and MAC.

SPAR was a finite-element structural analysis system developed initially at NASA Lewis in the early 1970s and upgraded extensively through 

the 1980s. SPAR was less powerful than NASTRAN but relatively interactive and easy to use for tasks involving iterative design and analysis. 

Chrysler Corporation used SPAR for designing body panels, starting in 

the 1980s.85 NASA Langley has made improvements to SPAR and has 

used it for many projects, including structural optimization, in conjunction with the Ames CONMIN program.86 SPAR evolved into the EAL program, which was used for the structural portion of structural-optical 

analyses at Marshall.87 Dryden Flight Research Center has used SPAR 

for Space Shuttle reentry thermal modeling.

Numerical Evaluation of Stochastic Structures under Stress 

(NESSUS) was the product of a Probabilistic Structural Analysis Methods 

(PSAM) project initiated in 1984 for probabilistic structural analysis 

of Shuttle and future spacecraft propulsion system components. The 

prime contractor was Southwest Research Institute (SwRI). NESSUS 

was designed for solving problems in which the loads, boundary conditions, and/or the material properties involved are best described by 

statistical distributions of values, rather than by deterministic (known, 

single) values. PSAM officially completed in 1995 with the delivery of 

NESSUS Version 6.2. SwRI was awarded another contract in 2002 for 

enhancements to NESSUS, leading to the release of Version 8.2 to NASA 

in December 2004 and commercially in 2005. Los Alamos National 

Laboratory has used NESSUS for weapon-reliability analysis under 

its Stockpile Stewardship program. Other applications included automotive collision analysis and prediction of the probability of spinal 

injuries during aircraft ejections, carrier landings, or emergency water 

landings. NESSUS is used in teaching and research at the University 

85. “Auto Design,” Spinoff, NASA, 1986, p. 72.

86. Gary L. Farley and Donald J. Baker, “Graphics and Composite Material Computer Program 

Enhancements for SPAR,” NASA TM-80209 (1980); and J.L. Rogers, Jr., “Programming Structural 

Synthesis System,” NASA Tech Briefs, vol. 10, no. 2 (May 1986).

87. Anees Ahmad and Lamar Hawkins, “Development of Software to Model AXAF-1 Image Quality,” NASA CR-203978 (1996), p. 1.
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of Texas at San Antonio.88 In some applications, NESSUS is coupled with commercially available deterministic codes offering greater 

structural analysis capability, with NESSUS providing the statistically 

derived inputs.89

Ceramics Analysis and Reliability Evaluation of Structures (SCARE/

CARES) was introduced as SCARE in 1985 and later renamed CARES. 

This program performed fast-fracture reliability and failure probability 

analysis of ceramic components. SCARE was built as a postprocessor 

to MSC/NASTRAN. Using MSC/NASTRAN output of the stress state in 

a component, SCARE performed the crack growth and structural reliability analysis of the component.90 Upgrades and a very comprehensive 

program description and user’s guide were introduced in 1990.91 In 1993, 

an extension, CARES/LIFE, was developed to calculate the time dependence of the reliability of a component as it is subjected to testing or 

use. This was accomplished by including the effects of subcritical crack 

growth over time.92 Another 1993 upgrade, CCARES (for CMC CARES), 

added the capability to analyze components made from ceramic matrix 

composite (CMC) materials, rather than just macroscopically isotropic 

materials.93 CARES/PC, introduced in 1994 and made publicly available 

through COSMIC, ran on a personal computer but offered a more limited capability (it did not include fast-fracture calculations).94

R&D Magazine gave an R&D 100 Award jointly to NASA Lewis 

and to Philips Display Components for application of CARES/Life to 

the development of an improved television picture tube in 1995. “Cares/

Life has been in high demand world-wide, although present technology 

88. Cody R. Godines and Randall D. Manteufel, “Probabilistic Analysis and Density Parameter 

Estimation Within Nessus,” NASA CR-2002-212008 (2002).

89. “Structural Analysis Made NESSUSary,” Spinoff, NASA, 2005, p. 94.

90. John P. Gyekenyesi, “SCARE: A Post-Processor Program to MSC/NASTRAN,” abstract from 

NASA Technical Reports Server at http://ntrs.nasa.gov.

91. N.N. Nemeth, et al., “Ceramics Analysis and Reliability,” NASA TP-2916 (1990) pp. 1–3.

92. Nemeth, Lynn M. Powers, L.A. Janosik, and J.P. Gyekenyesi, “Lifetime reliability evaluation of 

structural ceramic parts with the CARES/LIFE computer program,” in 34th AIAA(ASME)ASCE/AHS/

ASC Structures, Structural Dynamics, and Materials Conference, CA, Apr. 19–22, 1993; abstract 

from NASA Technical Reports Server at http://ntrs.nasa.gov.

93. Stephen F. Duffy and John P. Gyekenyesi, “CCARES: A computer algorithm for the reliability 

analysis of laminated CMC components,” NASA TM-111096 (1993).

94. Nemeth, “CARES: Ceramics Analysis and Reliability Evaluation of Structures,” NASA Lewis 

Research Center Report LEW-15168 (1994).
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transfer efforts are entirely focused on U.S.-based organizations. Success 

stories can be cited in numerous industrial sectors, including aerospace, 

automotive, biomedical, electronic, glass, nuclear, and conventional 

power-generation industries.”95

Integrated Composite Analyzer (ICAN) was developed in the early 

1980s to perform design and analysis of multilayered fiber composites. 

ICAN considered hygrothermal (humidity-temperature) conditions as 

well as mechanical loads and provided results for stresses, stress concentrations, and locations of probable delamination.96 ICAN was used 

extensively for design and analysis of composite space antennas and 

for analysis of engine components. Upgrades were developed, including new capabilities and a version that ran on a PC in the early 1990s.97

ICAN was adapted (as ICAN/PART) to analyze building materials under 

a cost-sharing agreement with Master Builders, Inc., in 1995.98

Goodyear began working with Glenn in 1995 to apply Glenn’s 

Micromechanics Analysis Code (MAC) to tire design. The relationship 

was formed, in part, as a result of Glenn’s involvement with the Great 

Lakes Industrial Technology Center (GLITeC) and the Consortium for 

the Design and Analysis of Composite Materials. NASA worked with 

Goodyear to tailor the code to Goodyear’s needs and provided onsite 

training. MAC was used to assess the effects of chord spacing, ply and 

belt configurations, and other tire design parameters. By 2002, Goodyear 

had several tires in production that had benefitted from the MAC 

design analysis capabilities. Dr. Steven Arnold was the Glenn point of 

contact in this effort.99

Goddard Space Flight Center

Goddard Space Flight Center was established in 1959, absorbing the 

U.S. Navy Vanguard satellite project and, with it, the mission of developing, launching, and tracking unpiloted satellites. Since that time, its 

roles and responsibilities have expanded to consider space science, Earth 

observation from space, and unpiloted satellite systems more broadly. 

95. Research & Technology, NASA Lewis Research Center, Cleveland, OH, 1996, p. 100.

96. P.L.N. Murthy and C.C. Chamis, “ICAN: Integrated Composites Analyzer,” 25th Structures, 

Structural Dynamics and Materials Conference, Palm Springs, CA (1984).

97. NASA Technical Reports Server at http://ntrs.nasa.gov, accessed May 14, 2009.

98. Research & Technology, p. 76.

99. “A ‘Tread’ Ahead of the Competition,” Spinoff, NASA, 2002, pp. 72–73.
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Structural analysis problems studied at Goddard included definition of 

operating environments and loads applicable to vehicles, subsystems, and 

payloads; modeling and analysis of complete launch vehicle/payload systems (generic and for specific planned missions); thermally induced loads 

and deformation; and problems associated with lightweight, deployable 

structures such as antennas. Control-structural interactions and multibody dynamics are other related areas of interest.

Goddard’s greatest contribution to computer structural analysis was, 

of course, the NASTRAN program. With public release of NASTRAN, 

management responsibility shifted to Langley. However, Goddard 

remained extremely active in the early application of NASTRAN to 

practical problems, in the evaluation of NASTRAN, and in the ongoing 

improvement and addition of new capabilities to NASTRAN: thermal 

analysis (part of a larger Structural-Thermal-Optical [STOP] program, 

which is discussed below), hydroelastic analysis, automated cyclic symmetry, and substructuring techniques, to name a few.100

Structural-Thermal-Optical analysis predicts the impact on the performance of a (typically satellite-based) sensor system due to the deformation of the sensors and their supporting structure(s) under thermal 

and mechanical loads. After NASTRAN was developed, a major effort 

began at GSFC to achieve better integration of the thermal and optical 

analysis components with NASTRAN as the structural analysis component. The first major product of this effort was the NASTRAN Thermal 

Analyzer. The program was based on NASTRAN and thereby inherited 

a great deal of modeling capability and flexibility. But, most importantly, the resulting inputs and outputs were fully compatible with 

NASTRAN: “Prior to the existence of the NASTRAN Thermal Analyzer, 

available general purpose thermal analysis computer programs were 

designed on the basis of the lumped-node thermal balance method. 

. . . They were not only limited in capacity but seriously handicapped 

by incompatibilities arising from the model representations [lumpednode versus finite-element]. The intermodal transfer of temperature 

100. J.B. Mason, “The NASTRAN Hydroelastic Analyzer,” NASA TM-X-65617 (1972); R.H. MacNeal, R.L. Harder, and Mason, “NASTRAN Cyclic Symmetry Capability,” Goddard Space Flight 

Center, Sept. 1973, abstract on NASA Technical Reports Server at http://ntrs.nasa.gov, accessed 

Apr. 20, 2009; and W. R. Case, “Dynamic Substructure Analysis of the International Ultraviolet 

Explorer (IUE) Spacecraft,” Goddard Space Flight Center, Sept. 1973, abstract on NASA Technical 

Reports Server at http://ntrs.nasa.gov, accessed Apr. 20, 2009.
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data was found to necessitate extensive interpolation and extrapolation. 

This extra work proved not only a tedious and time-consuming process 

but also resulted in compromised solution accuracy. To minimize such 

an interface obstacle, the STOP project undertook the development of 

a general purpose finite-element heat transfer computer program.”101

The capability was developed by the MacNeal Schwendler Corporation 

under subcontract from Bell Aerospace. “It must be stressed, however, 

that a cooperative financial and technical effort between [Goddard and 

Langley] made possible the emergence of this capability.”102

Another element of the STOP effort was the computation of “view 

factors” for radiation between elements: “In an in-house STOP project effort, GSFC has developed an IBM-360 program named ‘VIEW’ 

which computes the view factors and the required exchange coefficients 

between radiating boundary elements.”103 VIEW was based on an earlier view factor program, RAVFAC, but was modified principally for 

compatibility with NASTRAN and eventual incorporation as a subroutine in NASTRAN.104 STOP is still an important part of the analysis of 

many of the satellite packages that Goddard manages, and work continues toward better performance with complex models, multidisciplinary 

design, and optimization capability, as well as analysis.

Jet Propulsion Laboratory

Jet Propulsion Laboratory (JPL) began as an informal group of students 

and staff from the California Institute of Technology (Caltech) who 

experimented with rockets before and during World War II; evolved 

afterward into the Nation’s center for unpiloted exploration of the solar 

system and deep space, operating related tracking, and data acquisition systems; and was managed for NASA by Caltech.105 Dr. Theodore 

von Kármán, then head of Caltech’s Guggenheim Aeronautical 

Laboratory, shepherded this group to becoming a center of rocket 

101. H. Lee and J.B. Mason, “NASTRAN Thermal Analyzer: A General Purpose Finite Element Heat 

Transfer Computer Program,” in NASA Langley Research Center, 2nd NASTRAN Users’ Colloquium

(1972), p. 444.

102. Ibid., p. 445.

103. Ibid., p. 449.

104. E.F. Puccinelli, “View Factor Computer Program (Program VIEW) User’s Manual,” Goddard 

Space Flight Center, July 1973, abstract on NASA Technical Reports Server at http://ntrs.nasa.gov,

accessed Apr. 20, 2009.

105. JPL Annual Reports at http://www.jpl.nasa.gov/about/reports.cfm, accessed Apr. 25, 2009.
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research for the Army. Upon NASA’s formation in 1958, JPL came under 

NASA’s responsibility.106

Consistent with its origins and Caltech’s continuing role in its management, JPL’s orientation has always emphasized advanced experimental and analytical research in various disciplines, including structures. 

JPL developed efficiency improvements for NASTRAN as early as 1971.107

Other JPL research included basic finite element techniques, highvelocity impact effects, effect of spin on structural dynamics, geometrically nonlinear structures (i.e., structures that deflect sufficiently to 

significantly alter the structural properties), rocket engine structural 

dynamics, flexible manipulators, system identification, random processes, 

and optimization. The most notable of these are VISCEL, TEXLESP-S, 

and PID (AU-FREDI and MODE-ID).108

VISCEL (for Visco-Elastic and Hyperelastic Structures) and 

TEXLESP-S treat special classes of materials that general-purpose finite 

element codes typically cannot handle. VISCEL treats visco-elastic problems, in which materials exhibit viscosity (normally a fluid characteristic) as well as elasticity. VISCEL was introduced in 1971 and was adapted 

by industry over the next decade.109 In 1982, the Shell Oil Company 

used VISCEL to validate a proprietary code that was in development 

for the design of plastic products.110 In 1984, AiResearch was using 

VISCEL to analyze seals and similar components in aircraft auxiliary 

power units (APUs).111

JPL has been leading research in the structural dynamics of solid 

rockets almost since the laboratory was first established. TEXLESP-S 

was specifically developed for analysis of solid rocket fuels, which may be 

polymeric materials exhibiting such hyperelastic behavior. TEXLESP-S 

is a finite element code developed for large-strain (hyperelastic) problems, in which materials may be purely elastic but exhibit such large 

106. Jet Propulsion Laboratory, “NASA Facts: Jet Propulsion Laboratory,” NASA, http://www.jpl.

nasa.gov/news/fact_sheets/jpl.pdf, accessed Aug. 13, 2009.

107. R. Levy and S. Wall, “Savings in NASTRAN Decomposition Time by Sequencing to Reduce Active 

Columns,” in NASA Langley Research Center, NASTRAN: Users’ Experiences (1971), pp. 627–631.

108. Search of NASA Technical Reports Server at http://ntrs.nasa.gov, Apr. 22, 2009.

109. F.A. Akyuz and E. Heer, VISCEL Computer Program User Manual for Analysis of Linear Viscoelastic Structures, vol. 1: Users’ Manual, (Pasadena: NASA Jet Propulsion Laboratory, 1971).

110. NASA Scientific and Technical Information, “Computer Technology for Industry,” Spinoff Database at http://www.sti.nasa.gov/spinoff/database, accessed April 25, 2009.

111. NASA Scientific and Technical Information, “Auxiliary Power Units,” Spinoff Database.
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strain deformations that the geometric configuration of the structure 

is significantly altered. (This is distinct from the small-strain, largedeflection situations that can occur, for example, with long flexible 

booms on spacecraft.)112

System Identification/Parameter Identification (PID, including 

AU-FREDI and MODE-ID) is the use of empirical data to build or tune 

a mathematical model of a system. PID is used in many disciplines, 

including automatic control, flight-testing, and structural analysis.113

Ideally, excitation of the system is performed by systematically exciting 

specific modes. However, such controlled excitation is not always practical, and even under the best of circumstances, there is some uncertainty in the interpretation of the data. The MODE-ID program was 

developed in 1988 to estimate not only the modal parameters of a structure, but also the level of uncertainty with which those parameters have 

been estimated:

Such a methodology is presented which allows the precision 

of the estimates of the model parameters to be computed. 

It also leads to a guiding principle in applications. Namely, 

when selecting a single model from a given class of models, 

one should take the most probable model in the class based 

on the experimental data. Practical applications of this principle are given which are based on the utilization of measured 

seismic motions in large civil structures. Examples include 

the application of a computer program MODE-ID to identify 

modal properties directly from seismic excitation and response 

time histories from a nine-story steel-frame building at JPL 

and from a freeway overpass bridge.114

112. Eric B. Becker and Trent Miller, “Final Report: Development of Non-Linear Finite Element Computer Code,” NASA CR-179965 (1985).

113. James L. Beck, “Probabilistic System Identification in the Time Domain,” in USAF–NASA, USAF/

NASA Workshop of Model Determination for Large Space Systems (Pasadena: California Institute of 

Technology, 1988), abstract from NASA Technical Reports Server at http://ntrs.nasa.gov, accessed 

Apr. 24, 2009, notes: “Areas of application for system identification include the following: (1) Model 

Evaluation . . . (2) Model Improvement . . . (3) Empirical Modeling [using experimental data in the initial 

development of a model, when existing methods or information about the system are not sufficient]; and 

(4) Damage Detection and Assessment—continual or episodic updating of a structural model through 

vibration monitoring to detect and locate any structural damage.”

114. Beck, “Probabilistic System Identification.”
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Another system identification program, Autonomous Frequency 

Domain Identification (AU-FREDI), was developed for the identification 

of structural dynamic parameters and the development of control laws 

for large and/or flexible space structures. It was furthermore intended 

to be used for online design and tuning of robust controllers, i.e., to 

develop control laws real time, although it could be modified for offline 

use as well. AU-FREDI was developed in 1989, validated in the Caltech/

Jet Propulsion Laboratory’s Large Spacecraft Control Laboratory and 

made publicly available.115 This is just a small sample of the research 

that JPL has conducted and sponsored in system identification, control 

of flexible structures, integrated control/structural design, and related 

fields. While intended primarily for space structures, this research also 

has relevance for medicine, manufacturing technology, and the design 

and construction of large, ground-based structures.

Johnson Space Center

Johnson Space Center, a product of the “space age,” is NASA’s core center for human space flight, development of launch vehicles and systems, 

astronaut training, and human space flight operations. As a Center with 

significant hardware development and operational responsibilities, 

Johnson’s activities in analysis methods have been “usually directed 

to specific problems relating to developing hardware that the Center 

is responsible for.”116

Except for moderate downsizing in the 1980s and minor organizational changes such as separating Structures and Dynamics into two 

branches, the structures-related organization has been relatively stable over several decades. The Structural Engineering Division (ES) 

has approximately 120 employees divided into 5 branches: Structures, 

Dynamics, Thermal, Material, and Mechanisms. The Structures Branch 

(ES2) has responsibility for structural design, analysis (including computer methods), and testing.117 Johnson has some very significant test 

facilities, including a tower that can hold a full Apollo or similar-sized 

115. Y. Yam, “AU-FREDI—Autonomous Frequency Domain Identification,” JPL, 1994, abstract from 

NASA Technical Reports Server at http://ntrs.nasa.gov, accessed Apr. 25, 2009.

116. According to Modlin, who was the Johnson representative on the Ad Hoc Group for Structural 

Analysis leading up to the development of NASTRAN, in an e-mail message to author, Apr. 23, 2009.

117. Ross-Nazzal, e-mail to author, Apr. 7, 2009.
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vehicle and subject it to vibration testing.118 Current directions at Johnson 

include sustaining activity for the Space Shuttle and the International 

Space Station (ISS), and new work related to the Orion spacecraft.119

With the emphasis on hardware and systems development, rather 

than on methods development, Johnson has favored the use of computer 

programs already available when they can meet the need. According 

to Modlin:

Prior to NASTRAN we used the SAMIS program that was 

developed by JPL for stress and dynamics, our inputs regarding NASTRAN were directed to the NASTRAN office at NASA 

Langley after the program was delivered, but we did not do any 

development on our own. We and our contractor wanted to use 

NASTRAN on the Shuttle Orbiter, but required substructuring. 

This wasn’t delivered in time [as a NASTRAN capability] so the 

contractor continued with ASKA. . . . Some programs developed 

in house relate to: Lunar landing, Apollo Crew Module water 

landing and flight loads. One more general program that has 

wide use is NASGRO (formerly FLAGRO), which was developed by Royce Forman. It is a fracture mechanics routine.120

Although this paper has not attempted to cover fracture mechanics, 

it is worth noting that NASGRO, originally developed for space applications, has been enhanced with “many features specifically implemented 

to suit the needs of the aircraft industry,” because of increasing focus in 

the Federal Aviation Administration (FAA), NASA, and DOD on safety 

of aging aircraft.121

Other programs developed at Johnson or under Johnson sponsorship 

include TRASYS (Thermal Radiation Analysis System, 1973), FAMSOR 

(Frequencies and Modes of Shells Of Revolution, 1974), SNASOR (Static 

118. Interview of Aleck C. Bond by Rebecca Wright (No. 3), Houston, TX, July 15, 1999, in 

JSC Oral Histories project at http://www.jsc.nasa.gov/history/oral_histories/oral_histories.htm, 

accessed Apr. 25, 2009.

119. Modlin, e-mail message to author, Apr. 23, 2009; and Ross-Nazzal, e-mail message to 

author, Apr. 7, 2009.

120. Modlin, e-mail message to author, Apr. 23, 2009.

121. Shivakumar Mettu, et al, “NASGRO 3.0—A Software for Analyzing Aging Aircraft,” in 

NASA–FAA–DOD, 2nd Joint NASA/FAA/DoD Conference on Aging Aircraft, pt. 2, (Washington, 

DC: NASA–FAA–DOD, 1999), pp. 792–801.
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Nonlinear Analysis of Shells of Revolution, 1974), BUCKY (Plate buckling, 1992), and COMPAPP (Composite plate buckling, 1994).122

Langley Research Center

Langley was the first NACA laboratory, established in 1917. As such, it 

is the oldest and most distinguished of NASA aeronautics Centers, with 

a pedigree that dates to meetings held prior to the First World War to 

determine the future aeronautical laboratory structure of the Nation. 

Since the earliest days of American aviation, Langley has constantly 

anticipated, reacted, and adapted as necessary to meet the Nation’s 

aeronautical research needs, reflecting its broad technical capabilities 

and expertise in areas such as aerodynamics, aircraft and spacecraft 

structures, flight dynamics, crew systems, space environmental physics, and life sciences.

Among the very earliest NACA technical reports were several concerning loads calculation and structural analysis, some of which are 

cited in the introduction to this paper. These papers, and others that followed throughout the era of the NACA, were widely used in the aircraft 

industry. By the time NASA was founded, Langley had become a major 

Center for all forms of aeronautics research, engineering, and analysis.

Through the 1980s and 1990s, Langley had approximately 150 

technical professionals in the structural disciplines (not including Materials), covering both aircraft and spacecraft applications. 

This work was organized primarily in two divisions, Structural 

Mechanics (static problems) and Structural Dynamics, plus a separate 

Optimization Methods group of approximately 15 members.123 Structural 

Mechanics included Composites, Computational Structural Mechanics, 

Thermal Structures, Structural Concepts, and AeroThermal Loads.124

Structural Dynamics included Aeroelasticity, Unsteady Aerodynamics, 

Aeroservoelasticity, Landing and Impact Dynamics, Spacecraft Dynamics, 

and Interdisciplinary Research.125 (Reorganizations sometimes changed 

the specific delineation of responsibilities.) Langley researchers pursued 

122. Search of NASA Technical Reports Server at http://ntrs.nasa.gov, Apr. 20, 2009.

123. W.J. Stroud, (NASA Langley, retired), e-mail message to author, Mar. 29, 2009.

124. Kay S. Bales, “Structural Mechanics Division Research and Technology Plans for FY 1989 and 

Accomplishments for FY 1988,” NASA TM-101592 (1989).

125. Eleanor C. Wynne, “Structural Dynamics Division Research and Technology Accomplishments 

for FY 1990 and Plans for FY 1991,” NASA TM-102770 (1991), p. 1.
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many separate computational structural analysis studies and efforts, 

but overall, the Center was particularly (and intimately) involved with 

NASTRAN, the Design Analysis Methods for Vibration (DAMVIBS) 

rotorcraft structural dynamics modeling program, and efforts at 

integration and optimization.

After NASTRAN was developed during the period from 1965 to 1970, 

management of it was transferred from Goddard to Langley. Accordingly, 

a major emphasis at Langley through the 1970s was the maintenance and 

continuing improvement of NASTRAN. The first four Users’ Colloquia 

were held at Langley. While COSMIC handled the administrative aspects 

of NASTRAN distribution, the NSMO was responsible for technical 

management and coordinating NASTRAN development efforts across 

all Centers and many contractors. The program itself is discussed in 

greater detail elsewhere in this case.

The DAMVIBS research program, conducted from 1984 to 1991, 

reflected Langley’s long-standing heritage of research on rotorcraft structural dynamics. DAMVIBS achieved concrete advances in the industry 

state of the art in helicopter structural dynamic modeling, analysis-to-test 

matching, and, perhaps most importantly, acceptance of and confidence 

in modeling as a useful tool in designing helicopter rotor-airframe systems 

for low vibration. Key NASA program personnel were William C. Walton, 

Jr., who spearheaded program concept and initial direction (he retired 

in 1984); Raymond G. Kvaternik, who furnished program direction after 

1984; and Eugene C. Naumann, who supplied critical technical guidance. 

The industry participants were Bell Helicopter Textron, Boeing Helicopters, 

McDonnell-Douglas Helicopter Company, and Sikorsky Aircraft. The 

participants developed rotor-airframe finite element models, conducted 

ground vibration tests, made test/analysis comparisons, improved their 

models, and conducted further study into the “difficult components” that 

current state of the art rotorcraft analysis could not adequately model.126

Modeling “guides”—documented procedures—were identified from 

the start as a key element to the program:

This program emphasized the planning of the modeling . . . 

the NASA Technical Monitor insisted on a well thought out 

126. Raymond G. Kvaternik, “The NASA/Industry Design Analysis Methods for Vibration (DAMVIBS) Program—A Government Overview,” in Kvaternik, ed., A Government/Industry Summary of 

the Design Analysis Methods for Vibrations (DAMVIBS) Program, NASA CP-10114 (1993), p. 9.
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plan of attack, accompanied by detailed preplanned instructions. . . . The plan was reviewed by other industry representatives prior to undertaking the actual modeling. Another unique 

feature was that at the end of the modeling, deviations from 

the planned guides due to cause were reported.127

All of the participants reported that finite element modeling could 

predict vibrations more accurately than previously realized but required 

more attention to detail in the modeling, with finer meshes and the inclusion of secondary components not normally modeled for static strength 

and stiffness analysis. The participants further reported on specific 

improvements to dynamic modeling practice resulting from the exercise and on the increased use and acceptance of such modeling in the 

design phase at each respective company.128 As a result of DAMVIBS:

• 

• 

Bell and Boeing incorporated DAMVIBS lessons into the 

modeling of their respective portions of the V-22.129

Boeing made improvements the NASTRAN dynamic 

model of the CH47D, which was still in production, 

achieving greatly improved correlation to test data. Boeing 

127. R. Gabel, P. Lang, and D. Reed, “The NASA/Industry Design Analysis Methods for Vibration 

(DAMVIBS) Program—Boeing Helicopters Airframe Finite Element Modeling,” in Kvaternik, ed., A Government/Industry Summary of the Design Analysis Methods for Vibrations (DAMVIBS) Program, NASA 

CP-10114 (1993), p. 23.

128. For example, from Bell: “Structural optimization was found to be a useful tool and Bell is continuing with development of this methodology and integrating it into the design process to efficiently 

achieve minimum weight and vibration levels in future designs. The work that was accomplished 

under the NASA DAMVIBS program has had a major influence on the ‘hardening’ and growth of 

vibration technology in the helicopter industry. . . . ” From Sikorsky: “Prior to the DAMVIBS program, 

attempts to reduce [vibration] were usually limited to making modifications or adding vibration-control 

devices to an already designed and built airframe, in a trial-and-error fashion.” The Sikorsky team noted that DAMVIBS-funded work had “brought for the first time the introduction of low-vibration design 

into the design cycle at Sikorsky.” See James D. Cronkhite, “The NASA/Industry Design Analysis 

Methods for Vibration (DAMVIBS) Program—Bell Helicopter Textron Accomplishments,” in Kvaternik, 

ed., A Government/Industry Summary of the Design Analysis Methods for Vibrations (DAMVIBS) Program, NASA CP-10114 (1993), pp. 11, 22; and William J. Twomey, “The NASA/Industry Design 

Analysis Methods for Vibration (DAMVIBS) Program—Sikorsky Aircraft—Advances Toward Interacting 

with the Airframe Design Process,” in Kvaternik, ed., A Government/Industry Summary of the Design 

Analysis Methods for Vibrations (DAMVIBS) Program, NASA CP-10114 (1993), p. 47.

129. Gabel, Lang, and Reed, “DAMVIBS—Boeing,” p. 33. 130. Ibid., p. 33–34.
NASA’s Contributions to Aeronautics

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credited Eugene Naumann of Langley with identifying 

many of the needed changes.130

• 

• 

• 

McDonnell-Douglas improved its dynamic models of existing and newly developed products, achieving improved 

correlation with test results.131

Sikorsky developed an FEM model of the UH60A airframe “having a marked improvement in vibrationpredicting ability.”132

Sikorsky also developed a new program (PAREDYM, 

programmed in NASTRAN DMAP language) that could 

automatically adjust an FEM model so that its modal 

characteristics would match test values.133 PAREDYM 

then found use as a design tool: having the ability to 

modify a model of an existing design to better match 

test data, it also had the ability to modify a model of a 

new design not yet tested, to a set of desired modal characteristics. Designers could now specify a target (low) 

level of vibration response and let PAREDYM tune its 

model—essentially designing the airframe—to meet the 

goal. (The improvements would not be “free,” however, 

as the program could add weight in the process.) After 

discovering this usage mode, the developers then added 

facilities for minimizing the weight impact to achieve a 

desired level of vibration improvement.134

DAMVIBS ended in 1991, though this did not mark an end to Langley’s 

work on rotorcraft structural dynamics.135 Rather, it reflected a shift in 

emphasis away from the traditional helicopter to other aeronautics and 

130. Ibid., p. 33–34.

131. Mostafa Toossi, Richard Weisenburger, and Mostafa Hashemi-Kia, “The NASA/Industry 

Design Analysis Methods for Vibration (DAMVIBS) Program—McDonnell Douglas Helicopter Company Achievements,” in Kvaternik, ed., A Government/Industry Summary of the Design Analysis 

Methods for Vibrations (DAMVIBS) Program, NASA CP-10114 (1993), p. 44.

132. Twomey, “DAMVIBS—Sikorsky,” p. 47.

133. Ibid., p. 47.

134. Ibid., p. 52.

135. William T. Yeager, Jr., “A Historical Overview of Aeroelasticity Branch and Transonic Dynamics Tunnel Contributions to Rotorcraft Technology and Development,” NASA TM-2001-211054 

(2001), p. 38
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astronautics research ventures as well.136 As basic analysis capability had 

become relatively mature by around 1990, attention turned toward the 

integration of design, analysis, and optimization; to the integration of 

structural analysis with other disciplines; and to nondeterministic methods and the modeling of uncertainty.137 Projects included further work on 

rotorcraft, aircraft aerostructural optimization, control-structural optimization for space structures, and nondeterministic or “fuzzy” structures, to 

name a few.138 Many optimization projects at Langley used the CONMIN 

constrained function minimization program, developed at Ames, as the 

optimization driver, interfaced with various discipline-specific analysis 

codes developed at Langley or elsewhere.

In the 1970s, NASA Langley began what would prove to be some 

very significant multidisciplinary optimization (MDO) studies. Jaroslaw 

Sobieszczanski-Sobieski pioneered the Bi-Level Integrated System 

Synthesis (BLISS), a general approach that is applicable to design optimization in any set of disciplines and of any system, aircraft, or otherwise. 

His work at Langley, spanning from the 1970s to the present, is recognized 

throughout the aerospace industry and the MDO community. BLISS and 

related methods constitute one of the major classes of MDO techniques in 

widespread use today. Some of the early work on BLISS was concerned 

with improving the structural design process and addressing aerodynamic 

and structural problems concurrently. For example, in the late 1970s, 

Sobieszczanski-Sobieski developed methods for designing metal and/or 

composite wing structures of supersonic transports for minimum weight, 

including the effect of structural deformations on aeroelastic loads.139

This Langley work continued into the 1980s, when Langley researchers moved forward to apply the knowledge gained with BLISS to spacecraft, generating two other systems: the Integrated Design and Evaluation 

of Advanced Spacecraft (IDEAS) and Programming Structural Synthesis 

(PROSS). IDEAS did not perform optimization per se, but it did provide integration of design with analysis in multiple disciplines, includ136. Kvaternik, “The NASA/Industry Design Analysis Methods for Vibrations (DAMVIBS) Program: 

Accomplishments and Contributions,” NASA TM-104192 (1991), p. 13.

137. W.J. Stroud, (NASA Langley, retired), e-mail message to author, Mar. 28, 2009.

138. Victor W. Sparrow and Ralph D. Buehrle, “Fuzzy Structures Analysis of Aircraft Panels in 

NASTRAN,” AIAA Paper 2001-1320 (2001).

139. Jaroslaw Sobieszczanski-Sobieski, “An Integrated Computer Procedure for Sizing Composite 

Airframe Structures,” NASA TP-1300 (1979).
NASA’s Contributions to Aeronautics

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ing structures and structural dynamics.140 PROSS combined the 

Ames CONMIN optimizer with the SPAR structural analysis program 

(developed at NASA Lewis). PROSS was publicly released in 1983.141

Several subsequent releases incorporated either new optimization 

strategies and/or improved finite element analysis.142

One of these was ST-SIZE, which started as a hypersonic vehicle 

structural-thermal design code. In 1996, Collier Research Corporation 

obtained an exclusive license from Langley for the ST-SIZE program. 

Under a new model for NASA technology transfer, Collier agreed to 

pay NASA royalties from sales of Collier’s commercialized version of the 

code. This version, called HyperSizer (trademark of Collier Research 

Corporation), was intended to be applicable to a wide variety of uses, 

including office design and construction, marine systems, cargo containers, aircraft, and railcars. The program performed design, weight buildup, 

system-level performance assessments, structural analysis, and structural design optimization.143 In 2003, Spinoff reported that this model 

had worked well and that Collier and NASA were still working together 

to enhance the program, specifically by incorporating further analysis 

codes from NASA Glenn Research Center: Micromechanics Analysis Code 

with Generalized Method Cells (MAC/GMC) and higher-order theory for 

functionally graded materials (HOTGFM). Both of these were developed 

collaboratively between Glenn, University of Virginia, Ohio Aerospace 

Institute, and Tel Aviv University.144

Marshall Space Flight Center

Consistent with its mission to develop spacecraft technologies and with 

its heritage as the site where Wernher von Braun and his team had 

140. L. Bernard Garrett, “Interactive Modeling, Design and Analysis of Large Spacecraft,” NASA 

Langley Research Center (1982), on NASA Technical Reports Server at http://ntrs.nasa.gov,

accessed Apr. 4, 2009.

141. J.L. Rogers, Jr., J. Sobieszczanski-Sobieski, and R.B. Bhat, “Structural Optimization,” Tech 

Briefs, vol. 7, no. 2, p. 184, May 1983; abstract on NASA Technical Reports Server at http://

ntrs.nasa.gov, accessed Apr. 4, 2009.

142. Rogers, “NETS/PROSSS—Nets Coupled with the Programming System for Structural Synthesis,” 

LAR-14818, Langley Research Center; abstract on NASA Technical Reports Server at http://ntrs.nasa.

gov, accessed Apr. 4, 2009; and J.L. Rogers, Jr., “System for Structural Synthesis Combines FiniteElement Analysis and Optimization Programs,” NASA Tech Briefs, vol. 8, no. 2 (Nov. 1984), p. 242.

143. “Structural Analysis and Design Software,” Spinoff, NASA, 1997, p. 96.

144. “Efficient, Multi-Scale Design Takes Flight,” Spinoff, NASA, 2003, pp. 68–69.
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worked since 1950, Marshall Space Flight Center has always had a strong 

technical/analytical organization, engaged in science and engineering research as well as advanced design studies. Research areas have 

included basic finite element methods, shells, fluid-structure systems, 

and nonlinear structures, as well as quick-turnaround non-FEM methods for early design and feasibility studies.145

Applications have usually involved the structural and structuraldynamic problems of launch vehicles. As an example, computational 

techniques were used to help resolve “pogo” oscillations in both the first 

and second stages of the Saturn V launch vehicle. As the name implies, 

the pogo mode is a longitudinal tensile/compressive oscillation. Flight 

data from the unpiloted flight of the second Saturn V in 1968 showed 

severe vibrations from 125 to 135 seconds into the first-stage burn. The 

pogo mode is not always harmful, but in this case, there were concerns 

that it could upset the guidance system or damage the payload. The 

structural frequency was dependent on fuel load, and at a certain point 

in the flight, it would coincide with a natural frequency of the engine/

fuel/oxygen system, causing resonance. Using the models to evaluate the 

effects of various design changes, the working group assigned to the task 

determined that accumulators in the liquid oxygen (LOX) lines would 

alter the engine frequency sufficiently to resolve the issue. Subsequently, 

engineers examining flight data from the Apollo 8, 9, and 13 missions 

noticed a similar occurrence in the second stage. This was studied and 

resolved using similar techniques.146

The first-stage pogo issue occurred at a point in the Apollo program 

when time was of the essence in identifying, analyzing, and resolving 

the problem. The computer models were most likely no more complex 

than they had to be to solve the problem at hand. Marshall Space Flight 

Center has continued to develop and use fairly simple codes for early conceptual studies. Simple, quick-turnaround tools developed at Marshall 

include Cylindrical Optimization of Rings, Skin and Stringers (CORSS, 

1994) and the VLOADS launch loads and dynamics program (1997). 

VLOADS was developed as a Visual BASIC macro in Microsoft Excel. 

When released in COSMIC in 1997, it was also available in PC format. 

145. Search of NASA Technical Reports Server at http://ntrs.nasa.gov, Apr. 20, 2009.

146. Andrew J. Dunar and Stephen P. Waring, Power to Explore: A History of the Marshall Space 

Flight Center, 1960–1990, NASA SP-4313 (Washington, DC: NASA, 1999), pp. 44–50.
NASA’s Contributions to Aeronautics

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It was distributed on a single 3.5-inch diskette.147 This was a remarkable development from the days when the problem of launch vehicle 

dynamics occupied a sizable fraction of this Nation’s computing power!

Like researchers at Langley, Marshall’s personnel moved swiftly 

from single or limited application tools to finding ways to integrate them 

with other tools and processes and thereby achieve enhanced or previously unattainable capabilities. The Coupled Eulerian Lagrangian Finite 

Element (CELFE) code, developed collaboratively with NASA Lewis 

Research Center in 1978, included specialized nonlinear methods to calculate local effects of an impact. It was coupled to NASTRAN for calculation of the far-field response of the structure. Applications included 

space debris, micrometeor, and foreign object impact studies for aircraft engines.148 Marshall developed an interface between the PATRAN 

finite element preprocessor (normally used with NASTRAN) and the 

NASA Langley STAGS shell analysis code in 1990.149 Marshall sponsored 

Southwest Research Institute to develop an interface between Lewisdeveloped NESSUS probabilistic analysis and NASTRAN in 1996.150

Both STAGS and NESSUS have been widely used outside NASA. This 

review of NASA Centers and their work on computational structural 

analysis has offered only a glimpse of the variety of structural problems 

that exist and the corresponding variety of methods developed and used 

at the various NASA Centers and then shared with industry.

Applying Computational Structural Analysis to Flight Research

We now turn to an area of activity that provides, for aviation, the ultimate 

proof of design techniques and predictive capabilities: flight-testing. 

While there are many fascinating projects that could be discussed, we will 

consider only five that had particular relevance to the subject at hand, 

either because they collected data that were specifically intended to 

provide validation of computational predictions of structural behavior, or 

because they demonstrated unique structural design approaches.

147. J.B. Graham and P.L. Luz, “Preliminary In-Flight Loads Analysis of In-Line Launch Vehicles Using 

the VLOADS 1.4 Program,” NASA TM-1998-208472 (1998).

148. C. Chamis, “CELFE/NASTRAN code for the analysis of structures subjected to high velocity 

impact,” NASA TM-79048 (1978); abstract on NASA Technical Report Server at http://ntrs.nasa.

gov, accessed Apr. 26, 2009.

149. Neil Otte, “PATRAN-STAGS Translator (PATSTAGS),” NASA TM-100388 (1990).

150. Southwest Research Institute, “NESSUS/NASTRAN Interface,” NASA CR-202778 (1996).
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Two of these are the YF-12 Thermal Loads project and the Rotor 

Aerodynamic Limits survey, both of which collected data for validating and improving predictive methods. The remaining three are the 

Highly Maneuverable Aircraft Technology (HiMAT) digital fly-by-wire 

(DFBW) enhanced agility composite-structured canard demonstrator, 

the AD-1 oblique wing demonstrator, and the Grumman X-29 forwardswept wing (FSW) research aircraft. These three projects exercised, in 

progressively more challenging ways, the concept of aeroelastic tailoring: that is, predicting airframe flexibility and having enough confidence 

in those predictions to design an airplane that takes advantage of elastic deformation, rather than just trying to minimize it. In all of these, 

NASA-rooted computational structural prediction proved of great, and 

even occasionally, critical, significance.

The investigation of aircraft structural mechanics or, indeed, of 

almost any discipline, can be considered to include the following activities: investigation by basic theory, computational analysis or simulation, laboratory test, and flight test (or, more generally, any test of the 

final product in its actual operating environment). Many arguments 

have been had over which is the most valuable. This author is of the 

opinion—based on his experience in the practice of engineering, on a 

certain amount of historical research, and on the teaching and example 

of mentors and peers—that theory, computation, laboratory test, and 

flight test all constitute imperfect but complementary views of reality. 

Thus, until someone comes up with a way to know the exact state of 

stress and deflection in every part of a vehicle under actual operating 

conditions, we must form our understanding of reality as a composite 

image, using what information we can gain from each available source:

• 

• 

Flight test, obviously, is the best representation we have 

of an aircraft in actual operational conditions. However, 

our ability to interrogate the system is most severely 

compromised in this activity. Many data parameters are 

not available unless special instrumentation is installed, 

if at all, and this is the most difficult environment in 

which to obtain stable, high-quality data.

Laboratory test offers better visibility into the operation of specific parts of the system and better control of 

experimental parameters, at the price of some separation from true operational conditions.
NASA’s Contributions to Aeronautics

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• 

• 

Computation offers even greater opportunity to interrogate the value of any data parameter at any time(s) 

and to simulate conditions that might be impossible, 

difficult, or dangerous to test. Computation also eliminates all physical complications of running the experiment and all physical sources of noise and uncertainty. 

But in stepping out of the physical world and into the 

analytical world, the researcher also becomes subject to 

the limited fidelity of his computational method: what 

effects are and are not included in the computation and 

how well the computation represents physical reality.

Theory is sometimes the best source of insight and 

of understanding what parameters might be changed 

to obtain some desired effect, but it does not provide 

the detailed quantitative data necessary to implement 

the solution.

In this light, the following flight programs are discussed. Much more 

could be said about each of them. The present discussion is necessarily 

confined to their significance to the development or validation of loads 

and structural computation methods.

YF-12 Thermal Loads and Structural Dynamics

NASA operated two Lockheed YF-12As and one “YF-12C” (actually an 

early nonstandard SR-71A, although the Air Force at that time could not 

acknowledge that it was allowing NASA to operate an SR-71) between 

1969 and 1979.151 These aircraft were used for a variety of research projects. In some projects, the YF-12s were the test articles, exploring their 

performance, handling qualities, and propulsion system characteristics 

in various baseline or modified configurations and modes of operation. 

In other projects, the YF-12s were used as “flying wind tunnels” to carry 

test models and other experiments into the Mach 3+ flight environment. 

Testing directly related to structural analysis methods and/or loads prediction included a series of thermal-structural load tests from 1969 to 

1972 and smaller projects concerning ventral fin loads and structural 

151. NASA DFRC Fact Sheet: YF-12A, at http://www.nasa.gov/centers/dryden/news/

FactSheets, accessed Aug. 13, 2009. For a detailed examination of this program, see a companion essay by William Flanagan.
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Temperature time histories from YF12 flight project. NASA.

dynamics.152 The flight-testing was conducted at Dryden, which was also 

responsible for project management. Ames, Langley, and Lewis Research 

Centers were all involved in technical planning, analysis, and supporting 

research activities, coordinated through NASA Headquarters. The U.S. 

Air Force and Lockheed also provided support in various areas.153 Gene 

Matranga of Dryden was the manager of the program before Berwin 

Kock later assumed that role.154

The thermal-structural loads project involved modeling and testing in Dryden’s unique thermal load facility. The purpose was to correlate in-flight and ground-test measurements and analytical predictions 

of temperatures, mechanical loads, strains, and deflections. “In all the 

X-15 work, flight conditions were always transient. The vehicle went 

to high speed in a matter of two to three minutes. It slowed down in a 

matter of three to five minutes. . . . The YF-12, on the other hand, could 

stay at Mach 3 for 15 minutes. We could get steady-state temperature 

152. Hallion and Gorn, On the Frontier, pp. 423–426.

153. James A. Albers, in James and Associates, eds., YF-12 Experiments Symposium, vol. 1, NASA 

CP-2054 (1978), p. 25.

154. Peter W. Merlin, Mach 3+: NASA/USAF YF-12 Flight Research, 1969–1979, NASA SP2001-4525, No. 25 in the Monographs in Aerospace History series (Washington, DC: NASA, 

2002), pp. 8, 95.
NASA’s Contributions to Aeronautics

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data, which would augment the X-15 data immeasurably.”155 The YF-12 

testing showed that it could take up to 15 minutes for absolute temperatures in the internal structure to approach steady state, and, even 

then, the gradients—which have a strong effect on stresses because of 

differential expansion—did not approach steady state until close to 30 

minutes into the cruise.156

NASTRAN and FLEXSTAB (a code developed by Boeing on contract 

to NASA Ames to predict aeroelastic effects on stability) were used to 

model the YF-12A’s aeroelastic and aerothermoelastic characteristics. 

Alan Carter and Perry Polentz of NASA oversaw the modeling effort, 

which was contracted to Lockheed and accomplished by Al Curtis. This 

effort produced what was claimed to be the most extensive full-vehicle 

NASTRAN model developed up to that time. The computational models 

were used to predict loads and deflections, and also to identify appropriate locations for the strain gauges that would take measurements in 

ground- and flight-testing. The instrumentation included strain gauges, 

thermocouples, and a camera mounted on the fuselage to record airframe deflection in flight. Most of the flights, from Flight 11 in April 1970 

through Flight 53 in February 1972, included data collection for this 

project, often mixed with other test objectives.157 Subsequently, the aircraft ceased flying for more than a year to undergo ground tests in the 

high-temperature loads laboratory. The temperatures measured in flight 

were matched on the ground, using heated “blankets” placed over different parts of the airframe. Ground-testing with no aerodynamic load 

allowed the thermal effects to be isolated from the aerodynamic effects.158

There were also projects involving the measurement of aerodynamic 

loads on the ventral fin and the excitation of structural dynamic modes. 

The ventral fin project was conducted to provide improved understanding of the aerodynamics of low aspect ratio surfaces. FLEXSTAB was 

used in this effort but only for linear aerodynamic predictions. Ground 

tests had shown the fin to be stiff enough to be treated as a rigid surface. 

Measured load data were compared to the linear theory predictions and 

155. Quoted in Merlin, Mach 3+, p. 90.

156. Jerald M. Jenkins and Albert E. Kuhl, “Recent Load Calibrations Experience with the YF 12 

Airplane,” in James and Associates, eds., YF-12 Experiments Symposium, vol. 1, NASA CP-2054 

(1978), p. 49.

157. Merlin, Mach 3+, p. 92; Hallion and Gorn, On the Frontier, pp. 188–189 and 423–424.

158. Merlin, Mach 3+, p. 90.
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to wind tunnel data.159 For the structural dynamics tests, which occurred 

near the end of NASA’s YF-12A program, “shaker vanes”—essentially 

oscillating canards—were installed to excite structural modes in flight. 

Six flights with shaker vanes between November 1978 and March 1979 

“provided flight data on aeroelastic response, allowed comparison with 

calculated response data, and thereby validated analytical techniques.”160

Experiences from the program were communicated to industry and 

other interested organizations in a YF-12 Experiments Symposium that 

was held at Dryden in 1978, near the end of the 10-year effort.161 There 

were also briefings to Boeing, specifically intended to provide information that would be useful on the Supersonic Transport (SST) program, 

which was canceled in 1971.162 There have been other civil supersonic 

projects since then—the High-Speed Civil Transport (HSCT)/High-Speed 

Research (HSR) efforts in the 1990s and some efforts related to supersonic business jets since 2000—but none have yet led to an operational 

civil supersonic aircraft.

Modern Rotor Aerodynamic Limits Survey

The Modern Rotor Aerodynamic Limits Survey was a 10-year program 

launched in 1984, which encompassed flight efforts in 1987 and 1993–

1994. In 1987, a Sikorsky UH-60A Black Hawk was tested with conventional structural instrumentation installed on the rotor blades. Then:

. . . Sikorsky Aircraft was [subsequently] contracted to build a 

set of highly instrumented blades for the Black Hawk test aircraft: a pressure blade with 242 absolute pressure transducers and a strain-gauge blade with an extensive suite of strain 

gauges and accelerometers . . . approximately 30 gigabytes of 

data were obtained in 1993–94 and installed in an electronic 

database that was immediately accessible to the domestic 

rotorcraft industry.163

159. Robert R. Meyer, Jr., and V. Michael DeAngelis, “Flight-Measured Aerodynamic Loads on a 

0.92 Aspect Ratio Lifting Surface,” in James and Associates, eds., YF-12 Experiments Symposium, 

vol. 1, NASA CP-2054 (1978), p. 77.

160. Merlin, Mach 3+, p. 42.

161. James and Associates, eds., YF-12 Experiments Symposium, vol. 1, NASA CP-2054 (1978).

162. Merlin, Mach 3+, pp. 93–94.

163. Borchers, et al., Flight Research at Ames, p. 71.
NASA’s Contributions to Aeronautics

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• 

• 

• 

• 

• 

• 

NASTRAN model and NASTRAN to static test comparison. NASA.

The two types of measurement systems are complementary. Strain 

gauges give an indication of the total load in a member, but little insight 

to the details of where and how the load is generated. The pressure taps 

show the distribution of the applied aerodynamic load, but only at given 

stations, so the total load estimate depends on how one computes the 

data through the unknown regions between the pressure transducers. The 

combination of both types of data is most useful to researchers trying to 

correlate computational loads predictions with the test data.

HiMAT

HiMAT was a small, unpiloted aircraft (23.5-feet long, 15.6-foot wingspan, 

weight just over 3,000 pounds) somewhat representative of a fighter type 

configuration, flown between 1979 and 1983, and developed to evaluate 

the following set of technologies and features:

Close-coupled canard.

Winglets.

Digital fly-by-wire flight control.

Composite structure.

Aeroelastic tailoring.

Supercritical airfoil.

It was intended that the benefits of these collected advances be shown 

together rather than separately and on an unpiloted platform, so that 
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HiMAT Electro-Optical Flight Deflection Measurement System. NASA.

the vehicle could be tested more aggressively without danger to 

a pilot.164

“Aeroelastic tailoring” refers to the design of a structure to achieve 

aerodynamically favorable deformation under load, rather than the more 

traditional approach of simply minimizing deformation. The goal of aeroelastic tailoring on the HiMAT “. . . was to achieve an aero-dynamically 

favorable spanwise twist distribution for maneuvering flight conditions” 

in the canard and the outboard wing. “The NASTRAN program was used 

to compute structural deflections at each model grid point. Verification 

of these deflections was accomplished by performing a loads test prior 

to delivery of the vehicle to NASA.” The ground-test loads were based on 

a sustained 8-g turn at Mach 0.9, which was one of the key performance 

design points of the aircraft. The NASTRAN model and a comparison 

between predicted and measured deflections are shown in the accompanying figure. Canard and wing twist were less than predicted. The difference was attributed to insufficient understanding of the matrix-dominated 

laminate material properties.165

The vehicle was also equipped with a system to measure deflections 

of the wing surface in flight. Light emitting diodes (LEDs)—referred to 

as targets—on the wing upper surface were detected by a photodiode 

164. Hallion and Gorn, On The Frontier, pp. 276–278.

165. Richard C. Monaghan, “Description of the HiMAT Tailored Composite Structure and Laboratory Measured Vehicle Shape Under Load,” NASA TM-81354 (1981), pp. 4–5, 7.
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array mounted on the fuselage, at a location overlooking the wing. Three 

inboard targets were used to determine a reference plane, from which 

the deflection of the remaining targets could be measured. To measure 

wing twist, targets were positioned primarily in pairs along the front 

and rear wing spars.166 The HiMAT wing had a relatively small number of targets—only two pairs besides the inboard reference set—so the 

in-flight measurements were not a detailed survey of the wing by any 

means. Rather, they provided measurement at a few key points, which 

could then be compared with the NASTRAN data and the ground loads 

test data. Target and receiver locations are illustrated here, together with 

a sample of the deflection data at the 8-g maneuver condition. In-flight 

deflection data showed similar twist to the ground-test data, indicating 

that the aerodynamic loads were well predicted.167

The HiMAT was an early step in the development of aeroelastic tailoring capability, providing a set of NASTRAN data, static load test data, 

and flight-test data, for surface deflection at a given loading condition. The 

project also proved out the electro-optical system for in-flight deflection 

measurements, which would later be used in the X-29 project.

AD-1 Oblique Wing Demonstrator

The AD-1 was a small and inexpensive demonstrator aircraft intended to 

investigate some of the issues of an oblique wing. It flew between 1979 

and 1982. It had a maximum takeoff weight of 2,100 pounds and a maximum speed of 175 knots. It is an interesting case because (1) NASA had 

an unusually large role in its design and integration—it was essentially 

a NASA aircraft—and (2) because it provides a neat illustration of the 

prosecution of a particular objective through design, analysis, wind tunnel test, flight test, and planned follow-on development.168

The oblique wing was conceived by German aerodynamicists in the 

midst of the Second World War. But it was only afterward, through the 

166. Walter J. Sefic and Karl F. Anderson, “NASA High Temperature Loads Calibration Laboratory,” 

NASA TM-X-1868 (1969), pp. 3–4.

167. Glenn B. Gilyard, “The Oblique Wing Research Aircraft: A Test Bed for Unsteady Aerodynamic 

and Aeroelastic Research,” in NASA Langley Research Center, Transonic Unsteady Aerodynamics and 

Aeroelasticity, pt. 2, Report N89-19247 (1987), p. 412; J.J. Burken, G.S. Alag, and G.B. Gilyard, 

“Aeroelastic Control of Oblique-Wing Aircraft,” NASA TM-86808 (1986).

168. Robert E. Curry and Alex G. Sim, “In-Flight Total Forces, Moments, and Static Aeroelastic 

Characteristics of an Oblique-Wing Research Airplane,” NASA TP-2224 (1984), p. 15.
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AD-1 three view. NASA.

brilliance and determination of NASA aerodynamicist Robert T. Jones that 

it advanced to actual flight. Indeed, Jones, father of the American swept 

wing, became one of the most persistent proponents of the oblique wing 

concept.169 The principal advantage of the oblique wing is that it spreads 

both the lift and volume distributions of the wing over a greater length 

than that of a simple symmetrically swept wing. This has the effect of 

reducing both the wave drag because of lift and the wave drag because 

of volume, two important components of supersonic drag. With this theoretical advantage come practical challenges. The challenges fall into 

two broad categories: the effects of asymmetry on the flight characteristics (stability and handling qualities) of the vehicle, and the aeroelastic stability of the forward-swept wing. The research objectives of the 

AD-1 were primarily oriented toward flying qualities. The AD-1 was not 

intended to explore structural dynamics or divergence in depth, other 

169. Jones continued to study it and advocate it until his death in 1999; the author had the pleasure 

of hearing him lecture on this topic in the aerodynamics class that Jones taught at Stanford University in 

the 1980s. For Jones, see statement of William Sears in Ames Research Center staff, “Collected Works 

of Robert T. Jones,” NASA TM-X-3334 (1976), pp. vii-ix; Hallion, “Lippisch Gluhareff, and Jones: The 

Emergence of the Delta Planform and the Origins of the Sweptwing in the United States,” Aerospace 

Historian, vol. 26, No. 1 (Mar. 1979), pp. 1–10; and Walter G. Vincenti, “Robert Thomas Jones,” in 

Biographical Memoirs, vol. 86 (Washington, DC: National Academy of Sciences, 2005), pp. 3–21.
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than establishing safety of flight. Mike Rutkowski analyzed the wing for 

flutter and divergence using NASTRAN and other methods.170

However, the project did make a significant accomplishment in the 

use of static aeroelastic tailoring. The fiberglass wing design by Ron Smith 

was tailored to bend just enough, with increasing g, to cancel out an aerodynamically induced rolling moment. Pure bending of the oblique wing 

increases the incidence (and therefore the lift) of the forward-swept tip 

and decreases the incidence (and lift) of the aft-swept tip. In a pullup 

maneuver, increasing lift coefficient (CL), and load factor at a given flight 

condition, this would cause a rollaway from the forward-swept tip. At the 

same time, induced aerodynamic effects (the downwash/upwash distribution) increase the lift at the tip of an aft-swept wing. On an aircraft with 

only one aft-swept tip, this would cause a roll toward the forward-swept 

side. The design intent for the AD-1 was to have these two effects cancel 

each other as nearly as possible, so that the net change in rolling moment 

because of increasing g at a given flight condition would be zero. The 

design condition was CL = 0.3 for 1-g flight at 170 knots, 12,500-foot altitude, and a weight of 1,850 pounds, with the wing at 60-degree sweep.171

An aeroelastically scaled one-sixth model was tested at full-scale 

Reynolds number in the Ames 12-Foot Pressure Wind Tunnel. A stiff aluminum wing was used for preliminary tests, then two fiberglass wings. 

The two fiberglass wings had zero sweep at the 25- and 30-percent chord 

lines, respectively, bracketing the full-scale AD-1 wing, which had zero 

sweep at 27-percent chord. The wings were tested at the design scaled 

dynamic pressure and at two lower values to obtain independent variation of wing load because of angle of attack and dynamic pressure at a 

constant angle of attack. Forces and moments were measured, and deflection was determined from photographs of the wing at test conditions.172

Subsequently, “. . . the actual wing deflection in bending and twist 

was verified before flight through static ground loading tests.” Finally, 

in-flight measurements were made of total force and moment coefficients and of aeroelastic effects. Level-flight decelerations provided 

angle-of-attack sweeps at constant load, and windup turns provided 

angle-of-attack sweeps at constant “q” (dynamic pressure). Results 

170. Steve Smith, interview with author, Apr. 6, 2009; and M.J. Rutkowski, “Aeroelastic stability 

analysis of the AD-1 Manned Oblique-Wing Aircraft,” NASA TM-78439 (1977).

171. Curry and Sim, “In-Flight Total Forces,” p. 3.

172. Ibid., pp. 3–4.
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were interpreted and compared with predictions. The simulator model, 

with aeroelastic effects included, realistically represented the dynamic 

responses of the flight vehicle.173

Provision had been made for mechanical linkage between the pitch 

and roll controls, to compensate for any pitch-roll coupling observed in 

flight. However, the intent of the aeroelastic wing was achieved closely 

enough that the mechanical interconnect was never used.174 Roll trim 

was not needed at the design condition (60-degrees sweep) nor at zero 

sweep, where the aircraft was symmetric. At intermediate sweep angles, 

roll trim was required. The correction of this characteristic was not pursued because it was not a central objective of the project. Also, the airplane experienced fairly large changes in rolling moment with angle 

of attack beyond the linear range. Vortex lift, other local flow separations, and ultimately full stall of the aft-swept wing, occurred in rapid 

succession as angle of attack was increased from 8 to approximately 12 

degrees. Therefore, it would be a severe oversimplification to say that 

the AD-1 had normal handling qualities.175

The AD-1 flew at speeds of 170 knots or less. On a large, high-speed 

aircraft, divergence of the forward-swept wing would also be a consideration. This would be addressed by a combination of inherent stiffness, 

aeroelastic tailoring to introduce a favorable bend-twist coupling, and, 

potentially, active load alleviation. The AD-1 project provided initial correlation of measured versus predicted wing bending and its effects on 

the vehicle’s flight characteristics. NASA planned to take the next step 

with a supersonic oblique wing aircraft, using the same F-8 airframe 

that had been used for earlier supercritical wing tests. These studies 

delved deeper into the aeroelastic issues: “Preliminary studies have 

been performed to identify critical DOF [Degree of Freedom] for flutter model tests of oblique configurations. An ‘oblique’ mode has been 

identified with a 5 DOF model which still retains its characteristics with 

the three rotational DOF’s. An interdisciplinary analysis code (STARS), 

which is capable of performing flutter and aeroservoelastic analyses, has 

been developed. The structures module has a large library of elements 

and in conjunction with numerical analysis routines, is capable of efficiently performing statics, vibration, buckling, and dynamic response 

173. Ibid., pp. 3–4, 10.

174. Smith, interview with author.

175. Curry and Sim, “In-Flight Total Forces,” pp. 7–8.
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analysis of structures. . . . ” The STARS code also included supersonic 

(potential gradient method) and subsonic (doublet lattice) unsteady aerodynamics calculations. “ . . . Linear flutter models are developed and 

transformed to the body axis coordinate system and are subsequently 

augmented with the control law. Stability analysis is performed using 

hybrid techniques. The major research benefit of the OWRA [Oblique 

Wing Research Aircraft] program will be validation of design and analysis tools. As such, the structural model will be validated and updated 

based on ground vibration test (GVT) results. The unsteady aero codes 

will be correlated with experimentally measured unsteady pressures.”176

While the OWRA program never reached flight, (NASA was ready 

to begin wing fabrication in 1987, expecting first flight in 1991), these 

comments illustrate the typical interaction of flight programs with analytical methods development and the progressive validation process that 

takes place. Such methods development is often driven by unconventional problems (such as the oblique wing example here) and eventually 

finds its way into routine practice in more conventional applications. 

For example, in the design of large passenger aircraft today, the loads 

process is typically iterated to include the effects of static aeroelastic 

deflections on the aerodynamic load distribution.177

X-29

The Grumman X-29 aircraft was an extraordinarily ambitious and productive flight-test program run between 1984 and 1992. It demonstrated a 

large (approximately 35 percent) unstable static margin in the pitch axis, 

a digital active flight control system utilizing three-surface pitch control 

(all-moving canards, wing flaps, and aft-mounted strake flaps), and a thin 

supercritical forward-swept wing, aeroelastically tailored to prevent structural divergence. The X-29 was funded by the Defense Advanced Research 

Projects Agency (DARPA) through the USAF Aeronautical Systems Division 

(ASD). Grumman was responsible for aircraft design and fabrication, 

including the primary structural analyses, although there was extensive involvement of NASA and the USAF in addressing the entire realm 

of unique technical issues on the project. NASA Ames Research Center/

Dryden Flight Research Facility was the responsible test organization.178

176. Gilyard, “The Oblique Wing Research Aircraft.”

177. Mendoza and Hazen, interview with author.

178. Sefic and Maxwell, “X-29A Overview,” p. 2.
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Careful treatment of aeroelastic stability was necessary for the thin 

FSW to be used on a supersonic, highly maneuverable aircraft. According 

to Grumman, “Automated design and analysis procedures played 

a major role in the development of the X-29 demonstrator aircraft.” 

Grumman used one of its programs, called FASTOP, to optimize the 

X-29’s structure to avoid aeroelastic divergence while minimizing the 

weight impact.179

In contrast to the AD-1, which allowed the forward-swept wing to 

bend along its axis, thereby increasing the lift at the forward tip, the 

X-29’s forward-swept wings were designed to twist when bending, in a 

manner that relieved the load. This was accomplished by orienting the 

primary spanwise fibers in the composite skins at a forward “kick angle” 

relative to the nominal structural axis of the wing. The optimum angle 

was found in a 1977 Grumman feasibility study: “Both beam and coarsegrid, finite-element models were employed to study various materials 

and laminate configurations with regard to their effect on divergence and 

flutter characteristics and to identify the weight increments required to 

avoid divergence.”180 While a pure strength design was optimum at zero 

kick angle, an angle of approximately 10 degrees was found to be best 

for optimum combined strength and divergence requirements.

When the program reached the flight-test phase, hardware-in-theloop simulation was integral to the flight program. During the functional and envelope expansion phases, every mission was flown on 

the simulator before it was flown in the airplane.181 In flight, the X-29 

No. 1 aircraft (of two that were built) carried extensive and somewhat 

unique instrumentation to measure the loads and deflections of the airframe, and particularly of the wing. This consisted of pressure taps on 

the left wing and canard, an optical deflection measurement system on 

the right wing, strain gages for static structural load measurement, and 

accelerometers for structural dynamic and buffet measurement.

The most unusual element of this suite was the optical system, which 

had been developed and used previously on the HiMAT demonstrator 

(see preceding description). Optical deflection data were sampled at a 

179. Philip Mason, Edward Lerner, and Lawrence Sobel, “Applications of Integrated Design/Analysis Systems in Aerospace Structural Design,” in NASA Langley Research Center, Recent Advances in 

Multidisciplinary Analysis and Optimization, pt. 1 (1989), p. 25.

180. Ibid., p. 26.

181. Sefic and Maxwell, “X-29A Overview,” p. 3.
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Loads measurement provisions on Grumman X-29A. NASA.

rate of 13 samples per channel per second. Data quality was reported 

to be very good, and initial results showed good match to predictions. 

In addition, pressure data from the 156 wing and 17 canard pressure 

taps was collected at a rate of 25 samples per channel per second. One 

hundred six strain gages provided static loads measurement as shown. 

Structural dynamic data from the 21 accelerometers was measured at 

400 samples per channel per second. All data was transmitted to ground 

station and, during limited-envelope phase, to Grumman in Calverton, 

NY, for analysis.182 “Careful analyses of the instrumentation requirements, 

flight test points, and maneuvers are conducted to ensure that data of 

sufficient quality and quantity are acquired to validate the design, fabrication, and test process.”183 The detailed analysis and measurements 

provided extensive opportunities to validate predictive methods.

The X-29 was used as a test case for NASA’s STARS structural analysis computer program, which had been upgraded with aeroservoelastic analysis capability. In spite of the exhaustive analysis done ahead 

of time, there were, as is often the case, several “discoveries” made 

during flight test. Handling qualities at high alpha were considerably 

better than predicted, leading to an expanded high-alpha control and 

maneuverability investigation in the later phases of the project. The X-29 

182. Ibid., pp. 3–4.

183. Ibid., p. 4.
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No. 1 was initially limited to 21-degree angle of attack, but, during subsequent Phase II envelope expansion testing, its test pilots concluded it 

had “excellent control response to 45 deg. angle of attack and still had 

limited controllability at 67 deg. angle of attack.”184

There were also at least two distinct types of aeroservoelastic phenomena encountered: buffet-induced modes and a coupling between 

the canard position feedback and the aircraft’s longitudinal aerodynamic and structural modes were observed.185 The modes mentioned 

involved frequencies between 11 and 27 hertz (Hz). Any aircraft with 

an automatic control system may experience interactions between the 

aircraft’s structural and aerodynamic modes and the control system. 

Typically, the aeroelastic frequencies are much higher than the characteristic frequencies of the motion of the aircraft as a whole. However, 

the 35-percent negative static margin of the X-29A was much larger 

than any unstable margin designed into an aircraft before or since. As 

a consequence, its divergence timescale was much more rapid, making 

it particularly challenging to tune the flight control system to respond 

quickly enough to aircraft motions, without being excited by structural 

dynamic modes. Simply stated, the X-29A provided ample opportunity 

for aeroservoelastic phenomena to occur, and such were indeed observed, 

a contribution of the aircraft that went far beyond simply demonstrating the aerodynamic and maneuver qualities of an unstable forwardswept canard planform.186

In sum, each of these five advanced flight projects provides important lessons learned across many disciplines, particularly the validation 

of computer methods in structural design and/or analysis. The YF-12 

project provided important correlation of analysis, ground-test data, 

and flight data for an aircraft under complex aerothermodynamic loading. The Rotor Aerodynamic Limits survey collected important data on 

helicopter rotors—a class of system often taken for granted yet one that 

184. Quoted in Dryden Flight Research Center “Fact Sheets: X-29,” http://www.nasa.gov/

centers/dryden/news/FactSheets, accessed Mar. 19, 2009.

185. David F. Voracek and Robert Clarke, “Buffet Induced Structural/Flight Control System Interaction of the X 29A Aircraft,” AIAA Paper 91-1053 (1991); and Michael W. Kehoe, Lisa J. Bjarke, 

and Edward J. Laurie, “An In-Flight Interaction of the X 29A Canard and Flight Control System,” 

NASA TM-101718 (1990).

186. K.K. Gupta, M.J. Brenner, and L.S. Volker, “Integrated Aeroservoelastic Analysis Capability 

With X 29A Analytical Comparisons,” AIAA Paper 87-0907 (1987).
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represent an incredibly complex interaction of aerodynamic, aeroelastic, and inertial phenomena. The HiMAT, AD-1, and X-29 programs each 

advanced the state of the art in aeroelastic design as applied to nontraditional, indeed exotic, planforms featuring unstable design, composite structures, and advanced flight control concepts. Finally, the data 

required to validate structural analysis and design methods do not automatically come from the testing normally performed by aircraft developers and users. Special instrumentation and testing techniques are 

required. NASA has developed the facilities and the knowledge base 

needed for many kinds of special testing and is able assign the required 

priority to such testing. As these cases show, NASA therefore plays a key 

role in this process of gaining knowledge about the behavior of aircraft 

in flight, evaluating predictive capabilities, and flowing that experience 

back to the people who design the aircraft.

Computational Methods, Industrial Transfer, and the Way Ahead

Having surveyed the development of computational structural analysis 

within NASA, the contributions of various Centers, and key flight projects that tested and validated structural design and analysis methods 

in their ultimate application, we turn to the current state of affairs as 

of 2010 and future challenges.

Overall, even a cursory historical examination clearly indicates that 

the last four decades have witnessed revolutionary improvements in all 

of the following areas:

• 

• 

• 

• 

• 

• 

• 

• 

• 

• 

• 

• 

• 

• 

Analysis capability.

Complexity of structures that can be analyzed.

Number of nodes.

Types of elements.

Complexity of processes simulated.

Nonlinearity.

Buckling.

Other geometric nonlinearity.

Material nonlinearity.

Time-dependent properties.

Yield or ultimate failure of some members.

Statistical/nondeterministic processes.

Thermal effects.

Control system interactions.
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• 

• 

• 

• 

• 

• 

• 

• 

Usability.

Execution time.

Hardware improvements.

Efficiency of algorithms.

Adequate but not excessive model complexity.

Robustness, diagnostics, and restart capability.

Computing environment.

Pre- and post-processing.

Before NASTRAN, capabilities generally available (i.e., not counting proprietary programs at the large aerospace companies) were limited to a few hundred nodes. In 1970, NASTRAN made it possible to 

analyze models with over 2,000 nodes. Currently, models with hundreds 

of thousands of nodes are routinely analyzed. The computing environment has changed just as dramatically, or more so: the computer used 

to be a shared resource among many users—sometimes an entire company, or it was located at a data center used by many companies—with 

punch cards for input and reams of paper for output. Now, there is a PC 

(or two) at every engineer’s desk. NASTRAN can run on a PC, although 

some users prefer to run it on UNIX machines or other platforms.

Technology has thus come full circle: NASA now makes extensive use 

of commercial structural analysis codes that have their roots in NASA 

technology. Commercial versions of NASTRAN have essentially superseded NASA’s COSMIC NASTRAN. That is appropriate, in this author’s 

opinion, because it is not NASA’s role to provide commercially competitive performance, user interfaces, etc. The existence and widespread use 

of these commercial codes indicates successful technology transition.

At the time of this writing, basic capability is relatively mature. 

Advances are still being made, but it is now possible to analyze the vast 

majority of macroscopic structural problems that are of practical interest in aeronautics and many other industries.

Improvements in the “usability” category are of greater interest to 

most engineers. Execution speed has improved orders of magnitude, but 

this has been partially offset by the corresponding orders-of-magnitude 

increases in model size. Engineers build models with hundreds of 

thousands of nodes, because they can.

Pre- and post-processing challenges remain. Building the model and 

interpreting the results typically take longer than actually running the 

analysis. It is by no means a trivial task to build a finite element model 
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of a complex structure such as a complete airframe, or a major portion 

thereof. Some commercial software can generate finite element models automatically from CAD geometry. However, many practitioners 

in the aircraft industry prefer to have more involvement in the modeling process, because of the complexity of the analysis and the safetycritical nature of the task. The fundamental challenge is to make the 

modeling job easier, while providing the user with control when required 

and the ability to thoroughly check the resulting model.187

In 1982, Thomas Butler wrote, “I would compare the state of graphics pre- and post-processors today with the state that finite elements 

were in before NASTRAN came on the scene in 1964. Many good features exist. There is much to be desired in each available package.”188

Industry practitioners interviewed today have expressed similar sentiments. There is no single pre- or post-processing product that meets 

every need. Some users deliberately switch between different pre- and 

post-processing programs, utilizing the strengths of each for different 

phases of the modeling task (such as creating components, manipulating them, and visualizing and interrogating the finished model). A reasonable number of distinct pre- and post-processing systems maintain 

commercial competition, which many users consider to be important.189

As basic analysis capability has become well established, researchers step back and look at the bigger picture. Integration, optimization, 

and uncertainty modeling are common themes at many of the NASA 

Centers. This includes integration of design and analysis, of analysis 

and testing, and of structural analysis with analysis in other disciplines. 

NASA Glenn Research Center is heavily involved in nondeterministic 

analysis methods, life prediction, modeling of failure mechanisms, and 

modeling of composite materials, including high-temperature material 

systems for propulsion applications. Research at Langley spans many 

fields, including multidisciplinary analysis and optimization of aircraft 

and spacecraft, analysis and test correlation, uncertainty modeling and 

“fuzzy structures,” and failure analysis.

In many projects, finite element analysis is being applied at the 

microscale to gain a better understanding of material behaviors. The 

187. Thomas Christy and John Splichal (aircraft industry), interview by author, Wichita, KS, May 

12, 2009; Jadic interview.

188. Butler, “Operating in the Age of NASTRAN,” p. xx.

189. Mendoza, Hazen, Jadic, and Christy-Splichal interviews by author.
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ability to perform such analysis is a noteworthy benefit coming from 

advances in structural analysis methods at the macroscopic level. Very 

real benefits to industry could result. The weight savings predicted from 

composite materials have been slow in coming, partly because of limitations on allowable stresses. In the civil aviation industry especially, 

such limitations are not necessarily based on the inherent characteristics of the material but on the limited knowledge of those characteristics. Analysis that gives insight into material behaviors near failure, 

documented and backed up by test results, may help to achieve the full 

potential of composite materials in airframe structures.

Applications of true optimization—such as rigorously finding the 

mathematical minimum of a “cost function”—are still relatively limited in the aircraft industry. The necessary computational tools exist. 

However, the combination of practical difficulties in automating complex analyses and a certain amount of cultural resistance has somewhat 

limited the application of true optimization in the aircraft industry up 

to the present time. There is untapped potential in this area. The path to 

reaching it is not necessarily in the development of better computer programs, but rather, in the development and demonstration of processes 

for the effective and practical use of capabilities that exist already. The 

DAMVIBS program (discussed previously in the section on the NASA 

Langley Research Center) might provide a model for how this kind 

of technology transfer can happen. In that program, industry teams 

essentially demonstrated to themselves that existing finite element 

programs could be useful in predicting and improving the vibration 

characteristics of helicopters—when coupled with some necessary 

improvements in modeling technique. All of the participants subsequently embraced the use of such methods in the design processes of 

their respective organizations. A comparable program could, perhaps, 

be envisioned in the field of structural and/or multidisciplinary optimization in aircraft design.190

Considering structural analysis as a stand-alone discipline, however, it can be stated without question that computational methods 

have been adopted throughout the aircraft industry. Specific processes 

vary between companies. Some companies perform more upfront 

optimization than others; some still test exhaustively, while others test 

190. Jadic interview; and author’s experience.
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minimally. But the aircraft industry as a whole has embraced computational structural analysis and benefited greatly from it.

The benefits of computational structural analysis may not be adequately captured in one concise list, but they include the following:

• 

• 

• 

• 

• 

• 

• 

• 

• 

Improved productivity of analysis.

Ability to analyze a more complete range of load cases.

Ability to analyze a structure more thoroughly than was 

previously practical.

Ability to correct and update analyses as designs and 

requirements mature.

Improved quality and consistency of analysis.

Improved performance of the end product. Designs can 

be improved through more cycles of design/analysis in 

the early stages of a project, and earlier identification of 

structural issues, than previously practical.

Improved capabilities in related disciplines: thermal 

modeling and acoustic modeling, for example. Some 

aircraft companies utilize finite element models in the 

design stage of an aircraft to develop effective noise 

reduction strategies.

Ability to analyze structures that could not be practically 

analyzed before. For example, composite and metallic 

airframes are different. Metal structures typically have 

more discrete load paths. Composite structures, such 

as honeycomb-core panels, have less distinct load paths 

and are less amenable to analysis by hand using classical methods. Therefore, finite element analysis enables 

airplanes to be built in ways that would not be possible 

(or, at least, not verifiable) otherwise.

Reduced cost and increased utility of testing. Analysis 

does not replace all testing, but it can greatly enhance 

the amount of knowledge gained from a test. For example, modeling performed ahead of a test series can help 

identify the appropriate locations for strain gages, accelerometers, and other instrumentation, and aid in the 

interpretation of the resulting test data. The most difficult or costly types of testing can certainly be reduced. 

In a greatly simplified sense, the old paradigm is that 
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testing was the proof of the structure; now, testing validates the model, and the model proves the structure. 

Practically speaking, most aircraft companies practice 

something in between these two extremes.

NASA’s contributions have included not only the development of 

the tools but also the development and dissemination of techniques to 

apply the tools to practical problems and the provision of opportunities—through unique test facilities and, ultimately, flight research projects—to prove, validate, and improve the tools.

In other industries also, there is now widespread use of computerized structural analysis for almost every conceivable kind of part that 

must operate under conditions of high mechanical and/or thermal stress. 

NASTRAN is used to analyze buildings, bridges, towers, ships, wind tunnels and other specialized test facilities, nuclear power plants, steam 

turbines, wind turbines, chemical processing plants, microelectronics, 

robotic systems, tools, sports equipment, cars, trucks, buses, trains, 

engines, transmissions, and tires. It is used for geophysical and seismic 

analysis, and for medical applications.

In conclusion, finite element analysis would have developed with or 

without NASA’s involvement. However, by creating NASTRAN, NASA 

provided a centerpiece: a point of reference for all other development 

and an open-ended framework into which new capabilities could be 

inserted. This framework gradually collected the best or nearly best 

methods in every area. If NASTRAN had not been developed, different advances would have occurred only within proprietary codes used 

internally by different industrial companies or marketed by different software companies. There would have been little hope of consolidating all the important capabilities into one code or of making such 

capabilities available to the general user. NASTRAN brought highpowered finite element analysis within reach of many users much sooner 

than would have otherwise been the case. At the same time, the job of 

predicting every aspect of structural performance was by no means finished with the initial release of NASTRAN—nor is it finished yet. NASA 

has been and continues to be involved in the development of many 

new capabilities—developing programs and new ways to apply existing 

programs—and making the resulting tools and methods available to 

users in the aerospace industry and in many other sectors of the 

U.S. economy.
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Appendix A: 

NASTRAN Users’ Colloquia, 1971–1993

Note: This appendix includes a list of the dates and locations of the 

NASTRAN Users’ Colloquia and NASTRAN applications presented at 

the Colloquia by “nontraditional” users, i.e., industry other than aerospace, Government agencies other than NASA, and universities. Not every 

paper from these sources is listed, only those that represent applications. 

Many other papers were presented on modeling techniques, capability 

improvements, etc., which are not listed.

NASTRAN USERS’ COLLOQUIA DATES AND LOCATIONS

# YEAR DATE LOCATION CHAIRPERSON(S) 

 / OTHER NOTES

1st 1971 Sept. 13–15 NASA Langley J. Philip Raney 

(NASTSRAN SMO)

2nd 1972 Sept. 11–12 NASA Langley J. Philip Raney

3rd 1973 Sept. 11–12 NASA Langley <not available>

4th 1975 Sept. 9–11 NASA Langley Deene J. Weidman

5th 1976 Oct. 5–6 NASA Ames Deene J. Weidman

6th 1977 Oct. 4–6 NASA Lewis Deene J. Weidman (Langley)

and Christos Chamis (Lewis)

7th 1978 Oct. 4–6 NASA Marshall Deene J. Weidman 

Robert L. McComas 

(Marshall)

(Langley)

8th 1979 Oct. 30–31 NASA Goddard Robert L. 

Reginal

Brugh (COSMIC)

9th 1980 Oct. 22–23 NASA Kennedy Robert 

Henry 

L. Brugh (COSMIC)

Harris (KSC)

Note: From this point on, locations were no longer at NASA Centers, individual co/chairs are 

not identified in the proceedings, and the NASA Scientific & Technical Information (STI) Branch 

(or program) is listed in the proceedings as the responsible organization.

10th 1982 May 13–14 New Orleans, LA Co-chairs not identified.
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11th 1983 May 2–6 San Francisco, CA

12th 1984 May 7–11 Orlando, FL

13th 1985 May 6–10 Boston, MA

14th 1986 May 5–9 San Diego, CA

15th 1987 May 4–8 Kansas City, MO

16th 1988 Apr. 25–29 Arlington, VA

17th 1989 Apr. 24–28 San Antonio, TX

18th 1990 Apr. 23–27 Portland, OR COSMIC, 

Branch.

under the STI 

19th 1991 Apr. 22–26 Williamsburg, VA

20th 1992 Apr. 27–May 1 Colorado 

CO

Springs, 

21st 1993 Apr. 26–30 Tampa, FL
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NONAEROSPACE INDUSTRY APPLICATIONS 

PRESENTED AT USERS’ COLLOQUIA

OF NASTRAN 

YEAR COMPANY DESCRIPTION

1972

Westenhoff and Novick Analysis and design 

track support.

of on-grade railroad 

General Motors NASTRAN and 

tive structures.

in-house code for automoWestinghouse (Hanford) Fuel handling machinery for reactors.

Kleber-Colombes Tires.

Control Data Corp (CDC) Structural analysis of 40-story building.

Computer 

(CSC)

Sciences Corporation Structural 

nuclear re

dynamic and thermal analysis 

actor vessel support system.

1975

B.F. Goodrich Tires.

Exxon Petroleum processing machinery.

Littleton Rsch & Eng, with CDC Propeller-induced ship vibration.

Westinghouse (Hanford) Seismic analysis 

structures.

of nuclear reactor 

Reactor Centrum 

Hazameyer B.V.

Nederland & Electromagnetic field problems.

General Motors Modeling and analysis of acoustic cavities.

1976

Sargent & 

(2 papers, 

Lundy 

1 with CSC)

Deformations of thick cylinders (power 

plants); seismic analysis of nuclear power 

plant control panel.

EBASCO 

Analytics

Services, with Universal Concrete cracking.

1977 Sperry Marine with Univ VA Analysis of pressure vessels.

1978

Tennessee Eastman Co. NASTRAN uses in petrochemical industry.

EBASCO Services with Grumman (2) Tokomak 

coil and 

Fusion Test Reactor toroidal 

vacuum vessel structures.

field 

B.F. Goodrich Rubber sonar dome window.

1979 B.F. Goodrich Belt tensioning.

1980

Ontario Hydro Seismic analysis.

NKF Engineering Problems 

motion.

involving enforced boundary 

Tennessee Eastman Analysis 

failures.

of heat-transfer fluid fill pipe 

1982 B.F. Goodrich Bead area 

interface.

contact load at tire-wheel 
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1984

Tennessee Eastman Support system for large compressor.

Hughes Offshore Bolted marine riser structure.

1985 John Deere Use of COSMIC 

department.

NASTRAN in design 

1986 Texas Instruments Nonlinear magnetic circuits.

1987

Texas Instruments Forces on magnetized bodies.

NKF Eng. HVAC duct hanger systems.

1988

Tiernay Turbines Stress and vibration 

components.

analysis of gas turbine 

Texas Instruments Magnetostatic 

printhead.

nonlinear model of 

1989

Deutsch Metal Components General product line improvement 

(hydraulics, pneumatics, other power 

system components).

Intergraph NASTRAN in integrated 

design environ.

conceptual 

Dynacs Eng. Flexible multibody dynamics and 

(NASTRAN with TREETOPS).

control 

Texas Instruments Micromechanical deformable mirror.

1990 Analex Corp., with NASA Lewis Low velocity impact analysis.

1991 Tennessee Eastman Distillation tray structures.

1993 Butler Analyses Seismic analysis.
NASA’s Contributions to Aeronautics

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8

OTHER GOVERNMENT AGENCY NASTRAN APPLICATIONS 

PRESENTED AT USERS’ COLLOQUIA

YEAR GOVERNMENT 

AGENCY

DESCRIPTION

Naval Air Dev Ctr F-14A boron horizontal stabilizer static and 

1971

dynamic.

U.S. 

R&D 

with 

Army Air Mobility 

Lab (USAAMRDL) 

NASA Langley

NASTRAN in structural design optimization.

1975

Naval Weapons Center Modeling and analysis of damaged wings.

Naval 

tems C

Underwater Sysenter (NUSC)Transient analysis 

boundaries.

of bodies with moving 

(David 

Rsch & 

SRDC)

Taylor) Naval Ship 

Dev Ctr (DTNDynamic analysis of submerged structures.

Argonne Nat Lab Fluid-coupled 

reactors).

concentric cylinders (nuclear 

1976

DTNSRDC Underwater shock response.

NUSC Fluid-structure interactions.

DTNSRDC Submerged structures.

USAAMRDL 

Vertol

with Boeing Thermal and 

transmission.

structural analysis of helicopter 

U.S. Army, Watervliet Crack problems.

1977

DTNSRDC Finite element solutions for free surface flows.

NUSC Analysis of magnetic fields.

U.S. Army, Watervliet Large-deformation 

shells.

analysis of fiber-wrapped 

1978 Wright-Patterson AFB Ceramic structures.

DTNSRDC Magnetostatic field problems.

1979

U.S. Army Armament 

Rsch & Dev Command 

(USAARDC) (2)

Stress concentrations 

plastic analysis.

in screw heads, elasticNUSC Dynamically loaded periodic structures.

1980

NUSC, with A.O. Smith Ring element dynamic stresses.

USAARDC (2) Simulated damage 

plastic analysis.

UH-1B tailboom, elastic-
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1982

DTNSRDC Magnetic field problems.

NUSC Axisymmetric 

problems.

fluid structure interaction 

USAARDC Analysis of overloaded breech ring.

1983

DTNSRDC Fluid-filled elastic piping systems.

NUSC (2) Wave propagation through plates (2).

U.S. Army Benet Lab Elastic-plastic analysis of annular plates.

1984 Dept. of the Navy Acoustic scattering from submerged structures.

1985

WPAFB with Rockwell NASTRAN in a computer aided design system.

Naval Wpns Ctr Missile inertia loads.

WPAFB Simulation of nuclear overpressures.

DTNSRDC Loss factors, 

treatment.

frequency-dependent damping 

U.S. Army (Harry Diamond Lab) with Advanced 

Tech & Rsch

Transient analysis of fuze assembly.

DTNSRDC Magnetic heat pump.

1986

DTNSRDC (3) Multidisciplinary design; acoustics (2).

Naval Ocean Sys Ctr (2) Stress concentrations; 

wings.

flutter of low aspect ratio 

NUSC Surface impedance analysis.

1987

DTNSRDC (2) Computer animation of modal and transient 

vibrations; analysis of ship structures to underwater explosion shocks.

DTNSRDC & NRL Acoustic scattering.

NUSC Patrol boat subject to planning loads.

1988 David Taylor Rsch 

(DTRC—renamed)

Ctr Static preload 

scattering.

effects in acoustic radiation and 

1989 DTRC Low frequency 

structures.

resonances of submerged 

1990 DTRC Scattering from fluid-filled structures.

1991

Los Alamos Nat Lab Computer animation of displacements.

DTRC (2) Transient fluid-structure interactions.

1992

Naval Surf. Warfare Ctr Vibration 

plates.

and shock of laminated composite 

DTRC Acoustics of axisymmetric fluid regions.

1993 U.S.A.F. Wright Lab Design optimization studies.
NASA’s Contributions to Aeronautics

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8

UNIVERSITY APPLICATIONS OF NASTRAN PRESENTED AT USERS’ COLLOQUIA

Year University Description

1971 Old Dominion Univ Space Shuttle dynamics model.

1972

Old Dominion Univ Vibrations 

deck.

of cross-stiffened ship’s 

Louisiana Tech Univ NASTRAN as a teaching aid.

1975

Univ of MD NASTRAN for simultaneous 

bolic equations.

paraUniv NB & Mayo Graduate 

Medicine, with IBM

School of Stress 

heart.

analysis of left ventricle of the 

Univ MD, with Army, Frankford Arsenal Nonlinear 

case neck 

analysis of 

separation 

cartridge 

malfunction.

1977

Univ VA (with Sperry Marine, 

Industry” table)

listed in “Other 

1978 Univ MO Rolla NASTRAN 

research.

in education and 

1982 Air Force Inst Tech Elastic aircraft airloads.

1985

Univ of GA Agricultural engineering—teaching 

and research.

Clemson Univ Plated bone fracture gap motion.

Univ MO Fillet weld stress.

1987 Univ of Naples, with NASA Langley NASTRAN for 

interior noise.

prediction of aircraft 

1989 GWU, with DTRC Electromagnetic fields and waves.
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NASTRAN Reference Sources

At time of writing, these are available from the NASA Technical Reports 

Server at http://ntrs.nasa.gov:

NASTRAN: Users’ Experiences, NASA TM-X-2378, 1971.

NASTRAN: Users’ Experiences (2nd), NASA TM-X-2637, 1972.

NASTRAN: Users’ Experiences (4th), NASA TM-X-3278, 1975.

NASTRAN: Users’ Experiences (5th), NASA TM-X-3428, 1976.

Sixth NASTRAN Users’ Colloquium, NASA CP-2018, 1977.

Seventh NASTRAN Users’ Colloquium, NASA CP-2062, 1978.

Eight NASTRAN Users’ Colloquium, NASA CP-2131, 1979.

Ninth NASTRAN Users’ Colloquium, NASA CP-2151, 1980.

Tenth NASTRAN Users’ Colloquium, NASA CP-2249, 1982.

Eleventh NASTRAN Users’ Colloquium, NASA CP-2284, 1983.

Twelfth NASTRAN Users’ Colloquium, NASA CP-2328, 1984.

Thirteenth NASTRAN Users’ Colloquium, NASA CP-2373, 1985.

Fourteenth NASTRAN Users’ Colloquium, NASA CP-2419, 1986.

Fifteenth NASTRAN Users’ Colloquium, NASA CP-2481, 1987.

Sixteenth NASTRAN Users’ Colloquium, NASA CP-2505, 1988.

Seventeenth NASTRAN Users’ Colloquium, NASA CP-3029, 1989.

Eighteenth NASTRAN Users’ Colloquium, NASA CP-3069, 1990.

Nineteenth NASTRAN Users’ Colloquium, NASA CP-3111, 1991.

Twentieth NASTRAN Users’ Colloquium, NASA CP-3145, 1992.

Twenty-First NASTRAN Users’ Colloquium, NASA CP-3203, 1993.
NASA’s Contributions to Aeronautics

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Appendix B:

Miscellaneous NASA Structural Analysis Programs

Note: Miscellaneous computer programs, and in some cases test 

facilities or other related projects, that have contributed to the advancement of the state of the art in various ways are described here. In some 

cases, there simply was not room to include them in the main body of 

the paper; in others, there was not enough information found, or 

not enough time to do further research, to adequately describe the 

programs and document their significance. Readers are advised that 

these are merely examples; this is not an exhaustive list of all computer programs developed by NASA for structural analysis to the 2010 

time period. Dates indicate introduction of capability. Many of the programs were subsequently enhanced. Some of the programs were eventually phased out.

1) FLEXSTAB (Ames, Dryden, and Langley Research Centers, 1970s)

FLEXSTAB was a method for calculating stability derivatives that 

included the effects of aeroelastic deformation. Originally developed in 

the early 1970s by Boeing under contract to NASA Ames, FLEXSTAB 

was also used and upgraded at Dryden. FLEXSTAB used panel-method 

aerodynamic calculations, which could be readily adjusted with empirical corrections. The structural effects were treated first as a steady deformation at the trim condition, then as “unsteady perturbations about 

the reference motion to determine dynamic stability by characteristic 

roots or by time histories following an initial perturbation or following penetration of a discrete gust flow field.”191 Comparisons between 

FLEXSTAB predictions and flight measurements were made at Dryden 

for the YF-12A, Shuttle, B1, and other aircraft. Initially developed for 

symmetric flight conditions only, FLEXSTAB was extended in 1981 to 

include nonsymmetric flight conditions.192 In 1984, a procedure was 

developed to couple a NASTRAN structural model to the FLEXSTAB 

elastic-aircraft stability analysis.193 NASA Langley and the Air Force 

191. A.R. Dusto, et al., A Method for Predicting the Stability Characteristics of an Elastic Airplane,

vol. 1: FLEXSTAB Theoretical Description, NASA CR-114712 (1974), p. xxi.

192. R.L. Sims, “User’s Manual for FSLIP-3, FLEXSTAB Loads Integration Program,” NASA TM81364 (1981), p. 1.

193. Lawrence S. Schuster, “NASTRAN /FLEXSTAB Procedure for Static Aeroelastic Analysis,” 

NASA TM-84897 (1984).
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Flight Dynamics Laboratory also funded upgrades to FLEXSTAB, leading to the DYLOFLEX program, which added aeroservoelastic effects.194

2) ANSYMP Computer Program (Glenn Research Center, 1983)

ANSYMP was developed to capture the key elements of local plastic 

behavior without the overhead of a full nonlinear finite element analysis. “Nonlinear, finite-element computer programs are too costly to 

use in the early design stages for hot-section components of aircraft 

gas turbine engines. . . . This study was conducted to develop a computer program for performing a simplified nonlinear structural analysis using only an elastic solution as input data. The simplified method 

was based on the assumption that the inelastic regions in the structure 

are constrained against stress redistribution by the surrounding elastic 

material. Therefore the total strain history can be defined by an elastic 

analysis. . . . [ANSYMP] was created to predict the stress-strain history 

at the critical fatigue location of a thermomechanically cycled structure 

from elastic input data. . . . Effective [inelastic] stresses and plastic strains 

are approximated by an iterative and incremental solution procedure.” 

ANSYMP was verified by comparison to a full nonlinear finite element 

code (MARC). Cyclic hysteresis loops and mean stresses from ANSYMP 

“were in generally good agreement with the MARC results. In a typical 

problem, ANSYMP used less than 1 percent of the central processor unit 

(CPU) time required by MARC to compute the inelastic solution.”195

3) Structural Tailoring of Engine Blades (STAEBL, Glenn, 1985)

This computer program “was developed to perform engine fan blade 

numerical optimizations. These blade optimizations seek a minimum 

weight or cost design that satisfies realistic blade design constraints, by 

tuning one to twenty design variables. The STAEBL system has been generalized to include both fan and compressor blade numerical optimizations. The system analyses have been significantly improved through the 

inclusion of an efficient plate finite element analysis for blade stress and 

frequency determinations. Additionally, a finite element based approximate severe foreign object damage (FOD) analysis has been included. 

The new FOD analysis gives very accurate estimates of the full nonlinear 

194. R.D. Miller, R.I. Kroll, and R.E. Clemmons, Dynamic Loads Analysis System (DYLOFLEX) Summary, vol. 1: Engineering Formulation, NASA CR-2846-1 (1979).

195. Albert Kaufman, “Simplified Method for Nonlinear Structural Analysis,” NASA TP-2208 (1983), p. 2.
NASA’s Contributions to Aeronautics

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bird ingestion solution. Optimizations of fan and compressor blades 

have been performed using the system, showing significant cost and 

weight reductions, while comparing very favorably with refined design 

validation procedures.”196

4) TRansfer ANalysis Code to Interface Thermal and Structural (3D 

TRANCITS, Glenn, 1985)

Transfer of data between different analysis codes has always been one of 

the challenges of multidisciplinary design, analysis, and optimization. 

Even if input and output format can be standardized, different types of 

analysis often require different types of information or different mesh 

densities, globally or locally. TRANCITS was developed to translate 

between heat transfer and structural analysis codes: “TRANCITS has 

the capability to couple finite difference and finite element heat transfer 

analysis codes to linear and nonlinear finite element structural analysis 

codes. TRANCITS currently supports the output of SINDA and MARC 

heat transfer codes directly. It will also format the thermal data output directly so that it is compatible with the input requirements of the 

NASTRAN and MARC structural analysis codes. . . . The transfer module can handle different elemental mesh densities for the heat transfer 

analysis and the structural analysis.”197 MARC is a commercial, generalpurpose, nonlinear finite element code introduced by MARC Analysis 

and Research Corp. in the late 1970s. Because of its nonlinear analysis 

capabilities, MARC was used extensively at Glenn for engine component analyses and for other applications, such as the analysis of a space 

station strongback for launch loads in 1992.198 Other commercial finite 

element codes used at Glenn included MSC/NASTRAN, which was used 

along with NASA’s COSMIC version of NASTRAN.

5) COmposite Blade STRuctural ANalyzer (COBSTRAN, Glenn, 1989)

COBSTRAN was a preprocessor for NASTRAN, designed to generate 

finite element models of composite blades. While developed specifically 

196. K.W. Brown, M.S. Hirschbein, and C.C. Chamis, “Finite Element Engine Blade Structural 

Optimization,” AIAA Paper 85-0645 (1985).

197. R.L. Thompson and R.J. Maffeo, “A Computer Analysis Program for Interfacing Thermal and 

Structural Codes,” NASA TM-87021 (1985).

198. Frank F. Monasa and Joseph M. Roche, “Collapse Analysis of a Waffle Plate Strongback for 

Space Station Freedom,” NASA TM-105412 (1992).
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for advanced turboprop blades under the Advanced Turboprop (ATP) 

project, it was subsequently applied to compressor blades and turbine blades. It could be used with both COSMIC NASTRAN and MSC/

NASTRAN, and was subsequently extended to work as a preprocessor 

for the MARC nonlinear finite element code.199

6) BLAde SIMulation (BLASIM), 1992

BLASIM calculates dynamic characteristics of engine blades before 

and after an ice impact event. BLASIM could accept input geometry 

in the form of airfoil coordinates or as a NASTRAN-format finite element model. BLASIM could also utilize the ICAN program (discussed 

separately) to generate ply properties of composite blades.200 “The ice 

impacts the leading edge of the blade causing severe local damage. 

The local structural response of the blade due to the ice impact is predicted via a transient response analysis by modeling only a local patch 

around the impact region. After ice impact, the global geometry of the 

blade is updated using deformations of the local patch and a free vibration analysis is performed. The effects of ice impact location, ice size 

and ice velocity on the blade mode shapes and natural frequencies 

are investigated.”201

7) NPLOT (Goddard, 1982)

NPLOT was a product of research into the visualization of finite element 

models, which had been ongoing at Goddard since the introduction of 

NASTRAN. A fast, hidden line algorithm was developed in 1982 and 

became the basis for the NPLOT plotting program for NASTRAN, publicly released initially in 1985 and in improved versions into the 1990s.202

199. Karen F. Bartos and Michael A. Ernst, “Evaluation of MARC for the Analysis of Rotating Composite Blades,” NASA TM-4423 (1993), p. 1.

200. E.S. Reddy and G.H. Abumeri, “Blade Assessment for Ice Impact (BLASIM), User’s Manual, 

Version 1.0,” NASA CR-19075 (1993), pp.1.1–1.2.

201. G.H. Abumeri, E.S. Reddy, P.L.N. Murthy, and C.C. Chamis, “Dynamic Analysis of a Pre-andPost Ice Impacted Blade,” NASA TM-105829 (1992).

202. G.K. Jones and K.J. McEntire, “NPLOT: An Interactive Plotting Program for NASTRAN 

Finite Element Models,” in NASA, 13th NASTRAN Users’ Colloquium (1985), pp. 110–132; 

K. McEntire, “NPLOT—NASTRAN Plot,” program abstract, Jan. 1994, on NASA Technical Reports 

Server at http://ntrs.nasa.gov, accessed Apr. 20, 2009.
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8) Integrated Modeling of Optical Systems (IMOS) (Goddard and JPL, 

1990s)

A combined multidisciplinary code, IMOS was developed during the 

1990s by Goddard and JPL: “Integrated Modeling of Optical Systems 

(IMOS) is a finite element-based code combining structural, thermal, 

and optical ray-tracing capabilities in a single environment for analysis of space-based optical systems.”203 IMOS represents a recent step in 

the continuing evolution of Structural-Thermal-Optical analysis capability, which has been an important activity at the Space Flight Centers 

since the early 1970s.

9) Dynamic Simulation of Controls & Structure (Goddard, 1970s–1990s)

Another important area of spacecraft structural modeling is in the interaction of control systems with flexible multibody structural systems. In 

a general sense, this is the spacecraft counterpart to aeroservoelasticity, 

although the driving mechanisms are very different. Dynamic Simulation 

of Controls & Structure (DISCOS) was developed in the late 1970s to 

perform this type of analysis. “The physical system undergoing analysis may be generally described as a cluster of contiguous flexible structures (bodies) that comprise a mechanical system, such as a spacecraft. 

The entire system (spacecraft) or portions thereof may be either spinning or nonspinning. Member bodies of the system may undergo large 

relative excursions, such as those of appendage deployment or rotor/

stator motion. The general system of bodies is, by its inherent nature, a 

feedback system in which inertial forces (such as those due to centrifugal and Coriolis acceleration) and the restoring and damping forces are 

motion-dependent. . . . The DISCOS program can be used to obtain nonlinear and linearized time response of the system, interaction constant 

forces in the system, total system resonance properties, and frequency 

domain response and stability information for the system. DISCOS is 

probably the most powerful computational tool to date for the computer 

simulation of actively controlled coupled multi-flexible-body systems,” 

according to the computer program abstract. The program was made 

available to approved licensees (for $1,000, in 1994) with the caveat that 

DISCOS “ . . . is not easy to understand and effectively apply, but is not 

203. Gregory Moore, “Integrated Modeling of Optical Systems (IMOS): An Assessment and Future 

Directions,” Goddard Space Flight Center and Jet Propulsion Laboratory, in FEMCI Workshop 

2001, abstract on NASA Technical Reports Server at http://ntrs.nasa.gov, accessed Apr. 20, 2009.
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intended for simple problems. The DISCOS user is expected to have 

extensive working knowledge of rigid-body and flexible-body dynamics, 

finite-element techniques, numerical methods, and frequency-domain 

analysis.” DISCOS was used extensively at least into the 1990s for 

spacecraft modeling.204 In 1983, a program for bridging DISCOS, 

NASTRAN, and SAMSAN—a large order control system design 

program—was also publicly released.205 A 1987 NASA-funded 

study by Honeywell (Space and Strategic Avionics Division) outlined some limitations of DISCOS and other contemporary multibody dynamics programs and made recommendations for future 

work in the field.206 Also in the late 1980s, GSFC began collaborating with a research group at the University of Iowa that was developing similar multibody modeling capabilities for mechanical 

engineering applications. The National Science Foundation (NSF), 

U.S. Army Tank Automotive Command, and about 30 other Government 

and industry laboratories were involved in this project through the 

Industry/University Cooperative Research Center (I/UCRC) at the 

University of Iowa. Goals of the I/UCRC were to achieve mutual enhancement of capabilities in the modeling, simulation, and control of complex 

mechanical systems, including man/machine interactions applicable to 

manufacturing processes.207

10) Antenna Design and Analysis (JPL)

JPL operates both ground-based and space-borne antennas. These pose 

their own array (no pun intended) of structural challenges because of 

their large size, the need to maintain precision alignment, and, in the case 

of space-borne antennas, the need for extremely light weight. JPL was 

one of the early users of NASTRAN, using it to calculate gravity defor204. Harold P. Frisch, “DISCOS—Dynamic Interaction Simulation of Controls and Structures (IBM 

Version),” Computer Program abstract, Goddard Space Flight Center, Jan. 1994, on NASA Technical Reports Server at http://ntrs.nasa.gov, accessed Apr. 20, 2009.

205. Frisch, “NASDS-NASTRAN/DISCOS/SAMSAN DMAP Bridging Program,” Computer 

Program abstract, Goddard Space Flight Center, Jan. 1994, on NASA Technical Reports Server at 

http://ntrs.nasa.gov, accessed Apr. 20, 2009.

206. D.W. Lips, “Approaches and Possible Improvements in the Area of Multibody Dynamics 

Modeling,” NASA CR-179227 (1987).

207. Frisch, “Control system software, simulation, and robotic applications,” in NASA Goddard 

Space Flight Center, Technology 2000, vol. 1 (Washington, DC: NASA, 1991), pp. 315–321.
NASA’s Contributions to Aeronautics

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mation effects on ground-based 210-foot-diameter antennas in 1971.208

In 1976, JPL developed a simplified stiffness formulation (translational 

degrees of freedom only at the nodes) coupled with a structural member 

sizing capability for design optimization: “Computation times to execute several design/analysis cycles are comparable to the times required 

by general-purpose programs for a single analysis cycle.”209 In the late 

1980s, JPL upgraded a set of 64-meter antennas to a new diameter of 70 

meters. This project afforded “the rare opportunity to collect field data 

to compare with predictions of the finite-element analytical models.” 

Static and dynamic tests were performed while the antenna structures 

were in a stripped-down configuration during the retrofit process. The 

data provided insight into the accuracy of the models that were used to 

optimize the original structural designs.210

11) TRASYS Radiative Heat Transfer (Johnson, 1980s–1990s)

TRASYS was developed by Martin Marietta for Johnson to calculate internode radiative heat transfer as well as heat transfer to a model from the 

surroundings. It was used extensively through the 1990s. Applications 

included propulsion analysis at Glenn Research Center and StructuralThermal-Optical analysis (when integrated with NASTRAN for structural 

calculations and MACOS/IMOS and POPOS optical codes) at Marshall 

Space Center and JPL.211 BUCKY was developed in-house. BUCKY was 

initially introduced as a basic plane stress and plate buckling program 

but was extensively developed during the 1990s to include plate bending, varying element thickness, varying edge and pressure loads, edge 

moments, plasticity, output formatting for visualization in I-DEAS (a 

CAD program developed by Structural Dynamics Research Corporation), 

three-dimensional axisymmetric capability, and improvements in 

execution time.212

208. M. Smoot Katow, “Static Analysis of the 64-m (210-ft) Antenna Reflector Structure,” in NASA Langley Research Center, NASTRAN: Users’ Experiences (Hampton, VA, NASA, 1971), pp. 123–129.

209. Roy Levy, “Computer-Aided Design of Antenna Structures and Components,” JPL, in 2nd 

National Symposium on Computerized Structural Analysis and Design, Washington, DC (1976); 

abstract in NASA Technical Report Server at http://ntrs.nasa.gov, accessed Apr. 24, 2009.

210. Levy, “Structural Optimization and Recent Large Ground Antenna Installations,” in NASA, 

Recent Advances in Multidisciplinary Analysis and Optimization, pt. 3 (Washington, DC: NASA, 

1989), pp. 1393–1416.

211. NASA Technical Reports Server at http://ntrs.nasa.gov, accessed Apr. 20, 2009.

212. James P. Smith, “BUCKY Instruction Manual Version 3.3,” NASA TM-104793 (1994), p. 2.
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12) Structural Analysis of General Shells (STAGS) (Marshall and Langley, 

1960s–present)

Structural Analysis of General Shells (STAGS) evolved from early shell 

analysis codes developed by Lockheed Palo Alto Research Laboratory 

and sponsored by the NASA Marshall Space Flight Center between 1963 

and 1968, with subsequent development funded primarily from Langley. 

B.O. “Bo” Almroth of Lockheed was the principal developer. The name 

STAGS seems to have first appeared around 1970.213 Thus, STAGS initial development was nearly concurrent with that of NASTRAN. While 

NASTRAN development aimed to stem the proliferation of analysis 

codes, and of shell analysis codes in particular, NASTRAN did not initially provide the full capability needed to replace such codes. In particular, STAGS from the beginning included nonlinear capability that 

was found necessary in the accurate modeling of shells with cutouts. In 

the mid- to late 1970s, STAGS was released publicly, with user manuals. “Under contract with NASA, STAGS has been converted from being 

more or less a pure research tool into a code that is suitable for use by 

the public for practical engineering analysis. Suggestions from NASALangley have resulted in considerable enhancement of the code and are 

to some degree the cause of its increasing popularity. . . . User reaction 

consistently seems to indicate that the run time with STAGS is surprisingly low in comparison to comparable codes. A STAGS input deck is 

usually compact and time for its preparation is short.”214 STAGS continued to be enhanced through the 1980s (as STAGS-C1, actually a family of versions), offering unique capabilities for modeling total collapse 

of structures and problems that bifurcate into multiple possible solutions.215 It was apparently popular and widely used. For example, in 

1990, Engineering Dynamics, Inc., of Kenner, LA, used STAGS-C1 to 

model and verify a repair design for a damaged offshore oil platform.216

213. Norman F. Knight, Jr., and Charles C. Rankin, “STAGS Example Problems Manual,” NASA 

CR-2006-214281 (2006), p. 5; B.O. Almroth and A.M. Holmes, “An experimental study of the 

strength and stability of thin monocoque shells with reinforced and unreinforced rectangular cutouts,” 

NASA CR-115267 (1971), abstract on NASA Technical Reports Server at http://ntrs.nasa.gov, 

accessed May 10, 2009.

214. B.O. Almroth and F.A. Brogan, “The STAGS Computer Code,” NASA CR-2950 (1978), pp. 1–2.

215. C.C. Rankin, P. Stehlin, and F.A. Brogan, “Enhancements to the STAGS Computer Code,” 

NASA CR-4000 (1986), pp. 1, 3.

216. “Structural Analysis,” Spinoff, NASA Technology Utilization Office, 1991, p. 57.
NASA’s Contributions to Aeronautics

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8

STAGS Version 5.0 was released in 2006, and STAGS is still used for failure analysis, analysis of damaged structures, and similar problems.217

13) Nonlinear Structures: PANES (1975) and AGGIE-I (1980) (Marshall)

Program for Analysis of Nonlinear Equilibrium and Stability (PANES) 

was developed for structural problems involving geometric and/or material nonlinear characteristics. AGGIE-I was a more comprehensive code 

capable of solving larger and more general problems, also involving geometric or material nonlinearities.218

14) Finite Element Modeling of Piping Systems (Stennis)

While Stennis is not active in structural methods research, there have 

been some activities applying finite element and structural health monitoring techniques to the complex fuel distribution systems at the facility. 

One such effort was presented at the 27th Joint Propulsion Conference 

in Sacramento, CA, in 1991: “A set of PC-based computational Dynamic 

Fluid Flow Simulation models is presented for modeling facility gas and 

cryogenic systems. . . . A set of COSMIC NASTRAN-based finite element 

models is also presented to evaluate the loads and stresses on test facility piping systems from fluid and gaseous effects, thermal chill down, 

and occasional wind loads. The models are based on Apple Macintosh 

software which makes it possible to change numerous parameters.”219

NASA was, in this case, its own spinoff technology customer.

217. Knight and Rankin, “STAGS Example Problems Manual”; Donald J. Baker, “Response of Damaged and Undamaged Tailored Extension-Shear-Coupled Composite Panels,” Journal of Aircraft, 

vol. 43, no. 2 (2008), pp. 517–527.

218. Search of NASA Technical Reports Server at http://ntrs.nasa.gov, accessed Apr. 20, 2009.

219. L. Dequay, A. Lusk, and S. Nunez, “Integrated flow and structural modeling for rocket engine 

component test facility propellant systems,” AIAA Paper 91-2402 (1991).
Case 8 | NASA and Computational Structural Analysis

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Appendix C: 

Structural Analysis and Loads Prediction Facilities

Test facilities have an important role in verifying and improving analysis methods. A few test facilities that had a lot to do with the development and validation of structural analysis methods are described 

below. In addition to those described, other “landmark” test facilities 

include large-scale launch vehicle structural test facilities at Johnson 

and Marshall Space Centers, and the crash dynamics test facility at 

Langley Research Center.

Structural Dynamics Laboratory (Ames Research Center, 1965)

During the 1960s, Ames and Langley collaborated on some of the structural dynamics and buffet problems of spacecraft during ascent. (This 

collaboration occurred through some of the same meetings at NASA 

Headquarters that led to the development of NASTRAN.) To help assess 

the structural dynamic characteristics of boosters, and to build confidence in predictive methods, a large structural dynamics test facility 

was built at Ames (completed in 1965). This facility was large enough 

to hold a full-size Atlas or Titan II, had provisions for exciting the structural modes of the test article, and could be evacuated to test the structural damping characteristics in zero or reduced ambient air density.220

The facility was also used for research on buffet during reentry and landing impacts.221 Much of the structural dynamics research at Ames was 

discontinued or relocated during the early 1970s. The laboratory is long 

since deactivated, but the large, pentagonal tower still stands, housing a 

machine shop and a wind tunnel that can simulate Mars’s atmosphere 

by evacuating the chamber and then filling to low pressure with CO2.222

Thermal Loads Laboratory (Dryden Flight Research Center, 1960s)

A 1973 accounting of NASA research facilities listed only one major 

ground laboratory at Dryden: the High Temperature Loads Calibration 

Laboratory.223 High supersonic and hypersonic flight research created 

a need (1) to test airframes on the ground under simultaneous thermal 

and structural loading conditions and (2) to calibrate loads instrumen220. Henry A. Cole, e-mail message to author, Apr. 19, 2009.

221. Glenn E. Bugos, e-mail to author, Apr. 3, 2009.

222. Smith interview by author (by telephone), Apr. 6, 2009.

223. Hallion and Gorn, On The Frontier, pp. 100, 358.
NASA’s Contributions to Aeronautics

554

8

tation at elevated temperatures, so that the data obtained in flight could 

be reliably interpreted. These needs “ . . . led to the construction of a 

laboratory for calibrating strain-gage installations to measure loads in 

an elevated temperature environment. The problems involved in measuring loads with strain gages . . . require the capability to heat and 

load aircraft under simulated flight conditions. . . . The laboratory has 

the capability of testing structural components and complete vehicles 

under the combined effects of loads and temperatures, and calibrating 

and evaluating flight loads instrumentation under [thermal] conditions 

expected in flight.”224

The laboratory is housed in a hangarlike building with attached 

shop, offices, and control room. Capabilities included:

• 

• 

• 

• 

• 

• 

• 

Hangar-door opening 40 feet high by 136 feet wide.

Unobstructed test area 150 by 120 by 40 feet allowed 

the testing of aircraft up to and including, for example, 

a YF-12 or SR-71.

Ten megawatts of electrical heating power via quartz 

lamps and reflectors.

Temperatures up to 3,000 °F.

Hydraulic power of 4.5 gallons/minute at 3,000 pounds 

per square inch (psi) to apply loads.

Fourteen channels closed-loop load or position control 

of up to 34 separate actuators.

Sensors including strain gages, thermocouples, load 

cells, and position transducers.

Slots in the floor provided flexible locations for tiedown points, as 

well as routing for hydraulic and electrical power, instrumentation wiring, compressed air, or water (presumably for cooling). Closed-loop analog control of both mechanical load and heating was provided, to any 

desired preprogrammed time history.

The facility was used in the YF-12 thermal loads project (discussed 

elsewhere in this paper), in Space Shuttle structural verification at high 

224. Sefic and Karl F. Anderson, “NASA High Temperature Loads Calibration Laboratory,” NASA 

TM-X-1868 (1969), p. 1.
Case 8 | NASA and Computational Structural Analysis

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8

temperatures, and for a variety of other studies.225 The loads laboratory 

made contributions to the validation of computational methods by providing the opportunity to compare computational predictions with test 

data obtained under known, controlled, thermal and structural loading conditions, applied together or independently as required. At time 

of this writing, the facility is still in use.226

Spin Rig (Glenn Research Center, 1983)

One particular facility of many, a spin test rig built at Lewis in 1983 is 

mentioned here because its stated purpose was not primarily the testing of engine parts to verify the parts but the testing of engine parts to 

verify analysis methods: “The Lewis Research Center spin rig was constructed to provide experimental evaluation of analysis methods developed under the NASA Engine Structural Dynamics Program. Rotors up 

to 51 cm (20 in.) in diameter can be spun to 16,000 rpm in vacuum by 

an air motor. Vibration forcing functions are provided by shakers that 

apply oscillatory axial forces or transverse moments to the shaft, by a 

natural whirling of the shaft, and by an air jet. Blade vibration is detected 

by strain gages and optical tip blade-motion sensors.”227

225. “Fact Sheets: NASA Dryden’s Contributions to Spaceflight,” http://www.nasa.gov/centers/

dryden/news/FactSheets, accessed Mar. 19, 2009.

226. Martin Brenner, e-mail message to author, May 8, 2009.

227. Gerald V. Brown, Robert E. Kielb, Erwin H. Meyn, Richard E. Morris, and Stephen J. Posta, 

“Lewis Research Center Spin Rig and Its Use in Vibration Analysis of Rotating Systems,” NASA TP2304 (1984), p. 1.
NASA’s Contributions to Aeronautics

556

8

Appendix D: 

NASA-Developed Finite Element Analysis

• 

• 

• 

• 

• 

• 

• 

• 

• 

NASTRAN (COSMIC versions). NASA case Nos. COS-10054, -10057, 

-10061, and -10064 through -10067.

ICAN Integrated Composite Analyzer. Linear finite element analysis 

of multilayered fiber composites. NASA case No. LEW-15592. Lewis 

Research Center.

AIRLOADS. Oscillatory airloads on turbosystem blades in nonuniform flow. NASA case no. LEW-14947. Lewis Research Center.

ACTON AutoCAD to NASTRAN translator. NASA case no. GSC13217. Goddard Space Flight Center.

ACTOMP, AutoCAD to Mass Properties. Develops mass properties 

data from AutoCAD models. NASA case no. GSC-13228. Goddard 

Space Flight Center.

GeoFEST, Geophysical Finite Element Simulation. Developed by Jet 

Propulsion Laboratory in 2002 and released directly through OCS.

FECAP Finite Element Composite Analysis Program for microcomputers. NASA case no. LAR-14109. Langley Research Center.

NASTPLT, plotting post-processor for NASTRAN. NASA case no. 

GSC-12833. Goddard Space Flight Center.

STARS. Integrated multidisciplinary structural, fluids, aeroelastic, 

and aeroservoelastic analysis. NASA case no. FRC-09537. Dryden 

Flight Research Center.
Case 8 | NASA and Computational Structural Analysis

557

8

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The X-43A Hyper-X test vehicle drops away from its Boeing NB-52B mother ship in November 

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569

9

High-Temperature 

Structures and Materials

T.A. Heppenheimer

CASE

9

Taking fullest advantage of the high-speed potential of rocket and airbreathing propulsion systems required higher-temperature structures. 

Researchers recognized that aerothermodynamics involved linking 

aerodynamic and thermodynamic understanding with the mechanics 

of thermal loading and deformation of structures. This drove use of new 

structural materials. NASA and other engineers would experiment with 

active and passive thermal protection systems, metals, and materials.

I

N AEROSPACE ENGINEERING, high-temperature structures and 

materials solve two problems. They are used in flight above Mach 2 

to overcome the elevated temperatures that occur naturally at such 

speeds. They also are extensively used at subsonic velocities, in building 

high-quality turbofan engines, and for the protection of structures 

exposed to heating.

Aluminum loses strength when exposed to temperatures above 210 

degrees Fahrenheit (°F). This is why the Concorde airliner, which was 

built of this material, cruised at Mach 2.1 but did not go faster.1 Materials 

requirements come to the forefront at higher speeds and escalate sharply 

as airplanes’ speeds increase. The standard solutions have been to use 

titanium and nickel, and a review of history shows what this has meant.

Many people wrote about titanium during the 1950s, but to reduce 

it to practice was another matter. Alexander “Sasha” Kartveli, chief 

designer at Republic Aviation, proposed a titanium F-103 fighter, but 

his vision outreached his technology, and although started, it never 

flew. North American Aviation’s contemporaneous Navaho missile program introduced chemical milling (etching out unwanted material) for 

aluminum as well as for titanium and steel, and was the first to use 

titanium skin in an aircraft. However, the version of Navaho that 

1. Erik M. Conway, High-Speed Dreams (Baltimore: Johns Hopkins University Press, 2005), p. 57.
NASA’s Contributions to Aeronautics

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9

The Lockheed Blackbird experienced a wide range of upper surface temperatures, up to 600 °F. NASA.

was to use these processes never flew, as the program was canceled 

in 1957.2

The Lockheed A-12 Blackbird, progenitor of a family of exotic Mach 

3.2 cruisers that included the SR-71, encountered temperatures as high 

as 1,050 °F, which required that 93 percent of its structural weight be titanium. The version selected was B-120 (Ti-13V-11Cr-3Al), which has the 

tensile strength of stainless steel but weighs only half as much. But titanium is not compatible with chlorine, cadmium, or fluorine, which led 

to difficulties. A line drawn on a sheet of titanium with a pen would eat 

a hole into it in a few hours. Boltheads tended to fall away from assemblies; this proved to result from tiny cadmium deposits made by tools. 

This brought removal of all cadmium-plated tools from toolboxes. Spotwelded panels produced during the summer tended to fail because the 

local water supply was heavily chlorinated to kill algae. The managers 

took to washing the parts in distilled water, and the problem went away.3

2. Richard A. DeMeis, “The Trisonic Titanium Republic,” Air Enthusiast, No. 7, (July–Sept. 1978), 

pp. 198–213; Dale D. Myers, “The Navaho Cruise Missile: A Burst of Technology,” Acta Astronautica, vol. 26 (Nov. 8–10, 1992), pp. 741–748.

3. David Robarge, Archangel: CIA’s Supersonic A-12 Reconnaissance Aircraft (McLean, VA: CIA 

Center for the Study of Intelligence, 2007), pp. 11–12; see also Paul F. Crickmore, Lockheed 

SR-71 Blackbird (London: Osprey Publishing Ltd., 1986), pp. 90–92.
Case 9 | High-Temperature Structures and Materials

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The SR-71 was a success. Its shop-floor practice with titanium at first 

was classified but now has entered the aerospace mainstream. Today’s 

commercial airliners—notably the Boeing 787 and the Airbus A-380, 

together with their engines—use titanium as a matter of routine. That 

is because this metal saves weight.

Beyond Mach 4, titanium falters and designers must turn instead to 

alternatives. The X-15 was built to top Mach 6 and to reach 1,200 °F. In 

competing for the contract, Douglas Aircraft proposed a design that was 

to use magnesium, whose properties were so favorable that the aircraft 

would only reach 600 °F. But this concept missed the point, for managers wanted a vehicle that would cope successfully with temperatures of 

1,200 °F. Hence it was built of Inconel X, a nickel alloy.4

High-speed flight represents one application of advanced metals. 

Another involves turbofans for subsonic flight. This application lacks 

the drama of Mach-breaking speeds but is far more common. Such 

engines use turbine blades, with the blade itself being fabricated from 

a single-crystal superalloy and insulated with ceramics. Small holes 

in the blade promote a circulation of cooler gas that is ducted downstream from high-pressure stages of the compressor. The arrangement 

can readily allow turbines to run at temperatures 750 °F above the melting point of the superalloy itself.5

The High-Speed Environment

During World War II the whole of aeronautics used aluminum. There 

was no hypersonics; the very word did not exist, for it took until 1946 

for the investigator Hsue-shen Tsien to introduce it. Germany’s V-2 

was flying at Mach 5, but its nose cone was of mild steel, and no one 

expected that this simple design problem demanded a separate term 

for its flight regime.6

A decade later, aeronautics had expanded to include all flight speeds 

because of three new engines: the liquid-fuel rocket, the ramjet, and the 

variable-stator turbojet. The turbojet promised power beyond Mach 3, 

while the ramjet proved useful beyond Mach 4. The Mach 6 X-15 was 

under contract. Intermediate-range missiles were in development, with 

4. “Evaluation Report on X-15 Research Aircraft Design Competition,” Aug. 5, 1955, pp. 99–101, 

Record Group 255, Philadelphia Federal Records Center, National Archives and Records Service.

5. “Briefing: Rolls-Royce,” The Economist (Jan. 10, 2009), pp. 60–62.

6. Willy Ley, Rockets, Missiles, and Space Travel (New York: Viking, 1957), p. 213.

9
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ranges of 1,200 to 1,700 miles, and people regarded intercontinental 

missiles as preludes to satellite launchers.

A common set of descriptions presents the flight environments within 

which designers must work. Well beyond Mach 3, engineers accommodate aerodynamic heating through materials substitutions. The aircraft 

themselves continue to accelerate and cruise much as they do at lower 

speeds. Beyond Mach 4, however, cruise becomes infeasible because of 

heating. A world airspeed record for air-breathing flight (one that lasted 

for nearly the next half century) was set in 1958 with the Lockheed X-7, 

which was made of 4130 steel, at Mach 4.31 (2,881 mph). The airplane 

had flown successfully at Mach 3.95, but it failed structurally in flight 

at Mach 4.31, and no airplane has approached such performance in 

the past half century.7

No aircraft has ever cruised at Mach 5, and an important reason 

involves structures and materials. “If I cruise in the atmosphere for 2 

hours,” said Paul Czysz of McDonnell-Douglas, “I have a thousand 

times the heat load into the vehicle that the Shuttle gets on its quick 

transit of the atmosphere.”8 Aircraft indeed make brief visits to such 

speed regimes, but they don’t stay there; the best approach is to pop 

out of the atmosphere and then return, the hallmark of a true transatmospheric vehicle.

At Mach 4, aerodynamic heating raises temperatures. At higher 

Mach, other effects are seen. A reentering intercontinental ballistic missile (ICBM) nose cone, at speeds above Mach 20, has enough kinetic 

energy to vaporize 5 times its weight in iron. Temperatures behind its 

bow shock reach 9,000 kelvins (K), hotter than the surface of the Sun. 

The research physicist Peter Rose has written that this velocity would 

be “large enough to dissociate all the oxygen molecules into atoms, 

dissociate about half of the nitrogen, and thermally ionize a considerable fraction of the air.”9

7. Jay Miller, The X-Planes, X-1 to X-45 (North Branch, MN: Specialty Press, 2001). See also Lee L. 

Peterson, “Evaluation Report on X-7A,” Report AFMDC ADJ 57-8184, Oct. 3, 1957; and William 

A. Ritchie, “Evaluation Report on X-7A (System 601B),” Report AFMDC DAS 58-8129, Jan. 1959, 

copies in the Archives of the Air Force Historical Research Agency, Maxwell AFB, Montgomery, AL.

8. Author interview with Paul Czysz, Mar. 13, 1986.

9. Time, June 13, 1960, p. 70; P.H. Rose and W.I. Stark, “Stagnation Point Heat-Transfer Measurements in Dissociated Air.” Journal of the Aeronautical Sciences (Feb. 1958), pp. 86–97; P.H. Rose, 

“Physical Gas Dynamics Research at the Avco Research Lab,” AGARD Report 145, July 1957, p. 1.
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Aircraft thus face a simple rule: they can cruise up to Mach 4 if built 

with suitable materials, but they cannot cruise at higher speeds. This rule 

applies not only to entry into Earth’s atmosphere but also to entry into 

the atmosphere of Jupiter, which is far more demanding but which an 

entry probe of the Galileo spacecraft investigated in 1995, at Mach 50.10

Other speed limits become important in the field of wind tunnel 

simulation. The Government’s first successful hypersonic wind tunnel was John Becker’s 11-inch facility, which entered service in 1947. 

It approached Mach 7, with compressed air giving run times of 40 seconds.11 A current facility, which is much larger and located at the National 

Aeronautics and Space Administration (NASA) Langley Research 

Center, is the Eight-Foot High-Temperature Tunnel—which also uses 

compressed air and operates near Mach 7.

The reason for such restrictions involves fundamental limitations 

of compressed air, which liquefies if it expands too much when seeking 

higher speeds. Higher speeds indeed are achievable but only by creating shock waves within an instrument for periods measured in milliseconds. Hence, the field of aerodynamics introduces an experimental 

speed limit of Mach 7, which describes its wind tunnels, and an operational speed limit of Mach 4, which sets a restriction within which cruising flight remains feasible. Compared with these velocities, the usual 

definition of hypersonics, describing flight beyond Mach 5, is seen to 

describe nothing in particular.

Metals, Ceramics, and Composites

Solid-state materials exist in one of these forms and may be reviewed 

separately. Metals and alloys, the latter being particularly common, exist 

usually as superalloys. These are defined as exhibiting excellent 

mechanical strength and creep resistance at high temperatures, good 

surface stability, and resistance to corrosion and oxidation. The 

base alloying element of a superalloy is usually nickel, cobalt, or nickeliron. These three elements are compared in Table 1 with titanium.12 

10. Richard E. Young, Martha A. Smith, and Charles K. Sobeck, “Galileo Probe: In Situ Observations of Jupiter’s Atmosphere,” Science (May 10, 1996), pp. 837–838.

11. James R. Hansen, Engineer in Charge: A History of the Langley Aeronautical Laboratory, 

1917–1958, NASA SP-4305 (Washington: GPO, 1987), pp. 343–347.

12. Data from Matthew J. Donachie, Jr., Superalloys Source Book (Metals Park, OH: American 

Society for Metals, 1984), pp. 3–19.

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9

TABLE 1:

COMPARISON OF TITANIUM WITH SELECTED SUPER ALLOYS

ELEMENT NUMBER MELTING POINT (K)

Titanium 22 1,941

Iron 26 1,810

Cobalt 27 1,768

Nickel 28 1,726

Superalloys generally are used at temperatures above 1,000 °F, or 
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TABLE 2:

SELECTED ALLOYING ADDITIONS AND THEIR EFFECTS

ELEMENT PERCENTAGES EFFECT

Iron-nickel- and nickel-base Cobalt-base

Chromium 5–25 19–30 Oxidation and hot 

corrosion resistance; 

solution hardening; 

carbides

Molybdenum, 

Tungsten

0–12 0–11 Solution hardening; 

carbides

Aluminum 0–6 0–4.5 Precipitation hardening; oxidation 

resistance

Titanium 0–6 0–4 Precipitation hardening; carbides

Cobalt 0–20 N/A Affects amount 

precipitate

of 

Nickel N/A 0–22 Stabilizes austenite; 

forms hardening 

precipitates

Niobium 0–5 0–4 Carbides; solution 

hardening; precipitation hardening (nickel-, 

iron-nickel-base)

Tantalum 0–12 0–9 Carbides; solution 

hardening; oxidation 

resistance

Table 2 presents a selection of alloying additions, together with 

their effects.13 The superalloys generally react with oxygen, oxidation being the prime environmental effect on these alloys. General 

oxidation is not a major problem up to about 1,600 °F, but at higher 

temperatures, commercial nickel-and cobalt-base superalloys are attacked 

by oxygen. Below about 1,800 °F, oxidation resistance depends on 

chromium content, with Cr2O3 forming as a protective oxide; at higher 

temperatures, chromium and aluminum contribute in an interactive 

fashion to oxidation protection, with aluminum forming the protective 

Al2O3. Because the level of aluminum is often insufficient to provide 

13. Donachie, Introduction to Superalloys, Ref. 13, p. 15.

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9

long-term protection, protective coatings are often applied. Cobaltbase superalloys are readily welded using gas-metal-arc (GMA) or 

gas-tungsten-arc (GTA) techniques. Nickel- and iron-nickel-base superalloys are considerably less weldable, for they are susceptible to hot 

cracking, postweld heat treatment cracking, and strain-age cracking. 

However, they have been successfully welded using GMA, GTA, electron-beam, laser, and plasma arc methods. Superalloys are difficult to 

weld when they contain more than a few percentage points of titanium 

and aluminum, but superalloys with limited amounts of these alloying 

elements are readily welded.14

So much for alloys. A specific type of fiber, carbon, deserves discussion in its own right because of its versatility. It extends the temperature 

resistance of metals by having the unparalleled melting temperature of 

6,700 °F. Indeed, it actually gains strength with temperature, being up 

to 50 percent stronger at 3,000 °F than at room temperature. It also has 

density of only l.50 grams per cubic centimeter (g/cm3). These properties allowed carbon fiber to serve in two path-breaking vehicles of recent 

decades. The Voyager aircraft, which flew around the world in 1986 on 

a single load of fuel, had some 90 percent of its structure made of carbon fibers in a lightweight matrix. The Space Shuttle also relies on carbon for thermal protection of the nose and wing leading edges.15

These areas needed particularly capable thermal protection, and 

carbon was the obvious candidate. It was lighter than aluminum and 

could be protected against oxidation with a coating. Graphite was 

initially the standard form, but it had failed to enter the aerospace 

mainstream. It was brittle and easily damaged, and it did not lend itself 

to use with thin-walled structures.

The development of a better carbon began in 1958 with Vought 

Missiles and Space Company (later LTV Aerospace) in the forefront. The 

work went forward with support from the Dyna-Soar and Apollo programs and brought the advent of an all-carbon composite consisting of 

graphite fibers in a carbon matrix. Existing composites had names such 

as carbon-phenolic and graphite-epoxy; this one was carbon-carbon.

It retained the desirable properties of graphite in bulk: lightweight, 

temperature resistance, and resistance to oxidation when coated. It had 

14. Ibid., pp. 3–19.

15. Deborah D.L. Chung, Carbon Fiber Composites (Boston: Butterworth-Heinemann, 1994), pp. 

5–10.
Case 9 | High-Temperature Structures and Materials

577

a very low coefficient of thermal expansion, which reduced thermal 

stress. It also had better damage tolerance than graphite.

Carbon-carbon was a composite. As with other composites, Vought 

engineers fabricated parts of this material by forming them as layups. 

Carbon cloth gave a point of departure, being produced by oxygen-free 

pyrolysis of a woven organic fiber such as rayon. Sheets of this fabric, 

impregnated with phenolic resin, were stacked in a mold to form the 

layup and then cured in an autoclave. This produced a shape made of 

laminated carbon cloth phenolic. Further pyrolysis converted the resin 

to its basic carbon, yielding an all-carbon piece that was highly porous 

because of the loss of volatiles. It therefore needed densification, which 

was achieved through multiple cycles of reimpregnation under pressure 

with an alcohol, followed by further pyrolysis. These cycles continued 

until the part had its specified density and strength.

The Shuttle’s design specified carbon-carbon for the nose cap and 

leading edges, and developmental testing was conducted with care. 

Structural tests exercised their methods of attachment by simulating 

flight loads up to design limits, with design temperature gradients. Other 

tests, conducted within an arc-heated facility, determined the thermal 

responses and hot-gas leakage characteristics of interfaces between the 

carbon-carbon and the rest of the vehicle.

Additional tests used articles that represented substantial portions of 

the orbiter. An important test item, evaluated at NASA Johnson, reproduced a wing-leading edge and measured 5 by 8 feet. It had two leadingedge panels of carbon-carbon set side by side, a section of wing structure 

that included its main spars, and aluminum skin covered with thermalprotection tiles. It had insulated attachments, internal insulation, and 

internal seals between the carbon-carbon and the tiles. It withstood simulated air loads, launch acoustics, and mission temperature-pressure 

environments—not once but many times.16

There was no doubt that left to themselves, the panels of carboncarbon that protected the leading edges would have continued to do so. 

Unfortunately, they were not left to themselves. During the ascent of the 

16. Paul R. Becker, “Leading-Edge Structural Material System of the Space Shuttle,” American 

Ceramic Society Bulletin, vol. 60, No. 11 (1981), pp. 1210–1214; L.J. Korb, C.A. Morant, R.M. 

Calland, and C.S. Thatcher, “The Shuttle Orbiter Thermal Protection System,” American Ceramic 

Society Bulletin, vol. 60 (1981), pp. 1188–1193; Anon., “Technical Overview: Oxidation Resistant Carbon-Carbon for the Space Shuttle,” Vought Missiles and Space Company, n.d. (c. 1970).

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9

Shuttle Columbia, on January 16, 2003, a large piece of insulating foam 

detached itself from a strut that joined the external tank to the front of 

the orbiter. The vehicle at that moment was slightly more than 80 seconds into the flight, traveling at nearly Mach 2.5. This foam struck a 

carbon-carbon panel and delivered what proved to be a fatal wound. In 

words of the accident report:

Columbia re-entered Earth’s atmosphere with a preexisting 

breach in the leading edge of its left wing. This breach, caused by 

the foam strike on ascent, was of sufficient size to allow superheated air (probably exceeding 5,000 degrees Fahrenheit) to 

penetrate the cavity behind the RCC panel. The breach widened, 

destroying the insulation protecting the wing’s leading edge 

support structure, and the superheated air eventually melted 

the thin aluminum wing spar. Once in the interior, the superheated air began to destroy the left wing. Finally, over Texas, the 

increasing aerodynamic forces the Orbiter experienced in the 

denser levels of the atmosphere overcame the catastrophically 

damaged left wing, causing the Orbiter to fall out of control.17

Three years of effort succeeded in securing the foam on future flights, 

and the Shuttle returned to flight in July 2006 with foam that stayed 

put. In contrast with the high tech of the Shuttle, carbon fibers also are 

finding use in such low-tech applications as automobiles. As with the 

Voyager round-the-world aircraft, what counts is carbon’s light weight, 

which promotes fuel economy. The Graphite Car employs carbon fiber 

epoxy-matrix composites for body panels, structural members, bumpers, wheels, drive shafts, engine components, and suspension systems. 

A standard steel auto would weigh 4,000 pounds, but this car weighs 

only 2,750 pounds, for a saving in weight of nearly one-third.18

Superalloys thus represent the mainstream in aerospace materials, 

with composites such as carbon fiber extending their areas of use. There 

also are ceramics, but these are highly specialized. They cannot compete with the temperature resistance of carbon or with its light weight. 

They nevertheless come into play as insulators on turbine blades that 

protect the underlying superalloy. This topic will be discussed separately.

17. NASA, Columbia Accident Investigation Report, (Washington, DC: GPO, 2003), p. 12.

18. Chung, Carbon Fiber Composites, pp. 116–118.
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579

Ablative and Radiative Structures

Atmosphere entry of satellites takes place above Mach 20, only slightly 

faster than the speed of reentry of an ICBM nose cone. The two phenomena nevertheless are quite different. A nose cone slams back at a 

sharp angle, decelerating rapidly and encountering heating that is brief 

but very severe. Entry of a satellite is far easier, taking place over a 

number of minutes.

To learn more about nose cone reentry, one begins by considering 

the shape of a nose cone. Such a vehicle initially has high kinetic energy 

because of its speed. Following entry, as it approaches the ground, its 

kinetic energy is very low. Where has it gone? It has turned into heat, 

which has been transferred both into the nose cone and into the air that 

has been disturbed by passage of the nose cone. It is obviously of interest to transfer as much heat as possible into the surrounding air. During 

reentry, the nose cone interacts with this air through its bow shock. For 

effective heat transfer into the air, the shock must be very strong. Hence 

the nose cone cannot be sharp like a church steeple, for that would 

substantially weaken the shock. Instead, it must be blunt, as H. Julian 

Allen of the National Advisory Committee for Aeronautics (NACA) first 

recognized in 1951.19

Now that we have this basic shape, we can consider methods for 

cooling. At the outset of the Atlas ICBM program, in 1953, the simplest method of cooling was the heat sink, with a thick copper shield 

absorbing the heat of reentry. An alternative approach, the hot structure, called for an outer covering of heat-resistant shingles that were 

to radiate away the heat. A layer of insulation, inside the shingles, was 

to protect the primary structure. The shingles, in turn, overlapped 

and could expand freely.

A third approach, transpiration cooling, sought to take advantage of 

the light weight and high heat capacity of boiling water. The nose cone 

was to be filled with this liquid; strong g-forces during deceleration in 

the atmosphere were to press the water against the hot inner skin. The 

skin was to be porous, with internal steam pressure forcing the fluid 

19. See George W. Sutton, “The Initial Development of Ablation Heat Protection, An Historical 

Survey,” Journal of Spacecraft and Rockets, (Jan.–Feb. 1982), pp. 3–11; and H. Julian Allen and 

A.J. Eggers, Jr., “A Study of the Motion and Aerodynamic Heating of Ballistic Missiles Entering the 

Earth’s Atmosphere at High Supersonic Speeds,” NACA TR-1381 (1953), which summarizes his 

early research and that of Alfred J. Eggers.

9
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9

An Atlas ICBM with a low-drag ablatively cooled nose cone. USAF.

through the pores and into the boundary layer. Once injected, steam was 

to carry away heat. It would also thicken the boundary layer, reducing 

its temperature gradient and hence its rate of heat transfer. In effect, 

the nose cone was to stay cool by sweating.

Still, each of these approaches held difficulties. Transpiration cooling 

was poorly understood as a topic for design. The hot-structure concept 

raised questions of suitably refractory metals along with the prospect 
Case 9 | High-Temperature Structures and Materials

581

of losing the entire nose cone if a shingle came off. Heat sinks appeared 

to promise high weight. But they seemed the most feasible way to 

proceed, and early Atlas designs specified use of a heat-sink nose cone.20

Atlas was an Air Force program. A separate set of investigations 

was underway within the Army, which supported hot structures but 

raised problems with both heat sink and transpiration. This work anticipated the independent studies of General Electric’s George Sutton, with 

both efforts introducing an important new method of cooling: ablation. 

Ablation amounted to having a nose cone lose mass by flaking off when 

hot. Such a heat shield could absorb energy through latent heat, when 

melting or evaporating, and through sensible heat, with its temperature rise. In addition, an outward flow of ablating volatiles thickened 

the boundary layer, which diminished the heat flow. Ablation promised 

all the advantages of transpiration cooling, within a system that could 

be considerably lighter and yet more capable, and that used no fluid.21

Though ablation proved to offer a key to nose cone reentry, experiments showed that little if any ablation was to be expected under the relatively mild conditions of satellite entry. But satellite entry involved high 

total heat input, while its prolonged duration imposed a new requirement for good materials properties as insulators. They also had to stay 

cool through radiation. It thus became possible to critique the usefulness of ICBM nose cone ablators for the new role of satellite entry.22

Heat of ablation, in British thermal units (BTU) per pound, had 

been a standard figure of merit. Water, for instance, absorbs nearly 1,000 

BTU/lb when it vaporizes as steam at 212 °F. But for satellite entry, with 

little energy being carried away by ablation, head of ablation could be 

irrelevant. Phenolic glass, a fine ICBM material with a measured heat 

of 9,600 BTU/lb, was unusable for a satellite because it had an unacceptably high thermal conductivity. This meant that the prolonged thermal soak of a satellite entry could have enough time to fry a spacecraft. 

20. C.E. Brown, W.J. O’Sullivan, and C.H. Zimmerman, “A Study of the Problems Related to High 

Speed, High Altitude Flight,” NACA Langley, 1953; Jacob Neufeld, The Development of Ballistic Missiles in the United States Air Force, 1945–1960 (Washington: Office of Air Force History, 1990).

21. Army Ballistic Missile Agency staff, Re-Entry Studies, vol. 1. (Redstone Arsenal, AL: ABMA,

Nov. 25, 1958), pp. 2, 24–25, 31, 37–45, 61; Sutton, “Initial Development of Ablation Heat Protection,” pp. 3–11.

22. E.R. Riddell and J.D. Teare, “The Differences Between Satellite and Ballistic Missile Re-Entry 

Problems,” in Morton Alperin and Hollingsworth F. Gregory, eds., Vistas in Astronautics, vol. 2 

(New York: Pergamon Press, 1959,) pp. 174–190.

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9

Teflon, by contrast, had a measured heat only one-third as large. It 

nevertheless made a superb candidate because of its excellent properties as an insulator.23

Hence it became possible to treat the satellite problem as an extension of the ICBM problem. With appropriate caveats, the experience 

and research techniques of the ICBM program could carry over to this 

new realm. The Central Intelligence Agency was preparing to recover 

satellite spacecraft at the same time that the Air Force was preparing 

to fly full-size Atlas nose cones, with both being achieved in April 1959.

The Army flew a subscale nose cone to intermediate range in August 

1958, which President Dwight Eisenhower displayed during a November 

news conference. The Air Force became the first to fly a nose cone to 

intercontinental range, in July 1958. Both flights carried a mouse, and 

both mice survived their reentry, but neither was recovered. Better 

success came the following April, when an Atlas launched the fullsize RVX-l nose cone, and the Discoverer II reconnaissance spacecraft 

returned safely through the atmosphere—though it fell into Russian, not 

American, hands.24

Hot Structures: Dyna-Soar

Reentry of ICBM nose cones and of satellites takes place at nearly the 

same velocity. Reentry of spacecraft takes place at a standard velocity 

of Mach 25, but there are large differences in the technical means that 

have been studied for the thermal protection. During the 1960s, it was 

commonly expected that such craft would be built as hot structures. In 

fact, however, the thermal protection adopted for the Shuttle was the 

well-known “tiles,” a type of reusable insulation.

The Dyna-Soar program, early in the ’60s, was first to face this issue. 

Dyna-Soar used a radiatively cooled hot structure, with the primary or 

load-bearing structure being of René 41. Trusses formed the primary 

structure of the wings and fuselage, with many of their beams meeting at joints that were pinned rather than welded. Thermal gradients, 

23. Leo Steg, “Materials for Re-Entry Heat Protection of Satellites,” American Rocket Society Journal

(Sept. 1960), pp. 815–822.

24. Time (Nov. 18, 1957), pp. 19–20; (Apr. 27, 1959), p. 16; Joel W. Powell, “Thor-Able and 

Atlas-Able,” Journal of the British Interplanetary Society (May 1984), pp. 219–225; Kevin C. Ruffner, 

Corona: America’s First Satellite Program (McLean, VA: Central Intelligence Agency, 1995); see also 

T.A. Heppenheimer, “Toward Transatmospheric Flight,” (in this volume) for additional details.
Case 9 | High-Temperature Structures and Materials

583

Schematic drawing of the Boeing X-20A Dyna-Soar. USAF.

imposing differential expansion on separate beams, caused these members to rotate at the pins. This accommodated the gradients without 

imposing thermal stresses. René 41 was selected as a commercially 

available superalloy that had the best available combination of oxidation resistance and high-temperature strength. Its yield strength, 

130,000 pounds per square inch (psi) at room temperature, fell off only 

slightly at 1,200 °F and retained useful values at 1,800 °F. It could be processed as sheet, strip, wire, tubes, and forgings. Used as primary structure of Dyna-Soar, it supported a design specification that stated that 

the craft was to withstand at least four reentries under the most severe 

conditions permitted.

As an alloy, René 41 had a standard composition of 19 percent chromium, 11 percent cobalt, 10 percent molybdenum, 3 percent titanium, 

and 1.5 percent aluminum, along with 0.09 percent carbon and 0.006 

percent boron, with the balance being nickel. It gained strength through 

age hardening, with the titanium and aluminum precipitating within the 

nickel as an intermetallic compound. Age-hardening weldments initially 

showed susceptibility to cracking, which occurred in parts that had been 

strained through welding or cold working. A new heat-treatment process 

9
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9

permitted full aging without cracking, with the fabricated assemblies 

showing no significant tendency to develop cracks.25

As a structural material, the relatively mature state of René 41 

reflected the fact that it had already seen use in jet engines. It nevertheless lacked the temperature resistance necessary for use in the metallic 

shingles or panels that were to form the outer skin of the vehicle, which 

were to reradiate the heat while withstanding temperatures as high as 

3,000 °F. Here there was far less existing art, and investigators at Boeing 

had to find their way through a somewhat roundabout path. Four refractory or temperature-resistant metals initially stood out: tantalum, tungsten, molybdenum, and columbium. Tantalum was too heavy. Tungsten 

was not available commercially as sheet. Columbium also appeared to 

be ruled out, for it required an antioxidation coating, but vendors were 

unable to coat it without rendering it brittle. Molybdenum alloys also 

faced embrittlement because of recrystallization produced by a prolonged soak at high temperature in the course of coating formation. A 

promising alloy, Mo-0.5Ti, overcame this difficulty through addition of 

0.07 percent zirconium. The alloy that resulted, Mo-0.5Ti-0.07Zr, was 

called TZM molybdenum. For a time it appeared as a highly promising 

candidate for all the other panels.26

Wing design also promoted its use, for the craft mounted a delta 

wing with leading-edge sweep of 73 degrees. Though built for hypersonic entry from orbit, it resembled the supersonic delta wings 

of contemporary aircraft such as the B-58 bomber. But this wing 

was designed using H. Julian Allen’s blunt-body principle, with the 

leading edge being thickly rounded (that is, blunted) to reduce the rate 

of heating. The wing sweep then reduced equilibrium temperatures 

along the leading edge to levels compatible with the use of TZM.27

25. ASD staff, Proceedings of 1962 X-20A (Dyna-Soar) Symposium, vol. 3, Structures and Materials

(Wright Patterson AFB: USAF Aeronautical Systems Division, 1963), DTIC AD-346912, pp. III-3-1-2 

to -5, III-3-1-18 to -23, II-4-2-2 to -8. Re: René 41, see Howard J. Middendorf, “Materials and 

Processes for X-20A (Dyna-Soar),” Air Force Systems Command, June 1964, DTIC AD-449685, 

pp. 28, 41.

26. William Cowie, “Utilization of Refractory Metals on the X-20A (Dyna-Soar),” Air Force Systems Command, June 1964, DTIC: AD-609169; AD-449685, pp. 3–5, 20; AD-346912, pp. III-3-1-7, III-4-1-2 to -3.

27. DTIC AD-346912, p. III-3-1-8; Alvin Seiff and H. Julian Allen, “Some Aspects of the Design of 

Hypersonic Boost-Glide Aircraft,” NACA RM-A55E26 (1955); and Clarence Geiger, “Strangled 

Infant: The X-20 Dyna-Soar,” in Richard P. Hallion, ed., Hypersonic Revolution, vol. 1 (Bolling AFB: 

USAF, 1998).
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Boeing’s metallurgists nevertheless held an ongoing interest in columbium, because in uncoated form it showed superior ease of fabrication 

and lack of brittleness. A new Boeing-developed coating method eliminated embrittlement, putting columbium back in the running. A survey of its alloys showed that they all lacked the hot strength of TZM. 

Columbium nevertheless retained its attractiveness because it promised 

less weight. Based on coatability, oxidation resistance, and thermal emissivity, the preferred alloy was Cb-10Ti-5Zr, called D-36. It replaced TZM 

in many areas of the vehicle but proved to lack strength against creep at 

the highest temperatures. Moreover, coated TZM gave more of a margin against oxidation than coated D-36 did, again at the most extreme 

temperatures. D-36 indeed was chosen to cover most of the vehicle, 

including the flat underside of the wing. But TZM retained its advantage for such hot areas as the wing leading edges.28

The vehicle had some 140 running feet of leading edges and 140 

square feet of associated area. This included leading edges of the vertical fins and elevons as well as of the wings. In general, D-36 served when 

temperatures during reentry did not exceed 2,700 °F, while TZM was used 

for temperatures between 2,700 and 3,000 °F. In accordance with the 

Stefan-Boltzmann law, all surfaces radiated heat at a rate proportional 

to the fourth power of the temperature. Hence for equal emissivities, a 

surface at 3,000 °F radiated 44 percent more heat than one at 2,700 °F.29

Panels of both TZM and D-36 demanded antioxidation coatings. 

These coatings were formed directly on the surfaces as metallic silicides 

(silicon compounds), using a two-step process that employed iodine as 

a chemical intermediary. Boeing introduced a fluidized-bed method 

for application of the coatings that cut the time for preparation while 

enhancing uniformity and reliability. In addition, a thin layer of silicon 

carbide, applied to the surface, gave the vehicle its distinctive black color. 

It enhanced the emissivity, lowering temperatures by as much as 200 °F.

It was necessary to show that complete panels could withstand 

aerodynamic flutter. A report of the Aerospace Vehicles Panel of the 

Air Force Scientific Advisory Board came out in April 1962 and 

singled out the problem of flutter, citing it as one that called for critical 

attention. The test program used two NASA wind tunnels: the 4-foot by 4-foot 

Unitary facility at Langley that covered a range of Mach 1.6 to 2.8 

28. DTIC: AD-609169; AD-4496845, pp. 3–5, 20; AD-346912, pp. III-3-1-7, III-4-1-2 -3.

29. DTIC: AD-346912, p. III-3-1-8; AD-449685, p. iii.

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and the 11-foot by 11-foot Unitary installation at Ames for Mach 1.2 to 

1.4. Heaters warmed test samples to 840 °F as investigators started with 

steel panels and progressed to versions fabricated from René nickel alloy.

“Flutter testing in wind tunnels is inherently dangerous,” a Boeing 

review declared. “To carry the test to the actual flutter point is to risk 

destruction of the test specimen. Under such circumstances, the safety 

of the wind tunnel itself is jeopardized.” Panels under test were as large 

as 24 by 45 inches; flutter could have brought failure through fatigue, with 

parts of a specimen being blown through the tunnel at supersonic speed. 

Thus, the work started at dynamic pressures of 400 and 500 pounds per 

square foot (psf) and advanced over a year and a half to exceed the design 

requirement of close to 1,400 psf. Tests were concluded in 1962.30

Between the outer panels and the inner primary structure, a corrugated 

skin of René 41 served as the substructure. On the upper wing surface and 

upper fuselage, where the temperatures were no higher than 2,000 °F, the 

thermal-protection panels were also of René 41 rather than of a refractory. Measuring 12 by 45 inches, these panels were spot-welded directly to 

the corrugations of the substructure. For the wing undersurface and for 

other areas that were hotter than 2,000 °F, designers specified an insulated 

structure. Standoff clips, each with four legs, were riveted to the underlying 

corrugations and supported the refractory panels, which also were 12 

by 45 inches in size.

The space between the panels and the substructure was to be filled 

with insulation. A survey of candidate materials showed that most of 

them exhibited a strong tendency to shrink at high temperatures. This 

was undesirable; it increased the rate of heat transfer and could create 

uninsulated gaps at seams and corners. Q-felt, a silica fiber from Johns 

Manville, also showed shrinkage. However, nearly all of it occurred at 

2,000 °F and below; above 2,000 °F, further shrinkage was negligible. This 

meant that Q-felt could be “pre-shrunk” through exposure to temperatures above 2,000 °F for several hours. The insulation that resulted had 

density no greater than 6.2 pounds per cubic foot, one-tenth that of water. 

In addition, it withstood temperatures as high as 3,000 °F.31

30. DTIC AD-346912, pp. III-3-6-2 to -15 (quotes, pp. -8, -11); “Scientific Advisory Board Meme 

Report of the Aerospace Vehicles Panel on Dyna-Soar Panel Flutter,” Apr. 20, 1962. See also Heppenheimer, “Toward Transatmospheric Flight.”

31. DTIC: AD-449685, pp. 49–50; AD-609169, pp. 13, 29; AD-346912, pp. II-3-1-6 to -7, 

III-3-6-2, III-3-6-13 to -14.
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TZM outer panels, insulated with Q-felt, proved suitable for wing 

leading edges. These were designed to withstand equilibrium temperatures of 2,825 °F and short-duration over-temperatures of 2,900 °F. But 

the nose cap faced temperatures of 3,680 °F along with a peak heat flux 

of 143 BTU/ft2/sec. This cap had a radius of curvature of 7.5 inches, 

making it far less blunt than the contemporary Project Mercury heat 

shield that had a radius of 120 inches.32 Its heating was correspondingly 

more severe. Reliable thermal protection of the nose was essential, so 

the program conducted two independent development efforts that used 

separate technical approaches. The firm of Chance Vought pursued the 

main line of activity, while Boeing also devised its own nose cap design.

The work at Vought began with a survey of materials that paralleled 

Boeing’s review of refractory metals for the thermal-protection panels. 

Molybdenum and columbium had no strength to speak of at the pertinent temperatures, but tungsten retained useful strength even at 4,000 

°F. But that metal could not be welded, while no coating could protect 

it against oxidation. Attention then turned to nonmetallic materials, 

including ceramics.

Ceramics of interest existed as oxides such as silica and magnesia, 

which meant that they could not undergo further oxidation. Magnesia 

proved to be unsuitable because it had low thermal emittance, while 

silica lacked strength. However, carbon in the form of graphite showed 

clear promise. It held considerable industrial experience; it was light in 

weight, while its strength actually increased with temperature. It oxidized readily but could be protected up to 3,000 °F by treating it with 

silicon, in vacuum and at high temperatures, to form a thin protective 

layer of silicon carbide. Near the stagnation point, the temperatures 

during reentry would exceed that level. This brought the concept of a 

nose cap with siliconized graphite as the primary material and with an 

insulated layer of a temperature-resistant ceramic covering its forward 

area. With graphite having good properties as a heat sink, it would rise 

in temperature uniformly and relatively slowly, while remaining below 

the 3,000 °F limit throughout the full time of the reentry.

Suitable grades of graphite proved to be available commercially 

from the firm of National Carbon. Candidate insulators included hafnia, thoria, magnesia, ceria, yttria, beryllia, and zirconia. Thoria was 

32. DTIC AD-346912, pp. III-3-1-8, III-3-4-4; “How Mercury Capsule Design Evolved,” Aviation 

Week (Sept. 21, 1959), pp. 52–59.

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the most refractory but was very dense and showed poor resistance to 

thermal shock. Hafnia brought problems of availability and of reproducibility of properties. Zirconia stood out. Zirconium, its parent metal, 

had found use in nuclear reactors; the ceramic was available from the 

Zirconium Corporation of America. It had a melting point above 4,500 °F, 

was chemically stable and compatible with siliconized graphite, offered 

high emittance with low thermal conductivity, provided adequate resistance to thermal shock and thermal stress, and lent itself to fabrication.33

For developmental testing, Vought used two in-house facilities that 

simulated the flight environment, particularly during reentry. A ramjet, 

fueled with JP-4 and running with air from a wind tunnel, produced an 

exhaust with velocity up to 4,500 ft/sec and temperature up to 3,500 °F. 

It also generated acoustic levels above 170 decibels (dB), reproducing 

the roar of a Titan III booster and showing that samples under test could 

withstand the resulting stresses without cracking. A separate installation, built specifically for the Dyna-Soar program, used an array of propane burners to test full-size nose caps.

The final Vought design used a monolithic shell of siliconized graphite that was covered over its full surface by zirconia tiles held in place by 

thick zirconia pins. This arrangement relieved thermal stresses by permitting mechanical movement of the tiles. A heat shield stood behind 

the graphite, fabricated as a thick disk-shaped container made of coated 

TZM sheet metal and filled with Q-felt. The nose cap was attached to the 

vehicle with a forged ring and clamp that also were of coated TZM. The 

cap as a whole relied in radiative cooling. It was designed to be reusable; like the primary structure, it was to withstand four reentries under 

the most severe conditions permitted.34

The backup Boeing effort drew on that company’s own test equipment. Study of samples used the Plasma Jet Subsonic Splash Facility, 

which created a jet with temperature as high as 8,000 °F that splashed 

over the face of a test specimen. Full-scale nose caps went into the Rocket 

Test Chamber, which burned gasoline to produce a nozzle exit velocity 

of 5,800 ft/sec and an acoustic level of 154 dB. Both installations were 

capable of long-duration testing, reproducing conditions during reentries that could last for 30 minutes.35

33. DTIC AD-346912, pp. III-4-5-3 to -6, III-4-5-16.

34. Ibid., pp. III-4-5-13 to -18, III-4-5-35.

35. Ibid., pp. III-3-4-4 to -6, III-3-4-20.
Case 9 | High-Temperature Structures and Materials

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The Boeing concept used a monolithic zirconia nose cap that was 

reinforced against cracking with two screens of platinum-rhodium wire. 

The surface of the cap was grooved to relieve thermal stress. Like its 

counterpart from Vought, this design also installed a heat shield that 

used Q-felt insulation. However, there was no heat sink behind the zirconia cap. This cap alone provided thermal protection at the nose, through 

radiative cooling. Lacking pinned tiles and an inner shell, its design was 

simpler than that of Vought.36

Its fabrication bore comparison to the age-old work of potters, who 

shape wet clay on a rotating wheel and fire the resulting form in a kiln. 

Instead of using a potter’s wheel, Boeing technicians worked with a 

steel die with an interior in the shape of a bowl. A paper honeycomb, 

reinforced with Elmer’s Glue and laid in place, defined the pattern of 

stress-relieving grooves within the nose cap surface. The working material was not moist clay but a mix of zirconia powder with binders, internal lubricants, and wetting agents.

With the honeycomb in position against the inner face of the die, a 

specialist loaded the die by hand, filling the honeycomb with the damp 

mix and forming layers of mix that alternated with the wire screens. The 

finished layup, still in its die, went into a hydraulic press. A pressure of 

27,000 psi compacted the form, reducing its porosity for greater strength 

and less susceptibility to cracks. The cap was dried at 200 °F, removed 

from its die, dried further, and then fired at 3,300 °F for 10 hours. The 

paper honeycomb burned out in the course of the firing. Following visual 

and x-ray inspection, the finished zirconia cap was ready for machining to shape in the attachment area, where the TZM ring-and-clamp 

arrangement waste anchor it to the fuselage.37

The nose cap, outer panels, and primary structure all were built 

to limit their temperatures through passive methods: radiation and 

insulation. Active cooling also played a role, reducing temperatures 

within the pilot’s compartment and two equipment bays. These used 

a “water wall” that mounted absorbent material between sheet-metal 

panels to hold a mix of water and a gel. The gel retarded flow of this 

fluid, while the absorbent wicking kept it distributed uniformly to 

prevent hotspots.

36. Ibid., p. III-3-4-4; Geiger, “Strangled Infant,” p. 357. See Heppenheimer, “Toward Transatmospheric Flight.”

37. DTIC AD-346912, pp. III-4-6-5 to -7.

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During reentry, heat reached the water walls as it penetrated into the 

vehicle. Some of the moisture evaporated as steam, transferring heat to 

a set of redundant water-glycol loops that were cooled by liquid hydrogen from an onboard supply. A catalytic bed combined the stream of 

warmed hydrogen with oxygen that again came from an onboard supply. 

This produced gas that drove the turbine of Dyna-Soar’s auxiliary power 

unit, which provided both hydraulic and electric power to the craft.

A cooled hydraulic system also was necessary, to move the control surfaces as on a conventional airplane. The hydraulic fluid operating temperature was limited to 400 °F by using the fluid itself as an 

initial heat-transfer medium. It flowed through an intermediate waterglycol loop that removed its heat by being cooled with hydrogen. Major 

hydraulic components, including pumps, were mounted within an 

actively cooled compartment. Control-surface actuators, along with 

associated valves and plumbing, were insulated using inch-thick blankets of Q-felt. Through this combination of passive and active cooling methods, the Dyna-Soar program avoided a need to attempt to 

develop truly high-temperature arrangements, remaining instead within 

the state of the art.38

Specific vehicle parts and components brought their own thermal 

problems. Bearings, both ball and antifriction, needed strength to carry 

mechanical loads at high temperatures. For ball bearings, the cobaltbase superalloy Stellite 19 was known to be acceptable up to 1,200 °F. 

Investigation showed that it could perform under high load for short 

durations at 1,350 °F. Dyna-Soar nevertheless needed ball bearings qualified for 1,600 °F and obtained them as spheres of René 41 plated with 

gold. The vehicle also needed antifriction bearings as hinges for control 

surfaces, and here there was far less existing art. The best available bearings used stainless steel and were suitable only to 600 °F, whereas DynaSoar again faced a requirement of 1,600 °F. A survey of 35 candidate 

materials led to selection of titanium carbide with nickel as a binder.39

Antenna windows demanded transparency to radio waves at similarly high temperatures. A separate program of materials evaluation 

led to selection of alumina, with the best grade being available from the 

Coors Porcelain Company.40

38. Ibid., pp. 3-1-10 to -12; Geiger, “Strangled Infant,” pp. 361–368.

39. DTIC: AD-346912, pp. III-4-6-9 to -11; AD-449685, pp. 63–65.

40. DTIC: AD-346912, pp. III-4-6-7 to -8; AD-449685, pp. 57–59.
Case 9 | High-Temperature Structures and Materials

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NASA concepts for passive and actively cooled ablative heat shields, 1960. NASA.

The pilot needed his own windows. The three main ones, facing forward, were the largest yet planned for a piloted spacecraft. They had 

double panes of fused silica, with infrared-reflecting surfaces on all 

surfaces except the outermost. This inhibited the inward flow of heat 

by radiation, reducing the load on the active cooling of the pilot’s compartment. The window frames expanded when hot; to hold the panes in 

position, those frames were fitted with springs of René. The windows 

also needed thermal protection so they were covered with a shield of 

D-36. It was supposed to be jettisoned following reentry, around Mach 

5, but this raised a question: what if it remained attached? The cockpit had two other windows, one on each side, which faced a less severe 

environment and were to be left unshielded throughout a flight. Over 

a quarter century earlier, Charles Lindbergh had flown the Spirit of St. 

Louis across the North Atlantic from New York to Paris using just side 

vision and a crude periscope. But that was a far cry from a plummeting 

lifting reentry vehicle. Now, test pilot Neil Armstrong flew Dyna-Soar–

like approaches and landings in a modified Douglas F5D-1 fighter with 

side vision only and showed it was still possible.41

41. DTIC AD-346912, pp. III-3-1-10 to -11; Dennis Jenkins, Space Shuttle (Stillwater, MN: Voyageur Press, 2001).

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9

The vehicle was to touch down at 220 knots. It lacked wheeled 

landing gear, for inflated rubber tires would have demanded their own 

cooled compartments. For the same reason, it was not possible to use 

a conventional oil-filled strut as a shock absorber. The craft therefore 

deployed tricycle landing skids. The two main skids, from Goodyear, 

were of Waspalloy nickel steel and mounted wire bristles of René 41. 

These gave a high coefficient of friction, enabling the vehicle to skid 

to a stop in a planned length of 5,000 feet while accommodating runway irregularities. In place of the usual oleo strut, a long rod of Inconel 

stretched at the moment of touchdown and took up the energy of impact, 

thereby serving as a shock absorber. The nose skid, from Bendix, was 

forged from René 41 and had an undercoat of tungsten carbide to resist 

wear. Fitted with its own energy-absorbing Inconel rod, the front skid 

had a reduced coefficient of friction, which helped to keep the craft 

pointing straight ahead during slide-out.42

Through such means, the Dyna-Soar program took long strides 

toward establishing hot structures as a technology suitable for operational use during reentry from orbit. The X-15 had introduced heat sink 

fabricated from Inconel X, a nickel steel. Dyna-Soar went considerably 

further, developing radiation-cooled insulated structures fabricated from 

René 41 and from refractory materials. A chart from Boeing made the 

point that in 1958, prior to Dyna-Soar, the state of the art for advanced 

aircraft structures involved titanium and stainless steel, with temperature limits of 600 °F. The X-15 with its Inconel X could withstand temperatures above 1,200 °F. Against this background, Dyna-Soar brought 

substantial advances in the temperature limits of aircraft structures.43

Hot Structures: ASSET

Dyna-Soar never flew, for Defense Secretary Robert S. McNamara canceled the program in December 1963. At that time, vehicles were well 

under construction but still were some 21⁄2 years away from first flight.

42. DTIC AD-346912, pp. III-3-1-14 to -15; Geiger, “Strangled Infant,” pp. 347–349, 360–361.

43. See Table 1 in Heppenheimer, “Toward Transatmospheric Flight,” for specific advances by 

structural element. See also Terry L. Sunday and John R. London, “The X-20 Space Plane: Past 

Innovation, Future Vision,” in John Becklake, ed., History of Rocketry and Astronautics, vol. 17 of 

the AAS History Series (San Diego: Univelt, 1995), pp. 253–284; Geiger, “Strangled Infant,” pp. 

344–346; R.L. Schleicher, “Structural Design of the X-15,” Journal of the Royal Aeronautical Society

(Oct. 1963), pp. 618–636.
Case 9 | High-Temperature Structures and Materials

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Geometric planform of the McDonnell ASSET reentry test vehicle. USAF.

 Still its technology remained available for further development, and 

44. Hallion, Hypersonic Revolution, vol. 1, p. II-xvi (intro to Geiger case study on X-20).

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9

body. It had a length of 59 inches and a span of 55 inches. Its bill of 

materials resembled that of Dyna-Soar, for it used TZM molybdenum to 

withstand 3,000 °F on the forward lower heat shield, graphite for similar temperatures on leading edges, and zirconia rods for the nose cap, 

which was rated at 4,000 °F. But ASSET avoided the use of René 41, 

with cobalt and columbium alloys being employed instead.45

ASSET was built in two varieties: the Aerothermodynamic Structural 

Vehicle (ASV) weighing 1,130 pounds and the Aerothermodynamic 

Elastic Vehicle (AEV) at 1,225 pounds. The AEVs were to study panel 

flutter along with the behavior of a trailing-edge flap, which represented 

an aerodynamic control surface in hypersonic flight. These vehicles did 

not demand the highest possible flight speeds and therefore flew with 

single-stage Thors as the booster. But the ASVs were built to study materials and structures in the reentry environment while taking data on temperatures, pressures, and heat fluxes. Such missions demanded higher 

speeds. These boost-glide craft therefore used the two-stage Thor-Delta 

launch vehicle, which resembled the Thor-Able that had conducted nose 

cone tests at intercontinental range in 1958.46

The program eventually conducted six flights:47 several of these craft 

were to be recovered. Following standard practice, their launches were 

scheduled for the early morning, to give downrange recovery crews the 

maximum hours of daylight. That did not help ASV-1, the first flight in 

the program, which sank into the sea. Still, it flew successfully and returned 

good data. In addition, this flight set a milestone, for it was the first time in 

aviation history that a lifting reentry spacecraft had traversed the demanding hypersonic reentry corridor from orbit down to the lower atmosphere.48

ASV-2 followed, using the two-stage Thor-Delta, but it failed when 

the second stage did not ignite. The next one carried ASV-3, with this 

mission scoring a double achievement. It not only made a good flight 

downrange, but it was also successfully recovered. It carried a liquidcooled double-wall test panel from Bell Aircraft along with a molybdenum heat-shield panel from Boeing, home of Dyna-Soar. ASV-3 also had a 

45. “Advanced Technology Program: Technical Development Plan for Aerothermodynamic/Elastic 

Structural System Environmental Tests (ASSET),” Air Force Systems Command, Sept. 9, 1963.

46. Ibid., pp. 4, 11–13; Hallion, “ASSET,” in Hypersonic Revolution, vol. 1, pp. 451, 464–465.

47. See Table 2 in Heppenheimer, “Toward Transatmospheric Flight,” for a list of these flights and 

performance objectives.

48. Hallion, “ASSET,” pp. 510–512 (quote, p. 512).
Case 9 | High-Temperature Structures and Materials

595

new nose cap. The standard ASSET type used zirconia dowels, l.5 inches 

long by 0.5 inches in diameter, which were bonded together with a zirconia cement. The new cap, from International Harvester, had a tungsten base covered with thorium oxide and was reinforced with tungsten.

A company advertisement stated that it withstood reentry so well 

that it “could have been used again,” and this was true for the craft as a 

whole. Historian Richard P. Hallion writes that “overall, it was in excellent condition. Water damage . . . caused some problems, but not so serious that McDonnell could not have refurbished and reflown the vehicle.” 

The Boeing and Bell panels came through reentry without damage, and 

the importance of physical recovery was emphasized when columbium 

aft leading edges showed significant deterioration. They were redesigned, 

with the new versions going into subsequent AEV and ASV spacecraft.49

The next two flights were AEVs, each of which carried a flutter test 

panel and a test flap. AEV-1 returned only one high-Mach data point, at 

Mach 11.88, but this sufficed to indicate that its panel was probably too 

stiff to undergo flutter. Engineers made it thinner and flew a new one on 

AEV-2, where it returned good data until it failed at Mach 10. The flap 

experiment also showed value. It had an electric motor that deflected it 

into the airstream, with potentiometers measuring the force required to 

move it, and it enabled aerodynamicists to critique their theories. Thus 

one treatment gave pressures that were in good agreement with observations, whereas another did not.

ASV-4, the final flight, returned “the highest quality data of the ASSET 

program,” according to the flight-test report. The peak speed of 19,400 ft/sec, 

Mach 18.4, was the highest in the series and was well above the design speed 

of 18,000 ft/sec. The long hypersonic glide covered 2,300 nautical miles and 

prolonged the data return, which presented pressures at 29 locations on the 

vehicle and temperatures at 39. An onboard system transferred mercury ballast to trim the angle of attack, increasing the lift-to-drag ratio (L/D) from its 

average of 1.2 to 1.4, and extending the trajectory. The only important problem came when the recovery parachute failed to deploy properly and ripped 

away, dooming ASV-4 to follow ASV-1 into the depths of the Atlantic.50

49. Ibid., pp. 512–516 (quote, p. 515); “ASSET ASV-3 Flight Test Report,” McDonnell Aircraft 

Corp., Jan. 4, 1965, DTIC AD-357523; advertisement, Aviation Week, May 24, 1965, p. 62.

50. “Fourth ASSET Glider Gathers Data,” Aviation Week, Nov. 2, 1964, pp. 25–26; “ASSET ASV-4 

Flight Test Report,” McDonnell Aircraft Corp., June 25, 1965, DTIC AD-366546; Hallion, “ASSET,” 

pp. 516–519.

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NASA concept for a hypersonic cruise wing structure formed of beaded, corrugated, and tubular structural panels, 1978. NASA.

On the whole, ASSET nevertheless scored a host of successes. It 

showed that insulated hot structures could be built and flown without 

producing unpleasant surprises, at speeds up to three-fourths of orbital 

velocity. It dealt with such practical issues of design as fabrication, fasteners, and coatings. In hypersonic aerodynamics, ASSET contributed to 

understanding of flutter and of the use of movable control surfaces. The 

program also developed and successfully used a reaction control system 

built for a lifting reentry vehicle. Only one flight vehicle was recovered 

in four attempts, but it complemented the returned data by permitting 

a close look at a hot structure that had survived its trial by fire.

Reusable Surface Insulation

Early in the 1960s, researchers at Lockheed introduced an entirely different approach to thermal protection, which in time became the standard. Ablatives were unrivalled for once-only use, but during that decade 

the hot structure continued to stand out as the preferred approach for 

reusable craft such as Dyna-Soar. As noted, it used an insulated primary 

or load-bearing structure with a skin of outer panels. These emitted 

heat by radiation, maintaining a temperature that was high but steady. 
Case 9 | High-Temperature Structures and Materials

597

Strength versus temperature for various superalloys, including René 41, the primary structural 

material used on the X-20 Dyna-Soar. NASA.

Metal fittings supported these panels, and while the insulation could be 

high in quality, these fittings unavoidably leaked heat to the underlying 

structure. This raised difficulties in crafting this structure of aluminum 

and even of titanium, which had greater heat resistance. On Dyna-Soar, 

only René 41 would do.51

Ablatives avoided such heat leaks while being sufficiently capable as 

insulators to permit the use of aluminum. In principle, a third approach 

combined the best features of hot structures and ablatives. It called for 

the use of temperature-resistant tiles, made perhaps of ceramic, which 

could cover the vehicle skin. Like hot-structure panels, they would radiate heat, while remaining cool enough to avoid thermal damage. In 

addition, they were to be reusable. They also were to offer the excellent 

insulating properties of good ablators, preventing heat from reaching 

the underlying structure—which once more might be of aluminum. This 

concept became known as reusable surface insulation (RSI). In time, it 

gave rise to the thermal protection of the Shuttle.

51. F.M. Anthony, R.R. Fisher, and R.G. Helenbrook, “Selection of Space Shuttle Thermal Protection 

Systems,” AIAA Paper 71-443 (1971).

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RSI grew out of ongoing work with ceramics for thermal protection. Ceramics had excellent temperature resistance, light weight, and 

good insulating properties. But they were brittle and cracked rather than 

stretched in response to the flexing under load of an underlying metal 

primary structure. Ceramics also were sensitive to thermal shock, as 

when heated glass breaks when plunged into cold water. In flight, such 

thermal shock resulted from rapid temperature changes during reentry.52

Monolithic blocks of the ceramic zirconia had been specified for the 

nose cap of Dyna-Soar, but a different point of departure used mats of 

solid fiber in lieu of the solid blocks. The background to the Shuttle’s tiles 

lay in work with such mats that took place early in the 1960s at Lockheed 

Missiles and Space Company. Key people included R.M. Beasley, Ronald 

Banas, Douglas Izu, and Wilson Schramm. A Lockheed patent disclosure of December 1960 gave the first presentation of a reusable insulation made of ceramic fibers for use as a heat shield. Initial research dealt 

with casting fibrous layers from a slurry and bonding the fibers together.

Related work involved filament-wound structures that used long 

continuous strands. Silica fibers showed promise and led to an early 

success: a conical radome of 32-inch diameter built for Apollo in 1962. 

Designed for reentry, it had a filament-wound external shell and a lightweight layer of internal insulation cast from short fibers of silica. The 

two sections were densified with a colloid of silica particles and sintered 

into a composite. This gave a nonablative structure of silica composite 

reinforced with fiber. It never flew, as design requirements changed during the development of Apollo. Even so, it introduced silica fiber into 

the realm of reentry design.

Another early research effort, Lockheat, fabricated test versions of 

fibrous mats that had controlled porosity and microstructure. These 

were impregnated with organic fillers such as Plexiglas (methyl methacrylate). These composites resembled ablative materials, though the 

filler did not char. Instead it evaporated or volatilized, producing an 

outward flow of cool gas that protected the heat shield at high heattransfer rates. The Lockheat studies investigated a range of fibers that 

included silica, alumina, and boria. Researchers constructed multilayer 

composite structures of filament-wound and short-fiber materials that 

resembled the Apollo radome. Impregnated densities were 40 to 60 lb/

52. Wilson B. Schramm, Ronald P. Banas, and Y. Douglas Izu, “Space Shuttle Tile—The Early 

Lockheed Years,” Lockheed Horizons, No. 13, 1983, pp. 2–15.
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ft3, the higher density being close to that of water. Thicknesses of no 

more than an inch gave acceptably low back-face temperatures during 

simulations of reentry.

This work with silica-fiber ceramics was well underway during 1962. 

Three years later, a specific formulation of bonded silica fibers was ready 

for further development. Known as LI-1500, it was 89 percent porous 

and had a density of 15 lb/ft3, one-fourth that of water. Its external surface was impregnated with filler to a predetermined depth, again to 

provide additional protection during the most severe reentry heating. 

By the time this filler was depleted, the heat shield was to have entered 

a zone of more moderate heating, where the fibrous insulation alone 

could provide protection.

Initial versions of LI-1500, with impregnant, were intended for use 

with small space vehicles similar to Dyna-Soar that had high heating 

rates. Space Shuttle concepts were already attracting attention—the 

January 1964 issue of Astronautics & Aeronautics, the journal of the 

American Institute of Aeronautics and Astronautics, presents the thinking of the day—and in 1965 a Lockheed specialist, Maxwell Hunter, 

introduced an influential configuration called Star Clipper. His design 

employed LI-1500 for thermal protection.

Like other Shuttle concepts, Star Clipper was to fly repeatedly, but 

the need for an impregnant in LI-1500 compromised its reusability. But 

in contrast to earlier entry vehicle concepts, Star Clipper was large, offering exposed surfaces that were sufficiently blunt to benefit from H. Julian 

Allen’s blunt-body principle. They had lower temperatures and heating 

rates, which made it possible to dispense with the impregnant. An unfilled 

version of LI-1500, which was inherently reusable, now could serve.

Here was the first concept of a flight vehicle with reusable insulation, bonded to the skin, which could reradiate heat in the fashion of 

a hot structure. However, the matted silica by itself was white and had 

low thermal emissivity, making it a poor radiator of heat. This brought 

excessive surface temperatures that called for thick layers of the silica 

insulation, adding weight. To reduce the temperatures and the thickness, the silica needed a coating that could turn it black for high emissivity. It then would radiate well and remain cooler.

The selected coating was a borosilicate glass, initially with an admixture of Cr2O3 and later with silicon carbide, which further raised the 

emissivity. The glass coating and the silica substrate were both silicon 

dioxide; this assured a match of their coefficients of thermal expansion, 

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to prevent the coating from developing cracks under the temperature 

changes of reentry. The glass coating could soften at very high temperatures to heal minor nicks or scratches. It also offered true reusability, 

surviving repeated cycles to 2,500 °F. A flight test came in 1968 as NASA 

Langley investigators mounted a panel of LI-1500 to a Pacemaker reentry test vehicle along with several candidate ablators. This vehicle carried instruments and was recovered. Its trajectory reproduced the peak 

heating rates and temperatures of a reentering Star Clipper. The LI-1500 

test panel reached 2,300 °F and did not crack, melt, or shrink. This proofof-concept test gave further support to the concept of high-emittance 

reradiative tiles of coated silica for thermal protection.53

Lockheed conducted further studies at its Palo Alto Research Center. 

Investigators cut the weight of RSI by raising its porosity from the 89 

percent of LI-1500 to 93 percent. The material that resulted, LI-900, 

weighed only 9 pounds per cubic foot, one-seventh the density of water.54 

There also was much fundamental work on materials. Silica exists in 

three crystalline forms: quartz, cristobalite, and tridymite. These not 

only have high coefficients of thermal expansion but also show sudden expansion or contraction with temperature because of solid-state 

phase changes. Cristobalite is particularly noteworthy; above 400 °F, 

it expands by more than 1 percent as it transforms from one phase to 

another. Silica fibers for RSI were to be glass, an amorphous rather than 

a crystalline state with a very low coefficient of thermal expansion and 

an absence of phase changes. The glassy form thus offered superb resistance to thermal stress and thermal shock, which would recur repeatedly during each return from orbit.55

The raw silica fiber came from Johns Manville, which produced it 

from high-purity sand. At elevated temperatures, it tended to undergo 

“devitrification,” transforming from a glass into a crystalline state. Then, 

when cooling, it passed through phase-change temperatures and the 

fiber suddenly shrank, producing large internal tensile stresses. Some 

53. Ibid.

54. Korb, et al., “The Shuttle Orbiter”; Wilson Schramm, “HRSI and LRSI—The Early Years,” American Ceramic Society Bulletin, vol. 60 (1981), pp. 1194–1195.

55. Korb, et al., “The Shuttle Orbiter”; NASA, NASA Space Shuttle Technology Conference, vol. 

2, “Structures and Materials,” NASA TM-X-2273 (1971); Richard C. Thuss, Harry G. Thibault, and 

Arnold Hiltz, “The Utilization of Silica Based Thermal Insulation for the Space Shuttle Thermal Protection System,” CASI 72A-10764, Oct. 1971.
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fibers broke, giving rise to internal cracking within the RSI and degradation of its properties. These problems threatened to grow worse during subsequent cycles of reentry heating.

To prevent devitrification, Lockheed worked to remove impurities 

from the raw fiber. Company specialists raised the purity of the silica to 

99.9 percent while reducing contaminating alkalis to as low as 6 parts 

per million. Lockheed proceeded to do these things not only in the laboratory but also in a pilot plant. This plant took the silica from raw 

material to finished tile, applying 140 process controls along the way. 

Established in 1970, the pilot plant was expanded in 1971 to attain a true 

manufacturing capability. Within this facility, Lockheed produced tiles 

of LI-1500 and LI-900 for use in extensive programs of test and evaluation. In turn, the increasing availability of these tiles encouraged their 

selection for Shuttle protection in lieu of a hot-structure approach.56

General Electric (GE) also became actively involved, studying types 

of RSI made from zirconia and from mullite, 3Al2O3+2SiO2, as well 

as from silica. The raw fibers were commercial grade, with the zirconia 

coming from Union Carbide and the mullite from Babcock and Wilcox. 

Devitrification was a problem, but whereas Lockheed had addressed 

it by purifying its fiber, GE took the raw silica from Johns Manville 

and tried to use it with little change. The basic fiber, the Q-felt of DynaSoar, also had served as insulation on the X-15. It contained 19 different 

elements as impurities. Some were present at a few parts per million, 

but others—aluminum, calcium, copper, lead, magnesium, potassium, 

sodium—ran from 100 to 1,000 parts per million. In total, up to 0.3 

percent was impurity.

General Electric treated this fiber with a silicone resin that served 

as a binder, pyrolyzing the resin and causing it to break down at high 

temperatures. This transformed the fiber into a composite, sheathing 

each strand with a layer of amorphous silica that had a purity of 99.98 

percent or more. This high purity resulted from that of the resin. The 

amorphous silica bound the fibers together while inhibiting their devitrification. General Electric’s RSI had a density of 11.5 lb/ft3

, midway 

between that of LI-900 and LI-1500.57

Many Shuttle managers had supported hot structures, but by mid1971 they were in trouble. In Washington, the Office of Management 

56. Schramm, et al., “Space Shuttle Tile”; Schramm, “HRSI.”

57. CASI 72A-10764; NASA TM-X-2273, pp. 39–93.

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and Budget (OMB) now was making it clear that it expected to impose 

stringent limits on funding for the Shuttle, which brought a demand 

for new configurations that could cut the cost of development. Within 

weeks, the contractors did a major turnabout. They abandoned hot structures and embraced RSI. Managers were aware that it might take time 

to develop for operational use, but they were prepared to use ablatives 

for interim thermal protection and to switch to RSI once it was ready.58

What brought this dramatic change? The advent of RSI production 

at Lockheed was critical. This drew attention from Max Faget, a longtime NACA–NASA leader who had kept his hand in the field of Shuttle 

design, offering a succession of conceptual design configurations that 

had helped to guide the work of the contractors. His most important 

concept, designated MSC-040, came out in September 1971 and served 

as a point of reference. It used RSI and proposed to build the Shuttle of 

aluminum rather than René 41 or anything similar.59

Why aluminum? “My history has always been to take the most conservative approach,” Faget explained subsequently. Everyone knew how 

to work with aluminum, for it was the most familiar of materials, but 

everything else carried large question marks. Titanium, for one, was literally a black art. Much of the pertinent shop-floor experience had been 

gained within the SR-71 program and was classified. Few machine shops 

had pertinent background, for only Lockheed had constructed an airplane—the SR-71—that used titanium hot structure. The situation was 

worse for columbium and the superalloys, for these metals had been used 

mostly in turbine blades. Lockheed had encountered serious difficulties as its machinists and metallurgists wrestled with titanium. With the 

Shuttle facing the OMB’s cost constraints, no one cared to risk an overrun 

while machinists struggled with the problems of other new materials.60

NASA Langley had worked to build a columbium heat shield for the 

Shuttle and had gained a particularly clear view of its difficulties. It was 

heavier than RSI but offered no advantage in temperature resistance. 

58. “Space Shuttle Program Definition: Phase B Extension Final Report,” Grumman B35-43 RP-33 (Mar. 

15, 1972); North American Rockwell 1971 Reports SV 71-50 and SV 71-59 (1971); “Interim Report 

to OMSF: Phase B System Study Extension,” McDonnell-Douglas report (Sept. 1, 1971), p. 36.

59. Heppenheimer, The Space Shuttle Decision, NASA SP-4221 (Washington, DC: NASA, 

1999); Jenkins, Space Shuttle, pp. 141–150.

60. Heppenheimer, Turbulent Skies: The History of Commercial Aviation (New York: John Wiley, 

1995); author interview with Max Faget, Mar. 4, 1997.
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In addition, coatings posed serious problems. Silicides showed promise 

of reusability and long life, but they were fragile and easily damaged. A 

localized loss of coating could result in rapid oxygen embrittlement at 

high temperatures. Unprotected columbium oxidized readily, and above 

the melting point of its oxide, 2,730 °F, it could burst into flame.61 “The 

least little scratch in the coating, the shingle would be destroyed during reentry,” Faget said. Charles Donlan, the Shuttle Program Manager 

at NASA Headquarters, placed this in a broader perspective in 1983:

Phase B was the first really extensive effort to put together 

studies related to the completely reusable vehicle. As we went 

along, it became increasingly evident that there were some problems. And then as we looked at the development problems, they 

became pretty expensive. We learned also that the metallic heat 

shield, of which the wings were to be made, was by no means 

ready for use. The slightest scratch and you are in trouble.62

Other refractory metals offered alternatives to columbium, but even 

when proposing to use them, the complexity of a hot structure also militated against its selection. As a mechanical installation, it called for 

large numbers of clips, brackets, standoffs, frames, beams, and fasteners. Structural analysis loomed as a formidable task. Each of many panel 

geometries needed its own analysis, to show with confidence that the 

panels would not fail through creep, buckling, flutter, or stress under 

load. Yet this confidence might be fragile, for hot structures had limited ability to resist over-temperatures. They also faced the continuing 

issue of sealing panel edges against ingestion of hot gas during reentry.63

In this fashion, having taken a long look at hot structures, NASA did 

an about-face as it turned toward the RSI that Lockheed’s Max Hunter had 

recommended as early as 1965. Then, in January 1972, President Richard 

Nixon gave his approval to the Space Shuttle program, thereby raising it to 

the level of a Presidential initiative. Within days, NASA’s Dale Myers spoke 

61. L.J. Korb and H.M. Clancy, “The Shuttle Thermal Protection System,” CASI 81A-44344 (1981), 

pp. 232–249; Korb, et al., “The Shuttle Orbiter,” p. 1189.

62. John Mauer interview of Charles Donlan, Oct. 19, 1983, Oral History Series, Shuttle Interviews, Box 1, JSC History Collection, University of Houston-Clear Lake; author interview of Faget, 

Mar. 4, 1997.

63. CASI 81A-44344, Ref. 62.

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to a conference in Houston and stated that the Agency had made the basic 

decision to use RSI. Requests for proposal soon went out, inviting leading 

aerospace corporations to bid for the prime contract on the Shuttle orbiter, 

and North American won this $2.6-billion prize in July. However, the RSI 

wasn’t Lockheed’s. The proposal specified mullite RSI for the undersurface and forward fuselage, a design feature that had been held over from 

the company’s studies of a fully reusable orbiter during the previous year.64

Still, was mullite RSI truly the one to choose? It came from General 

Electric and had lower emissivity than the silica RSI of Lockheed but 

could withstand higher temperatures. Yet the true basis for selection lay 

in the ability to withstand 100 reentries as simulated in ground test. NASA 

conducted these tests during the last 5 months of 1972, using facilities 

at its Ames, Johnson, and Kennedy Centers, with support from Battelle 

Memorial Institute.

The main series of tests ran from August to November and gave 

a clear advantage to Lockheed. That firm’s LI-900 and LI-1500 went 

through 100 cycles to 2,300 °F and met specified requirements for maintenance of low back-face temperature and minimal thermal conductivity. The mullite showed excessive back-face temperatures and higher 

thermal conductivity, particularly at elevated temperatures. As test conditions increased in severity, the mullite also developed coating cracks 

and gave indications of substrate failure.

The tests then introduced acoustic loads, with each cycle of the simulation now subjecting the RSI to loud roars of rocket flight along with 

the heating of reentry. LI-1500 continued to show promise. By midNovember, it demonstrated the equivalent of 20 cycles to 160 decibels, 

the acoustic level of a large launch vehicle, and 2,300 °F. A month later, 

NASA conducted what Lockheed describes as a “sudden death shootout”: a new series of thermal-acoustic tests, in which the contending 

materials went into a single large 24-tile array at NASA Johnson. After 

20 cycles, only Lockheed’s LI-900 and LI-1500 remained intact. In separate tests, LI-1500 withstood 100 cycles to 2,500 °F and survived a thermal overshoot to 3,000 °F, as well as an acoustic overshoot to 174 dB. 

Clearly, this was the material NASA wanted.65

64. “Sortie Module May Cut Experiment Cost,” Aviation Week (Jan. 17, 1972), p. 17; North 

American Rockwell Report SV 72-19, 1972.

65. Schramm, et al., “Space Shuttle Tile,” Ref. 53, pp. 11–14; William S. Hieronymus, “Two Reusable Materials Studied for Orbiter Thermal Protection,” Aviation Week (Mar. 27, 1972), p. 48.
Case 9 | High-Temperature Structures and Materials

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Thermal protection system for the proposed National Hypersonic Flight Research Facility, 

1978. NASA.

As insulation, the tiles were astonishing. A researcher could heat 

one in a furnace until it was white hot, remove it, allow its surface 

to cool for a couple of minutes, and pick it up at its edges using his 

or her fingers, with its interior still at white heat. Lockheed won the 

thermal-protection subcontract in 1973, with NASA specifying LI-900 

as the baseline RSI. The firm responded with preparations for a fullscale production facility in Sunnyvale, CA. With this, tiles entered the 

mainstream of thermal protection.

Cooling

Hypersonics has much to say about heating, so it is no surprise that 

it also has something to say about cooling. Active cooling merits only 

slight attention, as in the earlier discussion of Dyna-Soar. Indeed, two 

books on Shuttle technology run for hundreds of pages and give complete treatments of tiles for thermal protection—but give not a word 

about active cooling.66 

66. Heppenheimer, Development of the Space Shuttle, 1972–1981, (Washington: Smithsonian 

Institution Press, 2002); Jenkins, Space Shuttle.

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9

 What the topic of cooling mostly comprises is the use of passive 

67. NASA SP-4221, pp. 335–340, 348–349; Grumman Report B35-43 RP-33; Michael L. 

Yaffee, “Program Changes Boost Grumman Shuttle,” Aviation Week (July 12, 1963), pp. 36–39.

68. “External LH2 Tank Study Final Report,” McDonnell-Douglas Report MDC E076-1, June 30, 1971.

69. NASA SP-4221, pp. 420–422; Heppenheimer, Development, p. 188.
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PRIME complemented ASSET, with both programs conducting flight 

tests of boost-glide vehicles. However, while ASSET pushed the state of 

the art in materials and hot structures, PRIME used ablative thermal 

protection for a more straightforward design and emphasized flight 

performance. Accelerated to near-orbital velocities by Atlas launch 

vehicles, the PRIME missions called for boost-glide flight from 

Vandenberg Air Force Base (AFB) to locations in the western Pacific 

near Kwajalein Atoll. The SV-5D had higher L/D than Gemini or Apollo 

did, and, as with those NASA programs, it was to demonstrate precision reentry. The plans called for cross range, with the vehicle flying up 

to 710 nautical miles to the side of a ballistic trajectory and then arriving within 10 miles of its recovery point.

The piloted X-24A supersonic lifting body, used to assess the SV-5 

shape’s approach and landing characteristics, was built of aluminum. 

The SV-5D also used this material for both its skin and primary structure. It mounted both aerodynamic and reaction controls, the former 

consisting of right and left body-mounted flaps set well aft. Deflected 

symmetrically, they controlled pitch; deflected individually (asymmetrically), they produced yaw and roll. These flaps were beryllium plates 

that provided a useful thermal heat sink. The fins were of steel honeycomb, likewise with surfaces of beryllium sheet.

Most of the vehicle surface obtained thermal protection from ESA 

3560 HF, a flexible ablative blanket of phenolic fiberglass honeycomb 

that used a silicone elastomer as the filler, with fibers of nylon and 

silica holding the ablative char in place during reentry. ESA 5500 HF, a 

high-density form of this ablator, gave added protection in hotter areas. 

The nose cap and the beryllium flaps used a different material: carbonphenolic composite. At the nose, its thickness reached 3.5 inches.70

The PRIME program made three flights that took place between 

December 1966 and April 1967. All returned data successfully, with 

the third flight vehicle also being recovered. The first mission reached 

25,300 ft/sec and flew 4,300 miles downrange, missing its target by only 

900 feet. The vehicle executed pitch maneuvers but made no attempt at 

cross range. The next two flights indeed achieved cross range, respec70. Miller, X-Planes, pp. 256–261; “SV-5D PRIME Final Flight Test Summary,” Martin Marietta Report 

ER-14465, Sept. 1967; William J. Normyle, “Manned Flight Tests to Seek Lifting-Body Technology,” 

Aviation Week (May 16, 1966), pp. 64–75; John L. Vitelli and Richard P. Hallion, “Project PRIME: 

Hypersonic Reentry from Space,” in Hallion, ed., Hypersonic Revolution, vol. 1, pp. 641, 648–649.

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Schematic of low– and high-temperature reusable surface insulation tiles, showing how they 

were bonded to the skin of the Space Shuttle. NASA.

tively of 500 and 800 miles, and the precision again was impressive. 

Flight 2 missed its aim point by less than 2 miles. Flight 3 missed by 

over 4 miles, but this still was within the allowed limit. Moreover, the 

terminal guidance radar had been inoperative, which probably contributed to the lack of absolute accuracy.71

A few years later, the Space Shuttle brought the question of whether 

its primary structure and skin should perhaps be built of titanium. 

Titanium offered a potential advantage because of its temperature resistance; hence, its thermal protection might be lighter. But the apparent 

weight saving was largely lost because of a need for extra insulation to 

protect the crew cabin, payload bay, and onboard systems. Aluminum 

could compensate for its lack of heat resistance because it had higher 

71. Martin Marietta Report ER-14465, p. I-1; Vitelli and Hallion, “Project PRIME,” in Hallion, ed., 

Hypersonic Revolution, vol. 1, pp. 694–702; B.K. Thomas, “USAF Nears Lifting Body Tests,” Aviation Week (July 10, 1967), pp. 99–101.
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thermal conductivity than titanium. It therefore could more readily 

spread its heat throughout the entire volume of the primary structure.

Designers expected to install RSI tiles by bonding them to the skin, 

and for this aluminum had a strong advantage. Both metals form thin layers of oxide when exposed to air, but that of aluminum is more strongly 

bound. Adhesive, applied to aluminum, therefore held tightly. The bond 

with titanium was considerably weaker and appeared likely to fail in 

operational use at around 500 °F. This was not much higher than the 

limit for aluminum, 350 °F, which showed that the temperature resistance of titanium did not lend itself to operational use.72

Turbine Blades

Turbine blades operate at speeds well below hypersonic, but this topic 

shares the same exotic metals that are used for flight structures at the 

highest speeds. It is necessary to consider how such blades use coatings to stay cool, an issue that represents another form of cooling. It 

also is necessary to consider directionally solidified and single-crystal 

castings for blades.

Britain’s firm of Rolls-Royce has traditionally possessed a strong 

standing in this field, and The Economist has noted its activity:

The best place to start is the surprisingly small, almost underwhelming, turbine blades that make up the heart of the giant 

engines slung beneath the wings of the world’s biggest planes. 

These are not the huge fan blades you see when boarding, but 

are buried deep in the engines. Each turbine blade can fit in 

the hand like an oversized steak knife. At first glance it may not 

seem much more difficult to make. Yet they cost about $10,000 

each. Rolls-Royce’s executives like to point out that their big 

engines, of almost six tonnes, are worth their weight in silver—

and that the average car is worth its weight in hamburger.73

Turbine blades are difficult to make because they have to survive high 

temperatures and huge stresses. The air inside big jet engines reaches 

about 2,900 °F in places, 750 degrees hotter than the melting point of 

72. Scott Pace, “Engineering Design and Political Choice: The Space Shuttle 1969-72,” master’s 

thesis (Cambridge: Massachusetts Institute of Technology, May 1982), pp. 179–188.

73. “Briefing: Rolls-Royce,” The Economist (Jan. 10, 2009), pp. 60–62.

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the metal from which the turbine blades are made. Each blade is grown 

from a single crystal of alloy for strength and then coated with tough 

ceramics. A network of tiny air holes then creates a thin blanket of cool 

air that stops it from melting.

The study of turbine blades brings in the topic of thermal barrier 

coatings (TBC). By attaching an adherent layer of a material of low thermal conductivity to the surface of an internally cooled turbine blade, a 

temperature drop is induced across the thickness of the layer. This results 

in a drop in the temperature of the metal blade. Using this approach, 

temperature reductions of up to 340 °F at the metal surface have been 

estimated for 150-micron-thick yttria stabilized zirconia coatings. The 

rest of the temperature decrease is obtained by cooling the blade using 

air from the compressor that is ducted downstream to the turbine.

The cited temperature reductions reduce the oxidation rate of the 

bond coat applied to the blades and so delay failure by oxidation. They 

also retard the onset of thermal fatigue. One should note that such coatings are currently used only to extend the life of components. They are 

not used to increase the operating temperature of the engine.

Modern TBCs are required to not only limit heat transfer through 

the coating but to also protect engine components from oxidation and 

hot corrosion. No single coating composition appears able to satisfy 

these requirements. As a result, a “coating system” has evolved. Research 

in the last 20 years has led to a preferred coating system consisting 

of four separate layers to achieve long-term effectiveness in the hightemperature, oxidative, and corrosive environment in which the blades 

must function. At the bottom is the substrate, a nickel – or cobalt-based 

superalloy that is cooled from the inside using compressor air. Overlaying 

it is the bond coat, an oxidation-resistant layer with thickness of 75–150 

microns that is typically of a NiCrAlY or NiCoCrAlY alloy. It essentially 

dictates the spallation failure of the blade. Though it resists oxidation, it 

does not avoid it; oxidation of this coating forms a third layer, the thermally grown oxide, with a thickness of 1 to 10 microns. It forms as Al2O3. 

The topmost layer, the ceramic topcoat, provides thermal insulation. It 

is typically of yttria-stabilized ZrO2. Its thickness is characteristically 

about 300 microns when deposited by air plasma spray and 125 microns 

when deposited by electron beam physical vapor deposition (EB-PVD).74

74. Nitin P. Padture, Maurice Gelland, and Eric H. Jordan, “Thermal Barrier Coatings for GasTurbine Engine Applications,” Science, vol. 296 (Apr. 12, 2002), pp. 280–284.
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Yttria-stabilized zirconia has become the preferred TBC layer material for use in jet engines because of its low thermal conductivity and its 

relatively high thermal expansion coefficient, compared with many other 

ceramics. This reduces the thermal expansion mismatch with the metals of high thermal expansion coefficient to which it is applied. It also 

has good erosion resistance, which is important because of the entrainment of particles having high velocity in the engine gases. Robert Miller, 

a leading specialist, notes that NASA and the NACA, its predecessor, have 

played a leading role in TBC development since 1942. Flame-sprayed 9

Rokide coatings, which extended the life of the X-15 main engine combustion chamber, represented an early success. Magnesia-stabilized zirconia later found use aboard the SR-71, allowing continuous use of the 

afterburner and sustained flight above Mach 3. By 1970, plasma-sprayed 

TBCs were in use in commercial combustors.75

These applications involved components that had no moving parts. 

For turbines, the mid-1970s brought the first “modern” thermal spray 

coating. It used yttria as a zirconia stabilizer and a bond coat that 

contained MCrAlY, and demonstrated that blade TBCs were feasible. 

C.W. Goward of Pratt & Whitney (P&W), writing of TBC experience 

with the firm’s J75 engine, noted: “Although the engine was run at relatively low pressures, the gas turbine engine community was sufficiently 

impressed to prompt an explosive increase in development funds and 

programs to attempt to achieve practical utilization of the coatings 

on turbine airfoils.”76

But tests in 1977 on the more advanced JT9D, also conducted at 

P&W, brought more mixed results. The early TBC remained intact on 

lower-temperature regions of the blade but spalled at high temperatures. This meant that further development was required. Stefan Stecura 

reported an optimum concentration of Y2O3 in ZrO2 of 6–8 percent. 

This is still the state of the art. H.G. Scott reported that the optimum 

phase of zirconia was t’-ZrO2. In 1987, Stecura showed that ytterbiastabilized zirconia on a ytterbium-containing bond coat doubled the 

blade life and took it from 300 1-hour cycles to 600 cycles. Also at that 

75. Robert A. Miller, “History of Thermal Barrier Coatings for Gas Turbine Engines,” NASA TM2009-215459 (2009).

76. G.W. Goward, “Seventeen Years of Thermal Barrier Coatings,” in Department of Energy, 

Proceedings of the Workshop on Coatings for Advanced Heat Engines, Castine, ME, July 27–30, 

1987 (Washington, DC: U.S. Department of Energy, 1987).

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time, P&W used a zirconia-yttria TBC to address a problem with endurance of vane platforms. A metallic platform, with no thermal barrier, 

showed burn-through and cracking from thermal-mechanical fatigue 

after 1,500 test cycles. Use of a TBC extended the service life to 18,000 

hours or 2,778 test cycles and left platforms that were clean, uncracked, 

and unburned. P&W shared these results with NASA, which led to the 

TBC task in the Hot Section Technology (HOST) program. NASA collaborated with P&W and four other firms as it set out to predict TBC 

lifetimes. A preliminary NASA model showed good agreement between 

experiment and calculation. P&W identified major degradation modes 

and gave data that also showed good correlation between measured 

and modeled lives. Other important contributions came from Garrett 

Turbine Co. and General Electric. The late 1980s brought Prescribed 

Velocity Distribution (PVD) blades that showed failure when they were 

nearly out of the manufacturer’s box. EV-PVD blades resolved this issue 

and first entered service in 1989 on South African Airways 747s. They 

flew from Johannesburg, a high-altitude airport with high mean temperatures where an airliner needed a heavy fuel load to reach London. 

EV-PVD TBCs remain the coating of choice for first-row blades, which 

see the hottest combustion gases. TBC research continues to this day, 

both at NASA and its contractors. Fundamental studies in aeronautics 

are important, with emphasis on erosion of turbine components. This 

work has been oriented toward rotorcraft and has brought the first 

EV-PVD coating for their blades. There also has been an emphasis on 

damping of vibration amplitudes. A new effort has dealt with environmental barrier coatings (EBCs), which Miller describes as “ceramic 

coatings, such as SiC, on top of ceramics.”77

Important collaborations have included work on coatings for diesels, where thick TBCs permit higher operating temperatures that yield 

increased fuel economy and cleaner exhaust. This work has proceeded 

with Caterpillar Tractor Co. and the Army Research Laboratory.78

Studies of supersonic engines have involved cooperation with P&W 

and GE, an industrial interaction that Miller described as “a useful 

77. Author interview with Robert A. Miller, June 10, 2009; see also Miller, NASA TM-2009-215459.

78. Dong-ming Zhu and Miller, “Investigation of Thermal High Cycle and Low Cycle Fatigue Mechanisms of Thick Thermal Barrier Coatings,” Materials Science and Engineering, vol. A245 (l998), pp. 

212–223.
Case 9 | High-Temperature Structures and Materials

613

reality check.”79 NASA has also pursued the Ultra Efficient Engine 

Technology program. Miller stated that it has not yet introduced engines 

for routine service but has led to experimental versions. This work 

has involved EBCs, as well as a search for low thermal conductivity. 

The latter can increase engine-operating temperatures and reduce 

cooling requirements, thereby achieving higher engine efficiency and 

lower emissions. At NASA Glenn, Miller and Dong-ming Zhu have 

built a test facility that uses a 3-kilowatt CO2 laser with wavelength 

of 10.6 microns. They also have complemented conventional ZrO2-

Y2O3 coatings with other rare-earth oxides, including Nd2O3-Yb2O3 

and Gd2O3-Yb2O3.80

Can this be reduced further? A promising approach involves development of new deposition techniques that give better control of TBC 

pore morphology. Air plasma spray deposition creates many intersplat 

pores between initially molten droplets, in what Miller described as “a 

messy stack of pancakes.” By contrast, TBC layers produced by EB-PVD 

have a columnar microstructure with elongated intercolumnar pores 

that align perpendicular to the plane of the coating. Alternate deposition methods include sputtering, chemical vapor deposition (CVD), 

and sol-gel approaches. But these approaches involve low deposition 

rates that are unsuitable for economic production of coated blades. 

CVD and sol-gel techniques also require the use of dangerous and costly 

precursor materials. In addition, none of these approaches permit the 

precise control and manipulation of pore morphology. Thus, improved 

deposition methods that control this morphology do not now exist.

Blade Fabrication

The fabrication of turbine blades represents a related topic. No blade 

has indefinite life, for blades are highly stressed and must resist 

creep while operating under continuous high temperatures. Table 

3 is taken from the journal Metallurgia and summarizes the stress 

to cause rupture in both wrought- and investment-case nickelbase superalloy.81

79. Miller interview.

80. Zhu and Miller, “Development of Advanced Low Conductivity Thermal Barrier Coatings,” International Journal of Applied Ceramic Technology, vol. l (2004), pp. 86–94.

81. R. McCallum, “Casting Critical Components,” Superalloys Source Book, pp. 286–291.

9
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9

TABLE 3:

STRESS TO CAUSE FAILURE OF VARIOUS ALLOYS

TYPE OF ALLOY STRESS TO CAUSE FAILURE AFTER:

Wrought Alloys: 100 hours at 1400 °F, MPa 50 hours at 1750 °F, MPa

Nimonic 80 340 48

Nimonic 105 494 127

Nimonic 115 571 201

Investment-Cast Alloys:

IN100 648 278

B1914 756 262

Mar-M246 756 309

An important development involved the introduction of directionally 

82. J.E. Northwood, “Improving Turbine Blade Performance by Solidification Control,” Superalloys 

Source Book, pp. 292–296.
Case 9 | High-Temperature Structures and Materials

615

A hypersonic scramjet configuration developed by Langley experts in the 1970s. The sharply 

swept double-delta layout set the stage for the National Aero-Space Plane program. NASA.

plate, it was separated from this plate by a helical single-crystal selector. 

A number of grains nucleated at the bottom of the selector, but most of 

them had their growth cut off by its walls, and only one grain emerged at 

the top. This grain was then allowed to grow and fill the entire mold cavity.

Creep-rupture tests showed that Alloy 454 had a temperature advantage of 75 to 125 °F over d.s. MAR-M200 + Hf, the strongest production-blade alloy. A 75 °F improvement in metal temperature capability 

corresponds to threefold improvement in life. Single-crystal Alloy 454 

thus was chosen as the material for the first-stage turbine blades of the 

JT9D-7R4 series of engines that were to power the Boeing 767 and the 

Airbus A310 aircraft, with engine certification and initial production 

shipments occurring in July 1980.83

Materials Research and Development: The NASP Legacy

The National Aero-Space Plane (NASP) program had much to contribute 

to metallurgy, with titanium being a particular point. It has lately come 

to the fore in aircraft construction because of its high strength-to-density ratio, high corrosion resistance, and ability to withstand moderately 

83. M. Gell, D.H. Duhl, and A.F. Giami, “The Development of Single Crystal Turbine Blades,” 

Superalloys Source Book, pp. 297–306.

9
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9

high temperatures without creeping. Careful redesign must be accomplished to include it, and it appears only in limited quantities in aircraft 

that are out of production. But newer aircraft have made increasing use 

of it, including the two largest manufacturers of medium- and longrange commercial jetliners, Boeing and Airbus, whose aircraft and their 

weight of titanium is shown in Table 4.

TABLE 4:

BOEING AND AIRBUS AIRCRAFT MAKING SIGNIFICANT USE OF TITANIUM

AIRCRAFT (INCLUDING THEIR ENGINES) WEIGHT OF TI, IN METRIC TONS

Boeing 787 134

Boeing 777 59

Boeing 747 45

Boeing 737 18

Airbus A380 145

Airbus A350 74

Airbus A340 32

Airbus A330 18

Airbus A320 12

These numbers offer ample evidence of the increasing prevalence of 

titanium as a mainstream (and hence no longer “exotic”) aviation material, mirroring its use in other aspects of the commercial sector. For 

example, in the 1970s, the Parker Pen Company used titanium in its T-1 

line of ball pens and rollerballs, which it introduced in 1971. Production 

stopped in 1972 because of the high cost of the metal. But hammerheads 

fabricated of titanium entered service in 1999. Their light weight allows 

a longer handle, which increases the speed of the head and delivers more 

energy to the nail while decreasing arm fatigue. Titanium also substantially diminishes the shock transferred to the user because it generates 

much less recoil than a steel hammerhead.

In advancing titanium’s use, techniques of powder metallurgy have 

been at the forefront. These methods give direct control of the microstructure of metals by forming them from powder, with the grains of 

powder sintering or welding together by being pressed in a mold at high 

temperature. A manufacturer can control the grain size independently of 

any heat-treating process. Powder metallurgy also overcomes restrictions 

on alloying by mixing in the desired additives as powdered ingredients.
Case 9 | High-Temperature Structures and Materials

617

Several techniques exist to produce the powders. Grinding a metal 

slab to sawdust is the simplest, though it yields relatively coarse grains. 

“Splat-cooling” gives better control. It extrudes molten metal onto the 

chilled rim of a rotating wheel that cools it instantly into a thin ribbon. 

This represents a quenching process that produces a fine-grained microstructure in the metal. The ribbon then is chemically treated with hydrogen, which makes it brittle so that it can be ground into a fine powder. 

Heating the powder then drives off the hydrogen.

The Plasma Rotating Electrode Process, developed by the firm of 

Nuclear Metals, has shown particular promise. The parent metal is 

shaped into a cylinder that rotates at up to 30,000 revolutions per minute (rpm) and serves as an electrode. An electric arc melts the spinning 

metal, which throws off droplets within an atmosphere of cool inert 

helium. The droplets plummet in temperature by thousands of degrees 

within milliseconds and their microstructures are so fine as to approach an 

amorphous state. Their molecules do not form crystals, even tiny ones, but 

arrange themselves in formless patterns. This process, called “rapid solidification,” has brought particular gains in high-temperature strength.

Standard titanium alloys lose strength at temperatures above 700 

to 900 °F. By using rapid solidification, McDonnell-Douglas raised this 

limit to 1,100 °F prior to 1986, when NASP got underway. Philip Parrish, 

the manager of powder metallurgy at the Defense Advanced Research 

Projects Agency (DARPA), notes that his agency spent some $30 million on rapid-solidification technology in the decade after 1975. In 1986, 

he described it as “an established technology. This technology now can 

stand along such traditional methods as ingot casting or drop forging.”84

Eleven-hundred degrees nevertheless was not enough. But after 1990, 

the advent of new baseline configurations for the X-30 led to an appreciation that the pertinent areas of the vehicle would face temperatures 

no higher than 1,500 °F. At that temperature, advanced titanium alloys 

could serve in metal matrix composites (MMCs), with thin-gauge metals being reinforced with fibers.

A particular composition came from the firm of Titanium Metals 

and was designated Beta-21S. That company developed it specifically for the X-30 and patented it in 1989. It consisted of Ti along with 

15Mo+2.8Cb+3Al+0.2Si. Resistance to oxidation proved to be its strong 

84. Heppenheimer, “Making Planes from Powder,” High Technology (Sept. 1986), pp. 54–55; 

author interview of Philip Parrish, Mar. 21, 1986.

9
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9

suit, with this alloy showing resistance that was two orders of magnitude 

greater than that of conventional aircraft titanium. Tests showed that it 

could also be exposed repeatedly to leaks of gaseous hydrogen without 

being subject to embrittlement. Moreover, it lent itself readily to being 

rolled to foil-gauge thicknesses of 4 to 5 mils in the fabrication of MMCs.85

There also was interest in using carbon-carbon for primary structure. Here the property that counted was not its heat resistance but its 

light weight. In an important experiment, the firm of LTV fabricated 

half of an entire wing box of this material. An airplane’s wing box is a 

major element of aircraft structure that joins the wings and provides a 

solid base for attachment of the fuselage fore and aft. Indeed, one could 

compare it with the keel of a ship. It extends to left and right of the aircraft centerline, with LTV’s box constituting the portion to the left of this 

line. Built at full scale, it represented a hot-structure wing proposed by 

General Dynamics. It measured 5 by 8 feet, with a maximum thickness 

of 16 inches. Three spars ran along its length, five ribs were mounted 

transversely, and the complete assembly weighed 802 pounds.

The test plan called for it to be pulled upward at the tip to reproduce the bending loads of a wing in flight. Torsion or twisting was to 

be applied by pulling more strongly on the front or rear spar. The maximum load corresponded to having the X-30 execute a pullup maneuver 

at Mach 2.2 with the wing box at room temperature. With the ascent 

continuing and the vehicle undergoing aerodynamic heating, the next 

key event brought the maximum difference in the temperatures of the 

top and bottom of the wing box, with the former being at 994 °F and the 

latter being at 1,671 °F. At that moment, the load on the wing box corresponded to 34 percent of the Mach 2.2 maximum. Farther in the flight the 

wing box was to reach peak temperature, 1,925 °F, on the lower surface. 

These three points were to be reproduced through mechanical forces 

applied at the ends of the spars and through the use of graphite heaters.

But several important parts delaminated during their fabrication, 

which seriously compromised the ability of the wing box to bear its 

specified loads. Plans to impose the peak or Mach 2.2 load were abandoned, with the maximum planned load being reduced to the 34 percent 

85. J. Sorensen, “Titanium Matrix Composites—NASP Materials and Structures Augmentation Program,” AIAA Paper 90-5207 (1990); Stanley W. Kandebo, “Boeing 777 to Incorporate New Alloy 

Developed for NASP,” Aviation Week (May 3, 1993), p. 36; “NASP Materials and Structures 

Program: Titanium Matrix Composites,” McDonnell-Douglas, DTIC ADB-192559, (Dec. 31, 1991).
Case 9 | High-Temperature Structures and Materials

619

associated with the maximum temperature difference. For the same reason, the application of torsion was deleted from the test program. Amid 

these reductions in the scope of the structural tests, two exercises went 

forward during December 1991. The first took place at room temperature and successfully reached the mark of 34 percent without causing 

further damage to the wing box.

The second test, a week later, reproduced the condition of peak temperature difference while briefly applying the calculated load of 34 percent. The plan then called for further heating to the peak temperature 

of 1,925 °F. As the wing box approached this value, a difficulty arose 

because of the use of metal fasteners in its assembly. Some were made 

from coated columbium and were rated for 2,300 °F, but most were a 

nickel alloy that had a permissible temperature of 2,000 °F. However, an 

instrumented nickel-alloy fastener overheated and reached 2,147 °F. The 

wing box showed a maximum temperature of 1,917 °F at that moment, 

and the test was terminated because the strength of the fasteners now 

was in question. This test nevertheless counted as a success because it 

had come within 8 degrees of the specified temperature.86

Both tests thus were marked as having achieved their goals, but 

their merits were largely in the mind of the beholder. The entire project 

would have been far more impressive if it had avoided delamination, 

had successfully achieved the Mach 2.2 peak load, and had subjected 

the wing box to repeated cycles of bending, torsion, and heating. This 

effort stood as a bold leap toward a future in which carbon-carbon might 

take its place as a mainstream material, but it was clear that this future 

would not arrive during the NASP program. However, the all-carboncomposite airplane, as distinct from one of carbon-carbon, has now 

become a reality. Carbon alone has high temperature resistance, whereas 

carbon composite burns or melts readily. The airplane that showcases 

carbon composites is the White Knight 2, built by Burt Rutan’s Scaled 

Composites firm as part of the Virgin Galactic venture that is to achieve 

commercial space flight. As of this writing, White Knight 2 is the world’s 

largest all-carbon-composite aircraft in service; even its control wires 

are carbon composite. Its 140-foot-span wing is the longest single carbon composite aviation component ever fabricated.87

86. John Bradley, “Test and Evaluation Report for Carbon/Carbon Wing Box Component.” General Dynamics, DTIC ADB-191627, (Feb. 21, 1992).

87. Leonard David, “Lift for Public Space Travel,” Aerospace America (Feb. 2009), pp. 24–29.

9
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9

Far below this rarefied world of transatmospheric tourism, carbon 

composites are becoming the standard material for commercial aviation. Aluminum has held this role up to Mach 2 since 1930, but after 

80 years, Boeing is challenging this practice with its 787 airliner. By 

weight, the 787 is 50 percent composite, 20 percent aluminum, 15 percent titanium, 10 percent steel, and 5 percent other. The 787 is 80 percent composite by volume. Each of them contains 35 tons of composite 

reinforced with 23 tons of carbon fiber. Composites are used on fuselage, wings, tail, doors, and interior. Aluminum appears at wing and 

tail leading edges, with titanium used mainly on engines. The extensive application of composites promotes light weight and long range. 

The 787 can fly nonstop from New York City to Beijing. The makeup 

of the 787 contrasts notably with that of the Boeing 777. Itself considered revolutionary when it entered service in 1995, it nevertheless had 

a structure that was 50 percent aluminum and 12 percent composite. 

The course to all-composite construction is clear, and if the path is not 

yet trodden, nevertheless, the goal is clearly in sight. As in 1930, when 

all-metal structures first predominated in American commercial aviation, the year 2010 marks the same point for the evolution of commercial composite aircraft.

The NASP program also dealt with beryllium. This metal had only 

two-thirds the density of aluminum and possessed good strength, but 

its temperature range was restricted. The conventional metal had a limit 

of some 850 °F, while an alloy from Lockheed called Lockalloy, which 

contained 38 percent aluminum, was rated only for 600 °F. It had never 

become a mainstream material like titanium, but, for the X-30, it offered 

the advantage of high thermal conductivity. Work with titanium had 

greatly increased its temperatures of use, and there was hope of achieving similar results with beryllium.

Initial efforts used rapid-solidification techniques and sought 

temperature limits as high as 1,500 °F. These attempts bore no fruit, 

and from 1988 onward the temperature goal fell lower and lower. In 

May 1990, a program review shifted the emphasis away from hightemperature formulations toward the development of beryllium as a 

metal suitable for use at cryogenic temperatures. Standard forms of this 

metal became unacceptably brittle when only slightly colder than –100 

°F, but cryoberyllium proved to be out of reach as well. By 1992, investigators were working with ductile alloys of beryllium and were sacrificing all prospect of use at temperatures beyond a few hundred degrees 
Case 9 | High-Temperature Structures and Materials

621

but were winning only modest improvements in low-temperature capability. Terence Ronald, the NASP materials director, wrote in 1995 of 

rapid-solidification versions with temperature limits as low as 500 °F, 

which was not what the X-30 needed to reach orbit.88

In sum, the NASP materials effort scored a major advance with Beta21S, but the genuinely radical possibilities failed to emerge. These included 

carbon-carbon as primary structure along with alloys of beryllium that 

were rated for temperatures well above 1,000 °F. The latter, if available, 

might have led to a primary structure with the strength and temperature 

resistance of Beta-21S but with less than half the weight. Indeed, such 

weight savings would have ramified throughout the entire design, leading to a configuration that would have been smaller and lighter overall.

Generally, work with materials fell well short of its goals. In dealing 

with structures and materials, the contractors and the National Program 

Office established 19 program milestones that were to be accomplished 

by September 1993. A General Accounting Office program review, issued 

in December 1992, noted that only six of them would indeed be completed.89 This slow progress encouraged conservatism in drawing up the 

bill of materials, but this conservatism carried a penalty.

When the scramjets faltered in their calculated performance and 

the X-30 gained weight while falling short of orbit, designers lacked 

recourse to new and very light materials, such as beryllium and 

carbon-carbon, that might have saved the situation. With this, NASP spiraled to its end. The future belonged to other less ambitious but more 

attainable programs, such as the X-43 and X-51. They, too, would press 

the frontier of aerothermodynamic structural design, as they pioneered 

the hypersonic frontier.

88. Rockwell International, “High Conductivity Composites: Executive Summary, Copper 

Materials, Beryllium Materials, Coatings, Ceramic Materials and Joining,” DTIC ADB-191898 

(Mar. 1993); Terence Ronald, “Status and Applications of Materials Developed for NASP,” AIAA 

Paper 95-6131 (1995).

89. Frank C. Conahan, “National Aero-Space Plane: Restructuring Further Research and Development Efforts,” General Accounting Office Report NSIAD-93-71 (Dec. 3, 1992), table, p. 20.

9
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9

Recommended Additional Readings

Reports, Papers, Articles, and Presentations:

H. Julian Allen and A.J. Eggers, Jr., “A Study of the Motion and 

Aerodynamic Heating of Ballistic Missiles Entering the Earth’s 

Atmosphere at High Supersonic Speeds,” NACA TR-1381 (1953).

Anon., “How Mercury Capsule Design Evolved,” Aviation Week (Sept. 21, 

1959).

Anon., “Fourth ASSET Glider Gathers Data,” Aviation Week (Nov. 2, 

1964).

Anon., “Sortie Module May Cut Experiment Cost,” Aviation Week (Jan. 

17, 1972).

F.M. Anthony, R.R. Fisher, and R.G. Helenbrook, “Selection of Space 

Shuttle Thermal Protection Systems,” AIAA Paper 71-443 (1971).

Paul R. Becker, “Leading-Edge Structural Material System of the Space 

Shuttle,” American Ceramic Society Bulletin, vol. 60, No. 11 (1981), 

pp. 1210–1214.

C.E. Brown, W.J. O’Sullivan, and C.H. Zimmerman, “A Study of the 

Problems Related to High Speed, High Altitude Flight,” NACA 

Langley Research Center, 1953.

Frank C. Conahan, “National Aero-Space Plane: Restructuring Further 

Research and Development Efforts,” General Accounting Office 

Report NSIAD-93-71 (Dec. 3, 1992).

William Cowie, “Utilization of Refractory Metals on the X-20A (DynaSoar),” Air Force Systems Command, June 1964, DTIC AD-609169.

Leonard David, “Lift for Public Space Travel,” Aerospace America (Feb. 

2009), pp. 24–29.

Richard A. DeMeis, “The Trisonic Titanium Republic,” Air Enthusiast, 

No. 7 (July–Sept. 1978), pp. 198–213.
Case 9 | High-Temperature Structures and Materials

623

Grumman Corporation, “Space Shuttle Program Definition: Phase B 

Extension Final Report,” B35-43 RP-33 (Mar. 15, 1972).

T.A. Heppenheimer, “Making Planes from Powder,” High Technology

(Sept. 1986), pp. 54–55.

William S. Hieronymus, “Two Reusable Materials Studied for Orbiter 

Thermal Protection,” Aviation Week (Mar. 27, 1972), p. 48.

Stanley W. Kandebo, “Boeing 777 to Incorporate New Alloy Developed 

for NASP,” Aviation Week (May 3, 1993), p. 36.

L.J. Korb, C.A. Morant, R.M. Calland, and C.S. Thatcher, “The Shuttle 

Orbiter Thermal Protection System,” American Ceramic Society 

Bulletin, vol. 60 (1981), pp. 1188–1193.

Martin Marietta, “SV-5D PRIME Final Flight Test Summary,” Report 

ER-14465 (Sept. 1967).

McDonnell Aircraft Corp., “ASSET ASV-3 Flight Test Report,” Jan. 4, 

1965, DTIC AD-357523.

McDonnell Aircraft Corp., “ASSET ASV-4 Flight Test Report,” June 25, 

1965, DTIC AD-366546.

McDonnell-Douglas, “External LH2 Tank Study Final Report,” MDC 

E076-1, June 30, 1971.

McDonnell-Douglas, “Interim Report to OMSF: Phase B System Study 

Extension,” Sept. 1, 1971.

McDonnell-Douglas, “NASP Materials and Structures Program: Titanium 

Matrix Composites,” Dec. 31, 1991.

Howard J. Middendorf, “Materials and Processes for X-20A (Dyna-Soar),” 

Air Force Systems Command, June 1964, DTIC AD-449685.

Robert A. Miller, “History of Thermal Barrier Coatings for Gas Turbine 

Engines,” NASA TM-2009-215459 (2009).

9
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624

9

Dale D. Myers, “The Navaho Cruise Missile: A Burst of Technology,” Acta 

Astronautica, vol. 26 (Nov. 8–10, 1992), pp. 741–748.

William J. Normyle, “Manned Flight Tests to Seek Lifting-Body 

Technology,” Aviation Week (May 16, 1966), pp. 64–75.

Scott Pace, “Engineering Design and Political Choice: The Space Shuttle 

1969–72,” master’s thesis (Cambridge: Massachusetts Institute of 

Technology, May 1982).

Nitin P. Padture, Maurice Gelland, and Eric H. Jordan, “Thermal Barrier 

Coatings for Gas-Turbine Engine Applications,” Science, vol. 296 

(Apr. 12, 2002), pp. 280–284.

Lee L. Peterson, “Evaluation Report on X-7A,” Report AFMDC ADJ 

57-8184, Oct. 3, 1957.

Joel W. Powell, “Thor-Able and Atlas-Able,” Journal of the British 

Interplanetary Society (May 1984), pp. 219–225.

William A. Ritchie, “Evaluation Report on X-7A (System 601B),” Report 

AFMDC DAS 58-8129, Jan. 1959.

Rockwell International, “High Conductivity Composites: Executive 

Summary, Copper Materials, Beryllium Materials, Coatings, Ceramic 

Materials and Joining” (Mar. 1993), DTIC ADB-191898.

Terence Ronald, “Status and Applications of Materials Developed for 

NASP,” AIAA Paper 95-6131 (1995).

P.H. Rose, “Physical Gas Dynamics Research at the Avco Research Lab,” 

NATO Advisory Group for Aeronautical Research and Development 

Report 145, July 1957.

P.H. Rose and W.I. Stark, “Stagnation Point Heat-Transfer Measurements 

in Dissociated Air,” Journal of the Aeronautical Sciences (Feb. 1958), 

pp. 86–97.
Case 9 | High-Temperature Structures and Materials

625

R.L. Schleicher, “Structural Design of the X-15,” Journal of the Royal 

Aeronautical Society (Oct. 1963), pp. 618–636.

Wilson B. Schramm, Ronald P. Banas, and Y. Douglas Izu, “Space Shuttle 

Tile—The Early Lockheed Years,” Lockheed Horizons, No. 13 (1983), 

pp. 2–15.

Wilson Schramm, “HRSI and LRSI—The Early Years,” American Ceramic 

Society Bulletin, vol. 60 (1981), pp. 1194–1195.

Alvin Seiff and H. Julian Allen, “Some Aspects of the Design of Hypersonic 

Boost-Glide Aircraft,” NACA RM-A55E26 (1955).

J. Sorensen, “Titanium Matrix Composites—NASP Materials and 

Structures Augmentation Program,” AIAA Paper 90-5207 (1990).

Leo Steg, “Materials for Re-Entry Heat Protection of Satellites,” American 

Rocket Society Journal (Sept. 1960), pp. 815–822.

George W. Sutton, “The Initial Development of Ablation Heat Protection, 

An Historical Survey,” Journal of Spacecraft and Rockets (Jan.–Feb. 

1982), pp. 3–11.

B.K. Thomas, “USAF Nears Lifting Body Tests,” Aviation Week (July 10, 

1967), pp. 99–101.

Richard C. Thuss, Harry G. Thibault, and Arnold Hiltz, “The Utilization 

of Silica Based Thermal Insulation for the Space Shuttle Thermal 

Protection System,” CASI 72A-10764, Oct. 1971.

Vought Missiles and Space Company, “Technical Overview: Oxidation 

Resistant Carbon-Carbon for the Space Shuttle,” n.d. (c. 1970).

Michael L. Yaffee, “Program Changes Boost Grumman Shuttle,” Aviation 

Week (July 12, 1963), pp. 36–39.

9
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9

Richard E. Young, Martha A. Smith, and Charles K. Sobeck, “Galileo 

Probe: In Situ Observations of Jupiter’s Atmosphere,” Science (May 

10, 1996), pp. 837–838.

U.S. Air Force Aeronautical Systems Division staff, Proceedings of 1962 

X-20A (Dyna-Soar) Symposium, vol. 3, Structures and Materials

(Wright Patterson AFB: HQ ASD, 1963).

U.S. Air Force Systems Command staff, “Advanced Technology Program: 

Technical Development Plan for Aerothermodynamic/Elastic Structural 

System Environmental Tests (ASSET),” HQ AFSC, Sept. 9, 1963.

U.S. Army Ballistic Missile Agency staff, Re-Entry Studies, vol. 1 

(Redstone Arsenal, AL: HQ ABMA, Nov. 25, 1958).

U.S. Department of Energy, Proceedings of the Workshop on Coatings for 

Advanced Heat Engines, Castine, ME, July 27–30, 1987 (Washington, 

DC: HQ DoE, 1987).

U.S. National Aeronautics and Space Administration staff, NASA Space 

Shuttle Technology Conference, vol. 2, Structures and Materials, NASA 

TM-X-2273 (1971).

Dong-ming Zhu and Robert A. Miller, “Investigation of Thermal High 

Cycle and Low Cycle Fatigue Mechanisms of Thick Thermal Barrier 

Coatings,” Materials Science and Engineering, vol. A245 (1998), pp. 

212–223.

Dong-ming Zhu and Robert A. Miller, “Development of Advanced Low 

Conductivity Thermal Barrier Coatings,” International Journal of 

Applied Ceramic Technology, vol. l (2004), pp. 86–94.

Books and Monographs:

Morton Alperin and Hollingsworth F. Gregory, eds., Vistas in Astronautics, 

vol. 2 (New York: Pergamon Press, 1959).

John Becklake, ed., History of Rocketry and Astronautics, vol. 17 of the 

AAS History Series (San Diego: Univelt, 1995).
Case 9 | High-Temperature Structures and Materials

627

Deborah D.L. Chung, Carbon Fiber Composites (Boston: ButterworthHeinemann, 1994).

Erik M. Conway, High-Speed Dreams (Baltimore: Johns Hopkins 

University Press, 2005).

Paul F. Crickmore, Lockheed SR-71 Blackbird (London: Osprey Publishing 

Ltd., 1986).

Matthew J. Donachie, Jr., Superalloys Source Book (Metals Park, OH: 

American Society for Metals, 1984).

Richard P. Hallion, ed., The Hypersonic Revolution: Case Studies in the 

History of Hypersonic Technology, vols. 1 and 2 (Bolling AFB: USAF, 

1998).

James R. Hansen, Engineer in Charge: A History of the Langley Aeronautical 

Laboratory, 1917–1958, NASA SP-4305 (Washington: GPO, 1987).

T.A. Heppenheimer, Turbulent Skies: The History of Commercial Aviation

(New York: John Wiley, 1995).

T.A. Heppenheimer, The Space Shuttle Decision, NASA SP-4221 

(Washington, DC: NASA, 1999).

T.A. Heppenheimer, Development of the Space Shuttle, 1972–1981

(Washington: Smithsonian Institution Press, 2002).

Dennis Jenkins, Space Shuttle (Stillwater, MN: Voyageur Press, 2001).

Willy Ley, Rockets, Missiles, and Space Travel (New York: Viking, 1957).

Jay Miller, The X-Planes, X-1 to X-45 (North Branch, MN: Specialty Press, 

2001).

NASA, Columbia Accident Investigation Report (Washington: GPO, 2003).

Jacob Neufeld, The Development of Ballistic Missiles in the United States 

Air Force, 1945–1960 (Washington: Office of Air Force History, 1990).

9
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9

David Robarge, Archangel: CIA’s Supersonic A-12 Reconnaissance Aircraft

(McLean, VA: CIA Center for the Study of Intelligence, 2007).

Kevin C. Ruffner, Corona: America’s First Satellite Program (McLean, VA: 

Central Intelligence Agency, 1995).
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The AFTI/F-16, shown here on a 1991 test flight, flew 15 years and over 700 research flights 

at NASA Dryden. NASA.
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Fly-By-Wire: 

Making the Electric Jet

Albert C. Piccirillo

T

HE EVOLUTION OF ADVANCED AIRCRAFT equipped with computerized flight and propulsion control systems goes back a long way 

and involved many players, both in the United States and internationally. During the Second World War, use of electronic sensors and 

subsystems began to become pervasive, adding new mission capabilities as well as increasing complexity. Autopilots were coupled to flight 

control systems, and electric trim was introduced. Hydraulically or 

electrically boosted flight control surfaces, along with artificial feel and 

stability augmentation systems, soon followed as did electronic engine 

control systems. Most significantly, the first uses of airborne computers 

emerged, a trend that soon resulted in their use in aircraft and missiles 

for their flight and mission control systems. Very early on, there was a 

realization that traditional mechanical linkages between the cockpit flight 

controls and the flight control surfaces could be replaced by a computercontrolled fly-by-wire (FBW) approach in which electric signals were transmitted from the pilot’s controls to the control surface actuators by wire. 

This approach was understood to have the potential to reduce mechanical complexity, lower weight, and increase safety and reliability. In addition, the processing power of the computer could be harnessed to enable 

unstable aircraft designs to be controlled. Properly tailored, these unstable designs could enable new aircraft concepts to be implemented that 

could fully exploit the advantages of active flight control. Such aircraft 

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Fly-by-wire (FBW) technology pioneered by NASA has enabled the 

design of highly unconventional airframe configurations. Continuing 

NASA FBW research has validated integrated digital propulsion 

and flight control systems. FBW has been applied to civil aircraft, 

improving their safety and efficiency. Lessons learned from NASA 

FBW research have been transferred to maritime design, and 

Agency experts have supported the U.S. Navy in developing digital 

electronic ship control.
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NASA’s Contributions to Aeronautics

would be more maneuverable and lighter, have better range, and also 

allow for the fully integrated control of aircraft, propulsion, navigation, 

and mission systems to optimize overall capability.

Two significant events unfolded during the 1960s that fostered the 

move to electronic flight control systems. The space race was a major 

influence, with most space systems relying on computerized fly-by-wire 

control systems for safe and effective operation. The Vietnam war provided a strong impetus for the development of more survivable aircraft 

systems as well for new aircraft with advanced performance features. 

The National Aeronautics and Space Administration (NASA) and the 

Air Force aggressively responded to these challenges and opportunities, 

resulting in the rapid transition of digital computer technology from the 

space program into aircraft fly-by-wire applications as exemplified by 

the Digital Fly-By-Wire F-8 and the AFTI/F-16 research programs. Very 

quickly after, a variety of flight research programs were implemented. 

These programs provided the basis for development and fielding of 

numerous military and civil aircraft equipped with advanced digital flyby-wire flight control systems. On the civil aviation side, safety has been 

improved by preventing aircraft flight envelope limitations from being 

exceeded. Operating efficiency has been greatly enhanced and major 

weight savings achieved from fly-by-wire and related electronic flight 

control system components. Integrated flight and propulsion control 

systems precisely adjust throttles and fuel tank selections. Rudder trim 

drag because of unbalanced engine thrust is reduced. Fuel is automatically transferred between tanks throughout the aircraft to optimize center of gravity during cruise flight, minimizing elevator trim drag. In the 

case of advanced military aircraft, electronically controlled active flight, 

propulsion, and mission systems have been fully integrated, providing 

revolutionary improvements in capabilities. Significantly, new highly 

unstable aircraft configurations are providing unprecedented levels of 

mission performance along with very low radar signatures, capabilities 

that have been enabled by exploiting digital fly-by-wire flight and propulsion control systems pioneered in NASA.

Aircraft Flight Control: Beginnings to the 1950s

Early aviation pioneers gradually came to realize that an effective flight 

control system was necessary to control the forces and moments acting on an aircraft. The creation of such a system of flight control was 

one of the great accomplishments of the Wright brothers, who used a 
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combination of elevator, rudder, and wing warping to achieve effective 

three-axis flight control in their 1903 Flyer. As aircraft became larger 

and faster, wing warping was replaced by movable ailerons to control 

motion around the roll axis. This basic three-axis flight control system 

is still used today. It enables the pilot to maneuver the aircraft about its 

pitch, yaw, and roll axes and, in conjunction with engine power adjustments, to control velocity and acceleration as well. For many generations after the dawn of manned, controllable powered flight, a system 

of direct mechanical linkages between the pilot’s cockpit controls and 

the aircraft’s control surfaces was used to both assist in stabilizing the 

aircraft as well as to change its flight path or maneuver. The pilot’s ability to “feel” the forces being transmitted to his flight controls, especially 

during rapid maneuvering, was critically important, because many early 

aircraft were statically unstable in pitch with the pilot having to exert 

a constant stabilizing influence with his elevator control. Well into the 

First World War, many aircraft on both sides of the conflict had poor 

stability and control characteristics, issues that would continue to challenge aircraft designers well into the jet age. Wartime experience showed 

that adequate stability and positive aircraft handling qualities, coupled 

with high performance (as exemplified by parameters such as low wing 

loading, high power-to-weight ratio, and good speed, turn, and climb 

rates) played a major role in success in combat between fighters.

By the Second World War, electrically operated trim tabs located on 

aerodynamic control surfaces and other applications of electrical control and actuation were emerging.1

 However, as aircraft performance 

increased and airframes grew larger and heavier, it became increasingly harder for pilots to maneuver their aircraft because of high aero1. An early example was the innovative Focke-Wulf FW 190 fighter of World War II fame. Its landing gear and flaps were electrically controlled and actuated by small reliable high-torque electric 

motors. A variable incidence electrically controlled and actuated movable horizontal stabilizer 

maintained trim about the longitudinal (pitch) axis and was regularly used to relieve the heavy 

control forces encountered during steep dives (the FW 190 exceeded Mach 0.8 during testing). 

Another highly advanced FW 190 feature was the BMW 801 Kommandogerät (command equipment). This consisted of a mechanical-hydraulic analog computer that automatically adjusted engine 

fuel flow, variable pitch propeller setting, supercharger setting, fuel mixture, and ignition timing in 

response to pilot commands via the single throttle lever. This pioneering step in computerized integrated propulsion control systems greatly simplified engine control. See Albert C. Piccirillo, “Electric 

Aircraft Pioneer—The Focke-Wulf Fw 190,” Society of Automotive Engineers (SAE) Paper 965631, 

Oct. 1996.
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dynamic forces on the control surfaces. World War II piston engine 

fighters were extremely difficult to maneuver in pitch and roll and often 

became uncontrollable as compressibility effects were encountered as 

speeds approached about Mach 0.8. The introduction of jet propulsion 

toward the later stages of the Second World War further exacerbated 

this controllability problem. Hydraulically actuated fight control surfaces were introduced to assist the pilot in moving the control surfaces 

at higher speeds. These “boosted” control surface actuators were connected to the pilot’s flight controls through a system consisting of cables, 

pulleys, and cranks, with hydraulic lines now also being routed through 

the airframe to power the control surface actuators.2

With a boosted flight control system, the pilot’s movements of the 

cockpit flight controls opens or closes servo valves in the hydraulic system, increasing or decreasing the hydraulic pressure powering the actuators that move the aircraft control surfaces. Initially, hydraulic boost 

augmented the force transmitted to the control surfaces by the pilot; such 

an approach is referred to as partial boost. However, fully boosted flight 

controls quickly became standard on larger aircraft as well as on those 

aircraft requiring high maneuverability at high indicated airspeeds and 

Mach numbers. The first operational U.S. jet fighter, the Lockheed P-80 

Shooting Star, used electric pitch and roll trim and had hydraulically 

boosted ailerons to provide roll effectiveness at higher airspeeds. The 

first jet-powered U.S. bomber to enter production, the four-engine North 

American B-45, flew for the first time in March 1947. It had hydraulically 

boosted flight control surfaces and an electrically actuated trim tab on 

the elevator that was used to maintain longitudinal trim. Despite their 

undeniable benefits, boosted flight control systems could also produce 

unanticipated hazards. The chief test pilot for the Langley Aeronautical 

Laboratory of what was then the National Advisory Committee for 

Aeronautics (NACA), Herbert “Herb” Hoover, was killed in the crash of 

a B-45 on August 14, 1952, when the aircraft disintegrated during a test 

mission near Barrowsville, VA.3

2. Control cables were replaced with rigid pushrods in German World War II flight control systems, 

for example in the FW 190 and the Fiesler Fi-103 pulsejet-powered cruise missile (the V-1). Pushrods minimized effects produced by control cables stretching under load.

3. “NASA Dryden Flight Research Center Pilot Biographies.” http://www.nasa.gov/centers/dryden/

news/Biographies/Pilots/index.html, accessed July 21, 2009. Herb Hoover was the first civilian pilot 

to fly faster than the speed of sound. He exceeded Mach 1 in the Bell XS-1 on Mar. 10, 1948.
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As a NASA report noted: “The aerodynamic power of the trim-tabelevator combination [on the B-45] was so great that, in the event of 

an inadvertent maximum tab deflection, the pilot’s strength was insufficient to overcome the resulting large elevator hinge moments if the 

hydraulic boost system failed or was turned off. Total in-flight destruction of at least one B-45, the aircraft operated by NACA, was probably 

caused by this combination of circumstances that resulted in a normal 

load factor far greater than the design value.”4

The Air Force/NACA Bell XS-1 rocket-powered research aircraft was 

equipped with an electrically trimmed adjustable horizontal stabilizer. 

It enabled the pilot to maintain pitch control as the conventional elevator lost effectiveness as the speed of sound was approached.5

 Using this 

capability, U.S. Air Force (USAF) Capt. Chuck Yeager exceeded Mach 

1 in level flight in the XS-1 in October 1947, followed soon after by a 

North American XP-86 Sabre (although in a dive). Hydraulically boosted 

controls and fully movable horizontal tails were rapidly implemented 

on operational high-performance jet aircraft, an early example being 

the North American F-100.6 To compensate for the loss of natural feedback to the pilot with fully boosted flight controls, various devices such 

as springs and bob weights were integrated into the flight control system. These “artificial feel” devices provided force feedback to the pilot’s 

controls that was proportional to changes in airspeed and acceleration. 

Industry efforts to develop boosted fight control surfaces directly benefited from NACA flight-test efforts of the immediate postwar period.

Fly-By-Wire: The Beginnings

The Second World War witnessed the first applications of computercontrolled fly-by-wire flight control systems. With fly-by-wire, primary 

control surface movements were directed via electrical signals transmitted by wires rather than by the use of mechanical linkages. The 

German Army’s A-4 rocket (the famous V-2 that postwar was the basis 

for both U.S. and Soviet efforts to move into space) used an electronic 

4. Lawrence K. Loftin, “Quest for Performance: The Evolution of Modern Aircraft,” Part II, ch. 12, 

Science and Technology Branch, NASA SP-468 (1985).

5. Richard P. Hallion, “The Air Force and the Supersonic Breakthrough,” published in Technology 

and the Air Force: A Retrospective Assessment, Air Force History and Museums Program, Washington, DC, 1997.

6. Hallion, “On the Frontier: Flight Research at Dryden, 1946–1981,” NASA SP-4303 (1984).
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analog computer that modeled the differential equations governing the 

missile’s flight control laws. The computer-generated electronic signals 

were transmitted by wire to direct movement of the actuators that drove 

graphite vanes located in the rocket motor exhaust. The thrust of the 

rocket engine was thus vectored as required to stabilize the V-2 missile 

at lower airspeeds until the aerodynamic control surfaces on the fins 

became effective.7 Postwar, a similar analog computer-controlled flyby-wire thrust vectoring approach was used in the U.S. Army Redstone 

missile, perhaps not surprisingly, because Redstone was predominantly 

designed by a team of German engineers headed by Wernher von Braun 

of V-2 fame. The Redstone would be used to launch the Mercury space 

capsule that carried Alan Shepard (the first American into space) in 1961.

The German Mistle (Mistletoe) composite aircraft of late World War 

II was probably the first example of the use of fly-by-wire for flight control in a manned aircraft application. Mistle consisted of a fighter (usually a Focke-Wulf FW 190) mounted on a support structure on a Junkers 

Ju 88 bomber.8 The Ju 88 was equipped with a 3,500-pound warhead and 

was intended to be flown to the vicinity of its target by the FW 190 pilot, 

at which time he would separate from the bomber and evade enemy 

defenses while the Ju 88 flew into its target. Potentiometers at the base 

of the FW 190 pilot’s control stick generated electrical commands that 

were transmitted via wire through the support structure to the bomber. 

These electrical commands activated electric motors that moved the system of pushrods leading to the Ju 88 control surfaces.9

Another electronic flight control system innovation related to the 

fly-by-wire concept had its origins in electronic feedback flight control 

research that began in Germany in the late 1930s and was published 

by Ernst Heinkel and Eduard Fischel in 1940. Their research was used 

in the 1944 development of a directional stability augmentation system 

for the Luftwaffe’s heavily armed and armored Henschel Hs 129 ground 

7. Michael J. Neufeld, “The Rocket and the Reich; Peenemünde and the Coming of the Ballistic Missile Era,” The Smithsonian Institution, published by The Free Press, a division of Simon and Schuster, 

New York, 1995. James E. Tomayko, “Computers Take Flight: A History of NASA’s Pioneering 

Digital Fly-by Wire Project,” NASA SP-2000-4224 (2000).

8. Much like the Space Shuttle is carried on top of the NASA Boeing 747s when it is ferried from 

Edwards AFB to Cape Kennedy, FL.

9. Tomayko, “Blind Faith: The United States Air Force and the Development of Fly-By-Wire Technology,” Technology and the Air Force: A Retrospective Assessment, Air Force History and Museums 

Program, Washington, DC, 1997.
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attack aircraft to compensate for an inherent Dutch roll10 instability that 

affected strafing accuracy with its large-caliber, low-rate-of-fire antitank 

cannon.11 This consisted of modifying the rudder portion of the flight 

control system for dual mode operation. The rudder was split into two 

sections, with the lower portion directly linked to the pilot’s flight controls. The upper section was electromechanically linked to a gyroscopic 

yaw rate sensor that automatically provided rudder corrections as yawing 

motions were detected.12 This was the first practical aircraft yaw damper. 

Northrop incorporated electronic stability augmentation devices into its 

YB-49 flying wing bomber that first flew in late 1947 in an attempt to 

compensate for serious directional stability problems. After the war, the 

NACA Ames Aeronautical Laboratory conducted extensive flight research 

into artificial stability. An NACA-operated Grumman F6F-3 Hellcat was 

modified to incorporate roll and yaw rate servos that provided stability 

augmentation, with flight tests beginning in 1948. In the following years, 

a number of other aircraft were modified by the NACA at Ames for variable stability research, including several variants of the North American 

F-86.13 By the 1950s, most high-performance swept wing jet-powered 

aircraft were designed with electronic stability augmentation devices.

Early Aircraft Fly-By-Wire Applications

By the 1950s, fully boosted flight controls were common, and the 

potential benefits of fly-by-wire were becoming increasingly apparent. 

Beginning during the Second World War and continuing postwar, fly-bywire and power-by-wire flight control systems had been fielded in various target drones and early guided missiles.14 However, most aircraft 

designers were reluctant to completely abandon mechanical linkages to 

10. Dutch roll is a term commonly used to describe an out of phase combination of yawing and 

rolling. Dutch roll instability can be improved by the use of a yaw damper.

11. The Hs 129 was armed with a variety of cannon that could include a 75-millimeter antitank gun.

12. Norman C. Weingarten, “History of In-Flight Simulation & Flying Qualities Research at CALSPAN,” AIAA Journal of Aircraft, vol. 42, no. 2, March/April 2005.

13. Paul F. Borchers, James A. Franklin, and Jay W. Fletcher, “Flight Research at Ames, 1940–

1997: Fifty-Seven Years of Development and Validation of Aeronautical Technology,” NASA 

SP-3300 (1998).

14. Rowland F. Pocock, “German Guided Missiles of the Second World War,” Arco Publishing 

Company, Inc., New York, 1967. These included Matador, Snark, BOMARC, Rascal, plus many others developed in a number of countries. See Gavin D. Jenny, James W. Morris, and Vernon R. Schmitt, 

“Fly-by-Wire, A Historical and Design Perspective,” The Society of Automotive Engineers, 1998.
NASA’s Contributions to Aeronautics

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flight control surfaces in piloted aircraft, an attitude that would undergo 

an evolutionary change over the next two decades as a result of a broad 

range of NACA–NASA, Air Force, and foreign research efforts.

Beginning in 1952, the NACA Langley Aeronautical Laboratory began 

an effort oriented to exploring various aspects of fly-by-wire, including 

the use of a side stick controller.15 By 1954, flight-testing began with what 

was perhaps the first jet-powered fly-by-wire research aircraft, a modified 

former U.S. Navy Grumman F9F-2 Panther carrier-based jet fighter used 

as an NACA research aircraft. The primary objective of the NACA effort 

was to evaluate various automatic flight control systems, including those 

based on rate and normal acceleration feedback. Secondary objectives 

were to evaluate use of fly-by-wire with a side stick controller for pilot 

inputs. The existing F9F-2 hydraulic flight control system, with its mechanical linkages, was retained with the NACA designing an auxiliary flight 

control system based on a fly-by-wire analog concept. A small, 4-inch-tall 

side stick controller was mounted at the end of the right ejection seat armrest. The controller was pivoted at the bottom and was used for both lateral (roll) and longitudinal (pitch) control. Only 4 pounds of force were 

required for full stick deflection. The control friction normally present in 

a hydromechanical system was completely eliminated by the electrically 

powered system. Additionally, the aircraft’s fuel system was modified to 

enable fuel to be pumped aft to destabilize the aircraft by moving the center of gravity rearward. Another modification was the addition of a steel 

container mounted on the lower aft fuselage. This carried 250 pounds of 

lead shot to further destabilize the aircraft. In an emergency, the shot could 

be rapidly jettisoned to restabilize the aircraft. Fourteen pilots flew the 

modified F9F-2, including NACA test pilots William Alford16 and Donald 

L. Mallick.17 Using only the side stick controller, the pilots conducted 

15. An electrical side stick controller was developed in Germany during the Second World War to guide 

air-launched Henshel Hs 293 and Ruhrstahl Fritz X missiles against ships maneuvering at sea. Command 

signals were transmitted from the controller to the missiles via radio link. A similar guidance approach 

was used on the surface-to-air Wasserfall (Waterfall) command-guided antiaircraft missile. Examples of 

these weapons are on display at the National Air and Space Museum’s Udvar-Hazy Center at Dulles 

Airport near Washington, DC, and at the National Museum of the U.S. Air Force in Dayton, OH.

16. Bill Alford was killed on Oct. 12, 1959, in the crash of a British Blackburn Buccaneer during 

a visiting test pilot evaluation flight from the Aircraft and Armament Experimental Establishment at 

Boscombe Down in the United Kingdom.

17. Donald L. Mallick, with Peter W. Merlin, “The Smell of Kerosene: A Test Pilot’s Odyssey,” NASA 

SP-4108 (2003).
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takeoffs, stall approaches, acrobatics, and rapid precision maneuvers 

that included air-to-air target tracking, ground strafing runs, and precision approaches and landings. The test pilots quickly became used to 

flying with the side stick and found it comfortable and natural to use.18

In mid-1956, after interviewing aircraft flight control experts from 

the Air Force Wright Air Development Center’s Flight Control Laboratory, 

Aviation Week magazine concluded:

The time may not be far away when the complex mechanical linkage between the pilot’s control stick and the airplane’s 

control surface (or booster valve system) is replaced with 

an electrical servo system. It has long been recognized that 

this“fly-by-wire” approach offered attractive possibilities for 

reducing weight and complexity. However, airplane designers 

and pilots have been reluctant to entrust such a vital function to 

electronics whose reliability record leaves much to be desired.19

Even as the Aviation Week article was published, several noteworthy 

aircraft were under development that would incorporate various fly-bywire approaches in their flight control systems. In 1956, the British Avro 

Vulcan B.2 bomber flew with a partial fly-by-wire system that operated 

in conjunction with hydraulically boosted, mechanically activated flight 

controls. The supersonic North American A-5 Vigilante Navy carrierbased attack bomber flew in 1958 with a pseudo-fly-by-wire flight control 

system. The Vigilante served the fleet for many years, but its highly complex design proved very difficult to maintain and operate in an aircraft 

carrier environment. By the mid-1960s, the General Dynamics F-111 was 

flying with triple-redundant, large-authority stability and command augmentation systems and fly-by-wire-controlled wing-mounted spoilers.20

On the basic research side, the delta winged British Short S.C.1, first 

flown in 1957, was a very small, single-seat Vertical Take-Off and Landing 

18. S.A. Sjoberg, “Some Experience With Side Controllers,” Research Airplane Committee Report 

on the Progress of the X-15 Project, pp. 167–174; Conference held at NACA Langley Aeronautical 

Laboratory, Langley Field, VA, Oct. 25–26, 1956.

19. “Automatic Flight Control System Sought,” Aviation Week, Aug. 6, 1956, pp. 275–284.

20. Duane McRuer and Dunstan Graham, “A Flight Control Century: Triumphs of the Systems 

Approach,” Paper 617, Journal of Guidance, Control and Dynamics, vol. 27, no. 2, pp. 

161–173, AIAA, 2003.
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(VTOL) aircraft. It incorporated a triply redundant fly-by-wire flight control system with a mechanical backup capability. The outputs from the 

three independent fly-by-wire channels were compared, and a failure in a 

single channel was overridden by the other two. A single channel failure 

was relayed to the pilot as a warning, enabling him to switch to the direct 

(mechanical) control system. The S.C.1 had three flight control modes, as 

described below, with the first two only being selectable prior to takeoff.21

• 

• 

• 

Full fly-by-wire mode with aerodynamic surfaces and nozzles controlled electrically via three independent servo 

motors with triplex fail-safe operation in conjunction with 

three analog autostabilizer control systems.

A hybrid mode in which the reaction nozzles were servo/

autostabilizer (fly-by-wire) controlled and the aerodynamic 

surfaces were linked directly to the pilot’s manual controls.

A direct mode in which all controls were mechanically 

linked to the pilot control stick.

The S.C.1 weighed about 8,000 pounds and was powered by four vertically mounted Rolls-Royce RB.108 lift engines, providing a total vertical thrust of 8,600 pounds. One RB.108 engine mounted horizontally 

in the rear fuselage provided thrust for forward flight. The lift engines 

were mounted vertically in side-by-side pairs in a central engine bay 

and could be swiveled to produce vectored thrust (up to 23 degrees forward for acceleration or –12 degrees for deceleration). Variable thrust 

nose, tail, and wingtip jet nozzles (powered by bleed air from the four 

lift engines) provided pitch, roll, and yaw control in hover and at low 

speeds during which the conventional aerodynamic controls were ineffective. The S.C.1 made its first flight (a conventional takeoff and landing) on April 2, 1957. It demonstrated tethered vertical flight on May 

26, 1958, and free vertical flight on October 25, 1958. The first transition from vertical flight to conventional flight was made April 6, 1960.22

21. J.K.B. Illingworth and H.W. Chinn, “Variable Stability and Control Tests on the S.C.1 Aircraft 

in Jet-Borne Flight, with Particular Emphasis on Reference to Desirable VTOL Flying Qualities,” Royal 

Aircraft Establishment, Bedford, U.K., Her Majesty’s Stationery Office, London, 1969.

22. D. Lean and H.W. Chinn, “Review of General Operating Experience with a Jet-Lift VTOL 

Research Aircraft (Short S.C.1),” Aeronautical Research Council Current Paper C.P. No. 832, Her 

Majesty’s Stationery Office, London, 1965.
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During 10 years of flight-testing, the two S.C.1 aircraft made hundreds of flights and were flown by British, French, and NASA test pilots. 

A Royal Aircraft Establishment (RAE) report summarizing flight-test 

experience with the S.C.1 noted: “Of the visiting pilots, those from NASA 

[Langley’s John P. “Jack” Reeder and Fred Drinkwater from Ames] flew 

the aircraft 6 or 7 times each. They were pilots of very wide experience, 

including flight in other VTOL aircraft and variable stability helicopters, 

which was of obvious assistance to them in assessing the S.C.1.”23 On 

October 2, 1963, while hovering at an altitude of 30 feet, a gyro input 

malfunction in the flight control system produced uncontrollable pitch 

and roll oscillations that caused the second S.C.1 test aircraft (XG 905) 

to roll inverted and crash, killing Shorts test pilot J.R. Green. The aircraft was then rebuilt for additional flight-testing. The first S.C.1 (XG 

900) was used for VTOL research until 1971 and is now part of the 

Science Museum aircraft collection at South Kensington, London. The 

second S.C.1 (XG 905) is in the Flight Experience exhibit at the Ulster 

Folk and Transport Museum in Northern Ireland, near where the aircraft was originally built by Short Brothers.

The Canadian Avro CF-105 Arrow supersonic interceptor flew for 

the first time in 1958. Revolutionary in many ways, it featured a dual 

channel, three-axis fly-by-wire flight control system designed without 

any mechanical backup flight control capability. In the CF-105, the 

pilot’s control inputs were detected by pressure-sensitive transducers 

mounted in the pilot’s control column. Electrical signals were sent from 

the transducers to an electronic control servo that operated the valves 

in the hydraulic system to move the various flight control surfaces. The 

CF-105 also incorporated artificial feel and stability augmentation systems.24 In a highly controversial decision, the Canadian government canceled the Arrow program in 1959 after five aircraft had been built and 

flown. Although only about 50 flight test hours had been accumulated, 

the Arrow had reached Mach 2.0 at an altitude of 50,000 feet. During 

its development, NACA Langley Aeronautical Laboratory assisted the 

CF-105 design team in a number of areas, including aerodynamics, 

performance, stability, and control. After the program was terminated, 

23. Illingworth and Chinn, “Variable Stability and Control Tests on the S.C.1 Aircraft in Jet-Borne 

Flight, with Particular Emphasis on Reference to Desirable VTOL Flying Qualities.”

24. J.C. Floyd, “The Canadian Approach to All-Weather Interceptor Development, Fourteenth British 

Commonwealth Lecture,” The Journal of the Royal Aeronautical Society, vol. 62, No. 576, Dec. 1958.
NASA’s Contributions to Aeronautics

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many Avro Canada engineers accepted jobs with NASA and British or 

American aircraft companies.25 Although it never entered production and 

details of its pioneering flight control system design were reportedly little known at the time, the CF-105 presaged later fly-by-wire applications.

NACA test data derived from the F9F-2 fly-by-wire experiment were 

used in development of the side stick controllers in the North American 

X-15 rocket research plane, with its adaptive flight control system.26 First 

flown in 1959, the X-15 eventually achieved a speed of Mach 6.7 and 

reached a peak altitude of 354,200 feet. One of the two side stick controllers in the X-15 cockpit (on the left console) operated the reaction 

thruster control system, critical to maintaining proper attitude control 

at high Mach numbers and extreme altitudes during descent back into 

the higher-density lower atmosphere. The other controller (on the right 

cockpit console) operated the conventional aerodynamic flight control 

surfaces. A CALSPAN NT-33 variable stability test aircraft equipped with 

a side stick controller and an NACA-operated North American F-107A 

(ex-USAF serial No. 55-5120), modified by NACA engineers with a side 

stick flight control system, were flown by X-15 test pilots during 1958–

1959 to gain side stick control experience prior to flying the X-15.27

Interestingly, the British VC10 jet transport, which first flew in 1962, 

has a quad channel flight control system that transmits electrical signals 

directly from the pilot’s flight controls or the aircraft’s autopilot via electrical wiring to self-contained electrohydraulic Powered Flight Control 

Units (PFCUs) in the wings and tail of the aircraft, adjacent to the flight 

control surfaces. Each VC10 PFCU consists of an individual small selfcontained hydraulic system with an electrical pump and small reservoir. 

The PFCUs move the control surfaces based on electrical signals provided to the servo valves that are electrically connected to the cockpit 

flying controls.28 There are no mechanical linkages or hydraulic lines 

between the pilot and the PFCUs. The PFCUs drive the primary flight 

25. Thirty-one former CF-105 engineers were hired by NASA with several going on to hold important positions within the NASA Mercury, Gemini, and Apollo programs. Chris Gainor, “Arrows to 

the Moon: Avro’s Engineers and the Space Race,” (Burlington: Apogee Books, 2001).

26. “Experience with the X-15 Adaptive Flight Control System,” NASA TN-D-6208, NASA Flight 

Research Center, Edwards, CA (March 1971).

27. Hallion, “On the Frontier Flight Research at Dryden, 1946–1981.”

28. Allan F. Damp, “Evaluation Tests on Boulton-Paul VC-10 Aileron Integrated Flight Control Actuator,” Commercial Division, Boeing Company, Renton, WA, Mar. 10, 1970.
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control surfaces that consist of split rudders, ailerons, and elevators on 

separate electrical circuits. Thus, the VC10 has many of the attributes 

of fly-by-wire and power-by-wire flight control systems. It also features 

a backup capability that allows it to be flown using the hydraulically 

boosted variable incidence tail plane and differential spoilers that are 

operated via conventional mechanical linkages and separate hydraulic 

systems.29 The VC10K air refueling tanker was still in Royal Air Force 

(RAF) service as of 2009, and the latest Airbus airliner, the A380, uses 

the PFCU concept in its fly-by-wire flight control system.

The Anglo-French Concorde supersonic transport first flew in 1969 

and was capable of transatlantic sustained supercruise speeds of Mach 

2.0 at cruising altitudes well above 50,000 feet. In support of the Concorde 

development effort, a two-seat Avro 707C delta winged flight research 

aircraft was modified as a fly-by-wire technology testbed with a side 

stick controller. It flew 200 hours on fly-by-wire flight trails at the U.K. 

at Farnborough until September 1966.30 Concorde had a dual channel 

analog fly-by-wire flight control system with a backup mechanical capability. The mechanical system served in a follower role unless problems 

developed with the fly-by-wire control elements of the system, in which 

case it was automatically connected. Pilot movements of the cockpit controls operated signal transducers that generated commands to the flight 

control system. These commands were processed by an analog electrical controller that included the aircraft autopilot. Mechanically operated 

servo valves were replaced by electrically controlled ones. Much as with 

the CF-105, artificial feel forces were electrically provided to the Concorde 

pilots based on information generated by the electronic controller.31

Space Race and the War in Vietnam: Emphasis on FBW Accelerates

During the 1960s, two major events would unfold in the United States 

that had very strong influence on the development and eventual 

29. Molly Neal, “VC10: Vickers-Armstrongs’ Long-range Airliner,” Flight International, May 10, 1962.

30. Ray Sturtivant, “British Research and Development Aircraft: Seventy Years at the Leading Edge,” 

Haynes/Foulis, 1990; Kyrill Von Gersdorff, “Transfer of German Aeronautical Knowledge After 

1945,” in Hirschel, Ernst Heinrich, Horst Prem, and Gero Madelung, “Aeronautical Research in 

Germany (From Lilienthal to Today),” Springer-Verlag, Berlin Heidelberg, 2004.

31. B.S. Wolfe, “The Concorde Automatic Flight Control System: A Description of the Automatic 

Flight Control System for the Anglo-French Supersonic Transport and its Development to Date,” 

Aircraft Engineering and Aerospace Technology, vol. 39, issue 5, 1967.
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introduction into operational service of advanced computer-controlled 

fly-by-wire flight control systems. Early in his administration, President 

John F. Kennedy had initiated the NASA Apollo program with the goal 

of placing a man on the Moon and safely bringing him back to Earth 

by the end of the decade. The space program, and Apollo in particular, 

would lead to major strides in the application of the digital computer to 

manage and control sensors, systems, and advanced fly-by-wire vehicles 

(eventually including piloted aircraft). During the same period, America 

became increasingly involved in the expanding conflict in South Vietnam, 

an involvement that rapidly escalated as the war expanded into a conventional conflict with dimensions far beyond what was originally foreseen. As combat operations intensified in Southeast Asia, large-scale 

U.S. strike missions began to be flown against North Vietnam. In 

response, the Soviet Union equipped North Vietnamese forces with 

improved air defense weapons, including advanced fighters, air-to-air and 

surface-to-air missiles, and massive quantities of conventional antiaircraft 

weapons, ranging in caliber from 12.7 to 100 millimeters (mm). U.S. 

aircraft losses rose dramatically, and American warplane designs came 

under increasing scrutiny as the war escalated.32 Analyses of combat 

data revealed that many aircraft losses resulted from battle damage to 

hydromechanical flight control system components. Traditionally, primary and secondary hydraulic system lines had been routed in parallel through the aircraft structure to the flight control system actuators. 

In the Vietnam combat, experience revealed that loss of hydraulic fluid 

because of battle damage often led to catastrophic fires or total loss 

of aircraft control. Aircraft modification programs were developed to 

reroute and separate primary and secondary hydraulic lines to reduce 

the possibility of a total loss of fluid given a hit. Other changes to existing aircraft flight control systems improved survivability, such as a modification to the F-4 that froze the horizontal tail in the neutral position 

to prevent the aircraft from going out of control when hydraulic fluid 

was lost.33 However, there was an increasing body of opinion that felt a 

32. During the war in Southeast Asia, 8,961 U.S. aircraft were lost. Of these, Air Force losses 

totaled 2,251, the Navy 859, the Marine Corps 463, and the Army 5,388 (mostly helicopters). 

The McDonnell-Douglas F-4 Phantom was the predominant fighter/attack aircraft used by the USAF, 

the USMC, and the Navy during the Vietnam war. A total of 765 F-4s were lost.

33. Robert E. Ball, “The Fundamentals of Aircraft Combat Survivability Analysis and Design,” American Institute of Aeronautics and Astronautics (AIAA), New York, 1985.
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new approach to flight control system design was necessary and technically feasible.

The Air Force JB-47E Fly-By-Wire Project

The USAF Flight Dynamics Laboratory at Wright Patterson Air Force 

Base (AFB), OH, sponsored a number of technology efforts and flighttest programs intended to increase the survivability of aircraft flight 

control system components such as fly-by-wire hydraulic actuators. 

Beginning in 1966, a Boeing B-47E bomber was progressively modified (being redesignated JB-47E) to incorporate analog computercontrolled fly-by-wire actuators for both pitch and roll control, with pilot 

inputs being provided via a side stick controller. The program spanned 

three phases. For Phase I testing, the JB-47E only included fly-by-wire 

in its pitch axis. This axis was chosen because the flight control system in the standard B-47E was known to have a slow response in pitch 

because of the long control cables to the elevator stretching under load. 

Control signals to the pitch axis FBW actuator were generated by a 

transducer attached to the pilot’s control column. The pilot had a simple 

switch in the cockpit that allowed him to switch between the standard 

hydromechanical flight control system (which was retained as a backup) 

and the computer-controlled FBW system. Modified thus, the JB-47E 

flew for the first time, in December 1967. Test pilots reported that the 

modified B-47 had better handling qualities then were attainable with 

the standard B-47E elevator control system, especially in high-speed, 

low-level flight.34

Phase II of the JB-47E program added fly-by-wire roll control and 

a side stick controller that used potentiometers to measure pilot input. 

By the end of the flight-test program, over 40 pilots had flown the FBW 

JB-47E. The Air Force chief test pilot during Phase II, Col. Frank Geisler, 

reported: “In ease of control there is no comparison between the standard 

system and the fly-by-wire. The fly-by-wire is superior in every aspect 

concerning ease of control. . . . It is positive, it is rapid—it responds well—

and best of all the feel is good.”35 Before the JB-47E Phase III flight-test 

program ended in early 1969, a highly reliable four-channel redundant 

34. Gavin D. Jenny, “JB-47E Fly-By-Wire Flight Test Program (Phase I),” Air Force Flight Dynamics 

Laboratory TR-69-40, Wright-Patterson AFB, OH, Sept. 1969.

35. Gavin D. Jenny, James W. Morris, and Vernon R. Schmitt, “Fly-by-Wire, A Historical and Design 

Perspective,” The Society of Automotive Engineers (SAE), 1998.
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electrohydraulic actuator had been installed in the pitch axis and successfully evaluated.36 By this time, the Air Force had already initiated Project 

680J, the Survivable Flight Control System (SFCS), which resulted in 

the prototype McDonnell-Douglas YF-4E Phantom aircraft being modified into a testbed to evaluate the potential benefits of fly-by-wire in a 

high-performance, fighter-type aircraft.37 The SFCS YF-4E was intended 

to validate the concept that dispersed, redundant fly-by-wire flight control elements would be less vulnerable to battle damage, as well as to 

improve the performance of the flight control system and increase overall mission effectiveness.

Project 680J: Survivable Flight Control System YF-4E

In mid-1969, modifications began to convert the prototype McDonnellDouglas YF-4E (USAF serial No. 62-12200) for the SFCS program. A 

quadruple-redundant analog computer-based three-axis fly-by-wire flight 

control system with integrated hydraulic servo-actuator packages was 

incorporated and side stick controllers were added to both the front 

and back cockpits. Roll control was pure fly-by-wire with no mechanical backup. For initial testing, the Phantom’s mechanical flight control 

system was retained in the pitch and yaw axes as a safety backup. On 

April 29, 1972, McDonnell-Douglas test pilot Charles P. “Pete” Garrison 

flew the SFCS YF-4E for the first time from the McDonnell-Douglas factory at Lambert Field in St. Louis, MO. The mechanical flight control 

system was used for takeoff with the pilot switching to the fly-by-wire 

system during climb-out. The aircraft was then flown to Edwards AFB 

for a variety of additional tests, including low-altitude supersonic flights. 

After the first 27 flights, which included 23 hours in the full three-axis flyby-wire configuration, the mechanical flight control system was disabled. 

First flight in the pure fly-by-wire configuration occurred January 22, 

1973. The SFCS YF-4E flew as a pure fly-by-wire aircraft for the remainder of its flight-test program, ultimately completing over 100 flights.38

Whereas the earlier phases of the flight-test effort were primarily 

flown by McDonnell-Douglas test pilots, the next aspect of the SFCS 

36. Tomayko, “Blind Faith: The United States Air Force and the Development of Fly-By-Wire Technology,” Technology and the Air Force: A Retrospective Assessment, Air Force History and Museums 

Program, Washington, DC, 1997.

37. Michael L. Yaffee, “Survivable Controls Gain Emphasis,” Aviation Week, Feb. 2, 1970.

38. Jenny, “JB-47E Fly-By-Wire Flight Test Program (Phase I).”
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program was focused on an Air Force evaluation of the operational 

suitability of fly-by-wire and an assessment of the readiness of the technology for transition into new aircraft designs. During this phase, 15 

flights were accomplished by two Air Force test pilots (Lt. Col. C.W. 

Powell and Maj. R.C. Ettinger), who concluded that fly-by-wire was 

indeed ready and suitable for use in new designs. They also noted that 

flying qualities were generally excellent, especially during takeoffs and 

landings, and that the pitch transient normally encountered in the 

F-4 during rapid deceleration from supersonic to subsonic flight was 

nearly eliminated. Another aspect of the flight-test effort involved socalled technology transition and demonstration flights in the SFCS 

aircraft. At this time, the Air Force had embarked on the Lightweight 

Fighter (LWF) program. One of the two companies developing flight 

demonstrator aircraft (General Dynamics)had elected to use fly-by-wire 

in its new LWF design (the YF-16). A block of 11 flights in the SFCS 

YF-4E was allocated to three pilots assigned to the LWF test force at 

Edwards AFB (Lt. Col. Jim Ryder, Maj. Walt Hersman, and Maj. Mike 

Clarke). Based on their experiences flying the SFCS YF-4E, the LWF test 

force pilots were able to provide valuable inputs into the design, development, and flight test of the YF-16, directly contributing to the dramatic success of that program. An additional 10 flights were allocated 

to another 10 pilots, who included NASA test pilot Gary E. Krier and 

USAF Maj. Robert Barlow.39 Earlier, Krier had piloted the first flight of 

a digital fly-by-wire (DFBW) flight control system in the NASA DFBW 

F-8C on May 25, 1972. That event marked the first time that a piloted 

aircraft had been flown purely using a fly-by-wire flight control system 

without any mechanical backup provisions. Barlow, as a colonel, would 

command the Air Force Flight Dynamics Laboratory during execution of 

several important fly-by-wire flight research efforts. The Air Force YF-16 

and the NASA DFBW F-8 programs are discussed in following sections.

The Precision Aircraft Control Technology YF-4E Program

In order to evaluate the use of computer-controlled fly-by-wire systems 

to actively control a relaxed stability aircraft, the SFCS YF-4E was 

further modified under the Air Force Precision Aircraft Control 

Technology (PACT) program. Movable canard surfaces were mounted 

39. Ibid.
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ahead of the wing and above the YF-4E’s inlets. A dual channel 

electronic fly-by-wire system with two hydraulic systems directed the 

canard actuators. The canards, along with the capability to manage 

internal fuel to move the center of gravity of the aircraft aft, effectively 

reduced the static stability margin to as low as –7.5 percent, that is, fully 

unstable, in the pitch axis. Relaxing static stability by moving the center 

of gravity aft reduces trim drag and decreases the downward force that 

the horizontal tail needs to produce to trim the aircraft. However, as center of gravity moves aft, an aircraft becomes less and less stable in the 

pitch axis, leading to a need to provide artificial stability augmentation. 

A negative static margin means that the aircraft is unstable, it cannot 

maintain stable flight and will be uncontrollable without artificial stability augmentation. During its test program, the PACT aircraft was flown 

34 times, primarily by McDonnell-Douglas test pilots. The Pact aircraft 

demonstrated significant performance gains that included an increase 

in the 4 g maneuvering ceiling of over 4,000 feet (to 50,000 feet) along 

with an increase in turning radius.40 The approach used by the PACT 

YF-4E to create a relaxed stability research aircraft was soon mirrored 

by several foreign flight research programs that are discussed in a separate section. In January 1979, the PACT YF-4E aircraft was delivered 

by Army helicopter from the McDonnell-Douglas factory in St. Louis to 

the Air Force Museum at Wright-Patterson AFB, OH, for permanent display.41 By this time, the Air Force had initiated the Control Configured 

Vehicle (CCV) F-16 project. This was followed by the Advanced Fighter 

Technology Integration (AFTI) F-16 program. Those programs would 

carry forward the fly-by-wire explorations initiated by the PACT YF-4E.

European FBW Research Efforts

By the late 1960s, several European research aircraft using partial 

fly-by-wire flight control systems were in development. In Germany, the 

supersonic VJ-101 experimental Vertical Take-Off and Landing fighter 

technology demonstrator, with its swiveling wingtip mounted afterburning turbojet engines, and the Dornier Do-31 VTOL jet transport 

used analog computer-controlled partial fly-by-wire flight control systems. American test pilots were intimately involved with both programs.

George W. Bright flew the VJ-101 on its first flight in 1963, and NASA test 

40. Ibid.

41. Tomayko, “Computers Take Flight: A History of NASA’s Pioneering Digital Fly-by Wire Project.”
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pilot Drury W. Wood, Jr., headed the cooperative U.S.–German Do-31 

flight-test program that included representatives from NASA Langley 

and NASA Ames. Wood flew the Do-31 on its first flight in February 1967. 

He received the Society of Experimental Test Pilots’ Ivan C. Kinchloe 

Award in 1968 for his role on the Do-31 program.42 By that time, NASA 

test pilot Robert Innis was chief test pilot on the Do-31 program. The 

German VAK-191B VTOL fighter technology flight demonstrator flew 

in 1971. Its triply redundant analog flight control system assisted the 

pilot in operating its flight control surfaces, engines, and reaction control 

nozzles, but the aircraft retained a mechanical backup capability. Later in 

its flight-test program, the VAK-191B was used to support development of 

the partial fly-by-wire flight control system that was used in the multinational Tornado multirole combat aircraft that first flew in August 1974.43

In the U.K., a Hawker Hunter T.12 two-seat jet trainer was converted into a fly-by-wire testbed by the Royal Aircraft Establishment. It 

incorporated a three-axis, quadruplex analog Integrated Flight Control 

System (IFCS) and a “sidearm” controller. The mechanical backup flight 

control system was retained.44 First flown in April 1972, the Hunter was 

eventually lost in a takeoff accident.

In the USSR, a Sukhoi Su-7U two-seat jet fighter trainer was modified with forward destabilizing canards as the Projekt 100LDU fly-bywire testbed. It first flew in 1968 in support of the Sukhoi T-4 supersonic 

bomber development effort. Fitted with a quadruple redundant fly-bywire flight control system with a mechanical backup capability, the fourengine Soviet Sukhoi T-4 prototype first flew in August 1972. Reportedly, 

the fly-by-wire flight control system provided much better handling qualities than the T-4’s mechanical backup system. Four T-4 prototypes were 

built, but only the first aircraft ever flew. Designed for Mach 3.0, the T-4 

never reached Mach 2.0 before the program was canceled after only 10 

test flights and about 10 hours of flying time.45 In 1973–1974, the Projekt 

42. In a Dec. 30, 2006, e-mail, Drury Wood wrote: “I was the project test pilot on this airplane. 

Flew all of the test rigs and full size over 600 flights. Made the record flights, Awarded Bundesverdienstkreuz am Bande and Society of Experimental Test Pilots’ highest award, Kinchloe.”

43. Ulrich Butter, “Control, Navigation, Avionics, Cockpit,” in Ernest Heinrich Hirschel, Horst Prem, 

and Gero Madelung, Aeronautical Research in Germany (From Lilienthal to Today), Springer-Verlag, 

Berlin Heidelberg, 2004.

44. “RAE Electric Hunter,” Flight International, June 28, 1973, pp. 1010–1011.

45. Yefim Gordon, “An Industry of Prototypes: Sukhoi T-4, Russia’s Mach 3 Bomber,” Wings of 

Fame, vol. 9, Aerospace Publishing Limited, London, 1997.
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100LDU testbed was used to support development of the fly-by-wire system flight control system for the Sukhoi T-10 supersonic fighter prototype program. The T-10 was the first pure Soviet fly-by-wire aircraft with 

no mechanical backup; it first flew on May 27, 1977. On July 7, 1978, 

the T-10-2 (second prototype) entered a rapidly divergent pitch oscillation at supersonic speed. Yevgeny Solovyev, distinguished test pilot and 

hero of the Soviet Union, had no chance to eject before the aircraft disintegrated.46 In addition to a design problem in the flight control system, 

the T-10’s aerodynamic configuration was found to be incapable of providing required longitudinal, lateral, and directional stability under all 

flight conditions. After major redesign, the T-10 evolved into the highly 

capable Sukhoi Su-27 family of supersonic fighters and attack aircraft.47

Understanding of FBW Benefits

By the early 1970s, the full range of benefits that could be possible by 

the use of fly-by-wire flight control had become ever more apparent to 

aircraft designers and pilots. Relevant technologies were rapidly maturing, and various forms of fly-by-wire flight control had successfully been 

implemented in missiles, aircraft, and spacecraft. Fly-by-wire had many 

advantages over more conventional flight control systems, in addition 

to those made possible from the elimination of mechanical linkages. A 

computer-controlled fly-by-wire flight control system could generate 

integrated pitch, yaw, and roll control instructions at very high rates to 

maintain the directed flight path. It would automatically provide artificial stability by constantly compensating for any flight path deviations. 

When the pilot moved his cockpit controls, commands would be automatically be generated to modify the artificial stability enough to enable 

the desired maneuvers to be accomplished. It could also prevent the 

pilot from commanding maneuvers that would exceed established aircraft limits in either acceleration or angle of attack. Additionally, the 

46. In an interview published in Krylya Rodiny, No. 8, Aug. 1989, under the title “An Aircraft of 

the 21st Century,” the chief designer at the Sukhoi OKB (Experimental Design Bureau), Mikhail 

Petrovich Simonov, commented: “It is also time to talk about those for whom risk, courage, and a 

willingness to devote their lives to learning the unknown. . . . I am talking about the test pilots of 

our design bureau. . . . Back then we did not know how the frequency responses of control match 

the human capabilities. Zhenya [Solovyev] ended up in a resonant mode and was killed and the 

aircraft was destroyed.”

47. Gordon and Peter Davidson, “Sukhoi Su-27 Flanker,” Specialty Press, 2006.
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flight control system could automatically extend high-lift devices, such 

as flaps, to improve maneuverability.

Conceptual design studies indicated that active fly-by-wire flight 

control systems could enable new aircraft to be developed that featured 

smaller aerodynamic control surfaces. This was possible by reducing 

the inherent static stability traditionally designed into conventional 

aircraft. The ability to relax stability while maintaining good handling 

qualities could also lead to improved agility. Agility is a measure of an 

aircraft’s ability to rapidly change its position. In the 1960s, a concept 

known as energy maneuverability was developed within the Air Force in 

an attempt to quantify agility. This concept states that the energy state of 

a maneuvering aircraft can be expressed as the sum of its kinetic energy 

and its potential energy. An aircraft that possesses higher overall energy 

inherently has higher agility than another aircraft with lower energy. 

The ability to retain a high-energy state while maneuvering requires 

high excess thrust and low drag at high-lift maneuvering conditions.48

Aircraft designers began synthesizing unique conceptual fighter designs 

using energy maneuver theory along with exploiting an aerodynamic 

phenomenon known as vortex lift.49 This approach, coupled with computer-controlled fly-by-wire flight control systems, was felt to be a key 

to unique new fighter aircraft with very high agility levels.

Neutrally stable or even unstable aircraft appeared to be within the 

realm of practical reality and were the subject of ever increasing interest 

and widespread study in NASA and the Air Force, as well as in foreign 

governments and the aerospace industry. Often referred to at the time 

as Control Configured Vehicles, such aircraft could be optimized for 

specific missions with fly-by-wire flight control system characteristics 

48. David C. Aronstein and Albert C. Piccirillo, “The Lightweight Fighter Program: A Successful 

Approach to Fighter Technology Transition,” AIAA, 1996. M.J. Wendl, G.G. Grose, J.L. Porter, 

and V.R. Pruitt, “Flight/Propulsion Control Integration Aspects of Energy Management,” Society of 

Automotive Engineers, Paper No. 740480, 1974.

49. U.S. industry interest in “vortex lift” increased during the early 1960s as a result of NASA Langley aerodynamic studies of foreign delta wing aircraft such as the Concorde supersonic transport 

and the combination canard/delta wing Swedish AJ-37 Viggen fighter. Langley’s vortex lift research 

program was led by Edward Polhamus, with researcher Linwood McKinney studying the favorable 

effects of vortexes on lift produced by strong leading-edge vortex flow off slender lifting surfaces. 

See Chambers, Joseph R., Partners in Freedom: Contributions of the Langley Research Center to U.S. 

Military Aircraft of the 1990s, Monographs in Aerospace History No. 19, NASA, Washington, 

DC, 2000.
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designed to improve aerodynamic performance, maneuverability, and 

agility while reducing airframe weight. Other CCV possibilities included 

the ability to control structural loads while maneuvering (maneuver load 

control) and the potential for implementation of unconventional control 

modes. Maneuver load control could allow new designs to be optimized, 

for example, by using automated control surface deflections to actively 

modify the spanwise lift distribution to alleviate wing bending loads on 

larger aircraft. Unconventional or decoupled control modes would be 

possible by using various combinations of direct-force flight controls 

to change the aircraft flight path without changing its attitude or, alternatively, to point the aircraft without changing the aircraft flight path. 

These unconventional flight control modes were felt at the time to provide an improved ability to point and fire weapons during air combat.50

In summary, the full range of benefits possible through the application of active fly-by-wire flight control in properly tailored aircraft 

design applications was understood to include:

• 

• 

• 

• 

Enhanced performance and improved mission effectiveness made possible by the incorporation of relaxed static 

stability and automatically activated high-lift devices 

into mission-optimized aircraft designs to reduce drag, 

optimize lift, and improve agility and handling qualities throughout the flight and maneuvering envelope.

New approaches to aircraft control, such as the use of 

automatically controlled thrust modulation and thrust 

vectoring fully integrated with the movement of the aircraft’s aerodynamic flight control surfaces and activation of its high-lift devices.

Increased safety provided by automatic angle-of-attack 

and angle-of-sideslip suppression as well as automatic 

limiting of normal acceleration and roll rates. These measures protect from stall and/or loss of control, prevent 

inadvertent overstressing of the airframe, and give the 

pilot maximum freedom to focus on effectively maneuvering the aircraft.

Improved survivability made possible by the elimination 

50. Aronstein and Piccirillo, “The Lightweight Fighter Program: A Successful Approach to Fighter 

Technology Transition.”
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of highly vulnerable hydraulic lines and incorporation of fault tolerant flight control system designs and 

components.

Greatly improved flight control system reliability and 

lower maintenance costs resulting from less mechanical complexity and automated built-in system test and 

diagnostic capabilities.

Automatic flight control system reconfiguration to allow 

safe flight, recovery, and landing following battle damage or system failures.

• 

• 

Digital Fly-By-Wire: The Space Legacy

Both the Mercury and Gemini capsules controlled their reaction control 

thrusters via electrical commands carried by wire. They also used highly 

reliable computers specially developed for the U.S. manned space flight 

program. During reentry from space on his historic 1962 Mercury mission, the first American in space, Alan Shepard, took manual control of 

the spacecraft attitude, one axis at a time, from the automatic attitude 

control system. Using the Mercury direct side controller, he “hand-flew” 

the capsule to the retrofire attitude of 34 degrees pitch-down. Shepard 

reported that he found that the spacecraft response was about the same 

as that of the Mercury simulator at the NASA Langley Research Center.51

The success of fly-by-wire in the early manned space missions gave NASA 

confidence to use a similar fly-by-wire approach in the Lunar Landing 

Research Vehicle (LLRV), built in the early 1960s to practice lunar landing techniques on Earth in preparation for the Apollo missions to the 

Moon. Two LLRVs were built by Bell Aircraft and first flown at Dryden 

in 1964. These were followed by three Lunar Landing Training Vehicles 

(LLTVs) that were used to train the Apollo astronauts. The LLTVs used 

a triply redundant fly-by-wire flight control system based on the use 

of three analog computers. Pure fly-by-wire in their design (there was 

insufficient weight allowance for a mechanical backup capability), they 

proved invaluable in preparing the astronauts for actual landings on the 

surface of the Moon, flying until November 1972.52 A total of 591 flights 

were accomplished, during which one LLRV and two LLTVs crashed in 

51. Robert B. Voas, “Manned Control of Mercury Spacecraft,” Astronautics, vol. 7, no. 3, Mar. 

1962, p. 18.

52. “LLRV Fact Sheet,” FS-2002-09-026-DFRC, NASA Dryden Flight Research Center.
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spectacular accidents but fortunately did so without loss of life.53 During 

this same period, digital computers were demonstrating great improvements in processing power and programmability. Both the Apollo Lunar 

Module and the Command and Service Module used full-authority digital fly-by-wire controls. Fully integrated into the fly-by-wire flight control systems used in the Apollo spacecraft, the Apollo digital computer 

provided the astronauts with the ability to precisely maneuver their vehicles during all aspects of the lunar landing missions. The success of the 

Apollo digital computer in these space vehicles led to the idea of using 

this computer in a piloted flight research aircraft.

By the end of 1969, many experts within NASA and especially at 

the NASA Flight Research Center at Edwards Air Force Base were convinced that digital-computer-based fly-by-wire flight control systems 

would ultimately open the way to dramatic improvements in aircraft 

design, flight safety, and mission effectiveness. A team headed by Melvin 

E. Burke—along with Dwain A. Deets, Calvin R. Jarvis, and Kenneth J. 

Szalai—proposed a flight-test program that would demonstrate exactly 

that. The digital fly-by-wire proposal was evaluated by the Office of 

Advanced Research and Technology (OART) at NASA Headquarters. A 

strong supporter of the proposal was Neil Armstrong, who was by then 

the Deputy Associate Administrator for Aeronautics. Armstrong had 

been the first person to step on the Moon’s surface, in July 1969 during 

the Apollo 11 mission, and he was very interested in fostering transfer 

of technology from the Apollo program into aeronautics applications. 

During discussion of the digital fly-by-wire proposal with Melvin Burke 

and Cal Jarvis, Armstrong strongly supported the concept and reportedly 

commented: “I just went to the Moon with one.” He urged that they contact the Massachusetts Institute of Technology (MIT) Draper Laboratory 

to evaluate the possibility of using modified Apollo hardware and software.54 The Flight Research Center was authorized to modify a fighter 

type aircraft with a digital fly-by-wire system. The modification would 

be based on the Apollo computer and inertial sensing unit.

The NASA Digital Fly-By-Wire F-8 Program

A former Navy F-8C Crusader fighter was chosen for modification, with 

the goal being to both validate the benefits of a digital fly-by-wire aircraft 

53. LLRV No. 2 is on display at the Dryden Flight Research Center.

54. Tomayko, “Computers Take Flight: A History of NASA’s Pioneering Digital Fly-by Wire Project.”
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flight control system and provide additional confidence on its use. Mel 

Burke had worked with the Navy to arrange for the transfer of four 

LTV F-8C Crusader supersonic fighters to the Flight Research Center. 

One would be modified for the F-8 Super Cruise Wing project, one was 

converted into the F-8 DFBW Iron Bird ground simulator, another was 

modified as the DFBW F-8, and one was retained in its basic service configuration and used for pilot familiarization training and general proficiency flying. When Burke left for a job at NASA Headquarters, Cal 

Jarvis, a highly experienced engineer who worked on fly-by-wire systems on the X-15 and LLRV programs, took over as program manager. 

In March 1971, modifications began to create the F-8 DFBW Iron Bird 

simulator. The Iron Bird effort was planned to ensure that development 

of the ground simulator always kept ahead of conversion efforts on the 

DFBW flight-test aircraft. This, the very first F-8C built for the Navy in 

1958 (bureau No. 1445546), carried the NASA tail No. 802 along with 

a “DIGITAL FLY-BY-WIRE” logo painted in blue on its fuselage sides.

F-8 DFBW: Phase I

In implementing the DFBW F-8 program, the Flight Research Center 

chose to remove all the mechanical linkages and cables to the flight 

control surfaces, thus ensuring that the aircraft would be a pure digital 

fly-by-wire system from the start. The flight control surfaces would be 

hydraulically activated, based on electronic signals transmitted via circuits that were controlled by the digital flight control system (DFCS). 

The F-8C’s gun bays were used to house auxiliary avionics, the Apollo 

Display and Keyboard (DSKY) unit,55 and the backup analog flight control system. The Apollo digital guidance computer, its related cooling 

system, and the inertial platform that also came from the Apollo program were installed in what had been the F-8C avionics equipment bay. 

The reference information for the digital flight control system was provided by the Apollo Inertial Management System (IMS). In the conversion of the F-8 to the fly-by-wire configuration, the original F-8 hydraulic 

actuator slider values were replaced with specially developed secondary 

actuators. Each secondary actuator had primary and backup modes. In 

the primary mode, the digital computer sent analog position signals for 

a single actuation cylinder. The cylinder was controlled by a dual self55. The DSKY had been developed for the Apollo program and enabled input and output to the 

digital computer system. It was used during Phase I of the DFBW F-8 program.
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monitoring servo valve. One valve controlled the servo; the other was used 

as a model for comparison. If the position values differed by a predetermined amount, the backup was engaged. In the backup mode, three servo 

cylinders were operated in a three-channel, force-summed arrangement.56

The triply redundant backup analog-computer-based flight control 

system—known as the Backup Control System (BCS)—used an independent power supply and was based on the use of three Sperry analog computers.57 In the event of loss of electrical power, 24-volt batteries could 

keep the BCS running for about 1 hour. Flight control was designed to 

revert to the BCS if any inputs from the primary digital control system 

to the flight control surface actuators did not match up; if the primary 

(digital) computer self-detected internal failures, in the event of electrical power loss to the primary system; and if inputs to secondary actuators were lost. The pilot had the ability to disengage the primary flight 

control system and revert to the BCS using a paddle switch mounted 

on the control column. The pilot could also vary the gains58 to the digital flight control system using rotary switches in the cockpit, a valuable 

feature in a research aircraft intended to explore the development of a 

revolutionary new flight control system.

The control column, rudder pedals, and electrical trim switches 

from the F-8C were retained. Linear Differential Variable Transformers 

(LDVTs) installed in the base of the control stick were used to detect 

pilot control inputs. They generated electrical signals to the flight control system to direct aircraft pitch and yaw changes. Pilot inputs to the 

rudder pedals were detected by LDVTs in the tail of the aircraft. There 

were two LDVTs in each aircraft control axis, one for the primary (digital) flight control system and one for the BCS. The IMS supplied the 

flight control system with attitude, velocity, acceleration, and position 

change references that were compared to the pilot’s control inputs; the 

flight control computer would then calculate required control surface 

position changes to maneuver the aircraft as required.

By the end of 1971, software for the Phase I effort was well along, and 

the aircraft conversion was nearly complete. Extensive testing of the aircraft’s flight control systems was accomplished using the Iron Bird, and 

56. Tomayko, “Computers Take Flight: A History of NASA’s Pioneering Digital Fly-by Wire Project.”

57. These Sperry-developed analog computers were also used in the Air Force’s YF-4E fly-by-wire 

project, which was in progress at the same time as NASA’s DFBW F-8 effort. Ibid.

58. Gain is a measure of the sensitivity of the aircraft to command inputs to the flight control system.
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For the DFBW F-8 program, the Flight Research Center removed all mechanical linkages and 

cables to the flight control surfaces. NASA.

planned test mission profiles were evaluated. On May 25, 1972, NASA 

test pilot Gary Krier made the first flight ever of an aircraft under digital computer control, when he took off from Edwards Air Force Base. 

Envelope expansion flights and tests of the analog BCS followed with 

supersonic flight being achieved by mid-June. Problems were encountered with the stability augmentation system especially, in formation 

flight because of the degree of attention required by the pilot to control 

the aircraft in the roll axis. As airspeeds approached 400 knots, control 

about all axes became too sensitive. Despite modifications, roll axis control remained a problem with lag encountered between control stick 

movement and aircraft response. In September 1972, Tom McMurtry 

flew the aircraft, finding that the roll response was highly sensitive and 

could lead to lateral pilot-induced oscillations (PIOs). By May 1973, 

23 flights had been completed in the Phase I DFBW program. Another 

seven flights were accomplished in June and July, during which different gain combinations were evaluated at various airspeeds.
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In August 1973, the DFBW F-8 was modified to install an YF-16 side 

stick controller.59 It was connected to the analog BCS only. The center 

stick installation was retained. Initially, test flights by Gary Krier and 

Tom McMurtry were restricted to takeoff and landing using the center 

control stick, with transition to the BCS and side stick control being 

made at altitude. Aircraft response and handling qualities were rated 

as highly positive. A wide range of maneuvers, including takeoffs and 

landings, were accomplished by the time the side stick evaluation was 

completed in October 1973. The two test pilots concluded that the YF-16 

side stick control scheme was feasible and easy for pilots to adapt to. 

This inspired high confidence in the concept and resulted in the incorporation of the side stick controller into the YF-16 flight control design. 

Subsequently, four other NASA test pilots flew the aircraft using the 

side stick controller in the final six flights of the DFBW F-8 Phase I 

effort, which concluded in November 1973. Among these pilots was 

General Dynamics chief test pilot Phil Oestricher, who would later fly 

the YF-16 on its first flight in January 1974. The others were NASA test 

pilots William H. Dana (a former X-15 pilot), Einar K. Enevoldson, and 

astronaut Kenneth Mattingly. During Phase I flight-testing, the Apollo 

digital computer maintained its reputation for high reliability and the 

three-channel analog backup fly-by-wire system never had to be used.

DFBW F-8: Phase II

On November 16, 1973, the DFBW team received a NASA group achievement award for its highly impressive accomplishments during the Phase 

I effort. By that time, planning was well underway for the Phase II effort, 

with the first version of the software specification having already been 

issued in April 1973. Whereas Phase I had verified the feasibility of flight 

control using a digital computer, Phase II was intended to develop a more 

practical approach to the implementation of digital flight control, one 

that could be used to justify the incorporation of digital technology into 

production designs for both military and commercial use. In the Phase 

59. As noted earlier, the NACA had evaluated a side stick controller as early as 1952. Side 

stick controllers had been successfully used in the NASA F9F-2, F-107A, and X-15, as well as the 

Mercury, Gemini, and Apollo space vehicles, and they were planned to be used in the upcoming 

Space Shuttle. The Air Force had flight-tested a side stick controller in a B-47E and a C-141 in 

the late 1960s and was planning on its use in the fly-by-wire project 680J YF-4E Survivable Flight 

Control System (SFCS) test aircraft.
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II design, the single channel Apollo computer-based flight control system was replaced with a triply redundant flight control system approach 

using three International Business Machines (IBM) AP-101 digital computers. The challenge was how to program this multicomputer system to 

act as a single computer in processing flight control laws and directing 

aircraft maneuvers while functioning independently for purposes of fault 

tolerance.60 The 32-bit IBM AP-101 computer had been selected for use 

in the Space Shuttle. It consumed 370 watts of power, weighed about 50 

pounds, and had 32,000 words of memory.61 The DFBW program decided 

to also use the AP-101 computer in its Phase II effort, and a purchase 

contact with IBM was signed in August 1973. However, the reliability of 

the AP-101 computer, as measured by mean time between failures, left 

much to be desired. The computer would turn out to require major redesign, and it never came close to meeting its reliability projections. As Ken 

Szalai recently commented: “the IBM AP-101 computer was one of the 

last of the ‘beasts.’ It was big and ran hot. The circuit boards tended to fail 

as temperatures increased. This was found to be due to thermal expansion causing the layers within the circuit boards to separate breaking 

their electrical connections.” Szalai recounted that he notified the Space 

Shuttle team as soon as the issue was discovered. They were surprised, as 

they had never seen a similar problem with the AP-101. The reason soon 

became apparent. The AP-101s installed in the F-8 Iron Bird were being 

tested in a non–air-conditioned hangar; Space Shuttle flight control system testing had been in a 50 degree Fahrenheit (50 °F) cooled laboratory 

environment. When the Space Shuttle was tested on the flight line in typical outside temperatures encountered at Dryden, similar reliability problems were encountered. IBM subsequently changed the thermal coating 

process used in the manufacture of the AP-101 circuit boards, a measure 

that partly resolved the AP-101’s reliability problems.62

Software for Phase II was also larger and more complex than that 

used in Phase I because of the need for new pilot interface devices. Flight 

control modes still included the direct (DIR) mode, the stability augmentation (SAS) mode, and the control-augmentation (CAS) mode. A 

pitch maneuver-load-control feature was added to the CAS mode, and a 

digital autopilot was fitted that incorporated Mach hold, altitude-hold, 

60. Kenneth J. Szalai, telephone conversation with author, Mar. 11, 2009.

61. Tomayko, “Computers Take Flight: A History of NASA’s Pioneering Digital Fly-by Wire Project.”

62. Szalai, telephone conversation.
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and heading-hold selections. The software gradually matured to the point 

where pilots could begin verification evaluations in the Iron Bird simulator in early 1976. By July, no anomalies were reported in the latest 

software release, with the direct and stability-augmentation modes considered flight-ready. The autopilot and control-augmentation mode still 

required more development, but they were not necessary for first flight.

The backup analog flight control system was also redesigned for 

Phase II, and the secondary actuators were upgraded. Sperry supplied 

an updated version of the Phase I Backup Control System using the same 

technology that had been used in the Air Force’s YF-4E project. Signals 

from the analog computers were now force-summed when they reached 

the actuators, resulting in a quicker response. The redesigned secondary 

actuators provided 20 percent more force, and they were also more reliable. The hydraulic actuators used in Phase I had two sources of hydraulic 

pressure for the actuators; in those chosen for Phase II, there were three 

hydraulic sources that corresponded with the three channels in each of 

the primary and secondary flight control systems. The secondary electronic actuators had three channels, with one dedicated to each computer 

in the primary system. The actuators were shared by the analog computer bypass system in the event of failure of the primary digital system.

The final Phase II design review occurred in late May 1975, with both 

the Iron Bird and aircraft 802 undergoing modification well into 1976. By 

early April, Gary Krier was able to fly the Iron Bird simulator with flight 

hardware and software. Handling qualities were generally rated as very 

good, but actuator anomalies and transients were noted, as were some 

problems with the latest software releases. After these issues were resolved, 

a flight qualification review was completed on August 20. High-speed taxi 

tests beginning 3 days later, then, on August 27, 1976, Gary Krier took off 

on the first flight of the Phase II program. On the second Phase II flight, 

one of the AP-101 computers failed with the aircraft at supersonic speed. 

An uneventful landing was accomplished with the flight control system 

remaining in the primary flight control mode. This was in accordance 

with the established flight-test procedure in the event of a failure of one 

of the primary computers. Flight-testing was halted, and all AP-101s were 

sent back to IBM for refurbishment. After 4 months, the AP-101s were 

back at Dryden, but another AP-101 computer failure occurred on the 

very next flight. Again, the primary digital flight control system handled 

the failure well, and flights were soon being accomplished without incident, providing ever increasing confidence in the system.
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In the spring of 1977, the DFBW F-8 was modified to support the 

Space Shuttle program. It flew eight times with the Shuttle Backup Flight 

System’s software test package running in parallel with the F-8 flight 

control software. Data from this package were downlinked as the F-8 

pilots flew a series of simulated Shuttle landing profiles. Later in 1977, 

the unpowered Space Shuttle Enterprise was being used to evaluate the 

flight characteristics of the Space Shuttle during approach and landing 

in preparation for full-up shuttle missions. During the Shuttle Approach 

and Landing Test (ALT) program, the Enterprise was carried aloft atop 

the NASA 747 Shuttle carrier aircraft. After release, the Shuttle’s handling qualities and the responsiveness of its digital fly-by-wire system 

were evaluated. On the fifth and last of the shuttle ALT flights in October 

1977, a pilot-induced oscillation developed just as the Enterprise was 

landing. The DFBW F-8C was then used in a project oriented to duplicating the PIO problem encountered on the Shuttle during a series of 

flights in 1978 that were initially flown by Krier and McMurtry. They 

were joined by Einar K. Enevoldson and John A. Manke, who had extensive experience flying NASA lifting body vehicles. The lifting body vehicles used side stick controllers and had approach characteristics that 

were similar to those of the Space Shuttle.

Flying simulated Shuttle landing profiles with the DFBW F-8, the 

pilots gathered extremely valuable data that supported the Shuttle program in establishing sampling rates and control law execution limits. 

The DFBW F-8 flight control software had been modified to enable the 

pilot to vary transport delay times to evaluate their effect on control 

response. Transport delay is the elapsed time between pilot movement 

of his cockpit control and the actual movement of the flight control surfaces. It is a function of several factors, including the time needed to do 

analog-to-digital conversion, the time required to execute the appropriate flight control law, length of the electrical wires to the actuators, and 

the lag in response of the hydraulic system. If transport delay is too long, 

the pilot may direct additional control surface movement while his initial commands are in the process of being executed by the flight control 

system. This can result in overcontrol. Subsequent attempts to correct 

the overshoot can lead to a series of alternating overshoots or oscillations that are commonly referred to as a PIO. The range of transport 

delay times within which the Shuttle would be unlikely to encounter a 

PIO was determined using the DFBW F-8, enabling Dryden to develop 

a PIO suppression filter for the Shuttle. The PIO suppression filter was 
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successfully evaluated in the F-8, installed in the Shuttle prior to its first 

mission into space, and proved to effectively eliminate the PIO issue.63

During Phase II, 169 flights were accomplished with several other 

test pilots joining the program, including Stephen D. Ishmael, Rogers 

Smith, and Edward Schneider. In addition to its previously noted accomplishments, the DFBW F-8 successfully evaluated adaptive control law 

approaches that would later become standard in many FBW aircraft. 

It was used in the Optimum Trajectory Research Experiment (OPTRE). 

This involved testing data uplink and downlink between the F-8 and a 

computer in the then-new Remotely Piloted Vehicle Facility. This experiment demonstrated that an aircraft equipped with a digital flight control system could be flown using control laws that were operating in 

ground-based digital computers. The F-8 conducted the first in-flight evaluations of an automatic angle-of-attack limiter and maneuvering flaps. 

These features are now commonly used on nearly all military and commercial aircraft with fly-by-wire flight controls. The DFBW F-8 also successfully tested an approach that used a backup software system known 

as the Resident Backup System (REBUS) to survive potential software 

faults that could cause all three primary flight control system computers to fail. The REBUS concept was later used in other experimental 

aircraft, as well as in production fly-by-wire flight control systems. The 

final flight-test effort of the DFBW program involved the development 

of a methodology called analytical redundancy management. In this 

concept, dynamic and kinematic relationships between dissimilar sensors and measurements were used to detect and isolate sensor failures.64

DFBW F-8: An Appreciation

The NASA DFBW F-8 had conclusively proven that a highly redundant digital flight control system could be successfully implemented 

and all aspects of its design validated.65 During the course of the program, the DFBW F-8 demonstrated the ability to be upgraded to take 

advantage of emerging state-of-the-art technologies or to meet evolv63. Szalai, Calvin R. Jarvis, Gary E. Krier, Vincent A. Megna, Larry D. Brock, and Robert N. 

O’Donnell, “Digital Fly-by-Wire Flight Control Validation Experience,” NASA TM-72860 (Dec. 1978). 

R.E. Bailey, M.F. Schaeffer, R.E. Smith, and J.F. Stewart, “Flight Test Experience With Pilot-InducedOscillation Suppression Filters,” NASA TM-86028, NASA Dryden Research Center (Jan. 1984).

64. Digital Fly-By-Wire, “The All-Electric Airplane,” NASA Dryden TF-2001-02 DFRC.

65. Szalai, et al., “Digital Fly-By-Wire Flight Control Validation Experience.”
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ing operational requirements. It proved that digital fly-by-wire flight 

control systems could be adapted to the new design and employment 

concepts that were evolving in both the military and in industry at the 

time. Perhaps the best testimony to the unique accomplishments of 

the F-8 DFBW aircraft and its NASA flight-test team is encapsulated in 

the following observations of former NASA Dryden director Ken Szalai:

DFBW systems are ‘old hat’ today, but in 1972, only Apollo 

astronauts had put their life and missions into the hands of 

software engineers. We considered the F-8 DFBW a very high 

risk in 1972. That fact was driven home to us in the control 

room when we asked the EAFB [Edwards Air Force Base] 

tower to close the airfield, as was preplanned with the USAF, 

for first flight. It was the first time this 30-year-old FCS [Flight 

Control System] engineer had heard that particular radio 

call. . . . The project was both a pioneering effort for the technology and a key enabler for extraordinary leaps in aircraft 

performance, survivability, and superiority. The basic architecture has been used in numerous production systems, and 

many of the F-8 fault detection and fault handling/recovery 

technology elements have become ‘standard equipment.’ . . . In 

the total flight program, no software error/fault ever occurred 

in the operational software, synchronization was never lost in 

hundreds of millions of sync cycles, it was never required to 

transfer to the analog FBW backup system, there were zero 

nuisance channel failures in all the years of flying, and many 

NASA and visiting guest pilots easily flew the aircraft, including Phil Oestricher before the first YF-16 flight.66

In retrospect, the NASA DFBW F-8C is of exceptional interest in the 

history of aeronautics. It was the first aircraft to fly with a digital fly-bywire flight control system, and it was also the first aircraft to fly without any mechanical backup flight controls. Flown by Ed Schneider, the 

DFBW F-8 made its last flight December 16, 1985, completing 211 flights. 

The aircraft is now on display at the NASA Dryden Flight Research 

Center. Its sustained record of success over a 13-year period provided a 

66. Szalai, e-mail to the author, Mar. 11, 2009.
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high degree of high confidence in the use of digital computers in fly-bywire flight control systems. The DFBW F-8C also paved the way for a 

number of other significant NASA, Air Force, and foreign research programs that would further explore and expand the application of digital 

computers to modern flight control systems, providing greatly improved 

aircraft performance and enhanced flight safety.

The Lightweight Fighter Program and the YF-16

In addition to the NASA F-8 DFBW program, several other highly noteworthy efforts involving the use of computer-controlled fly-by-wire flight 

control technology occurred during the 1970s. The Air Force had initiated the Lightweight Fighter program in early 1972. Its purpose was “to 

determine the feasibility of developing a small, light-weight, low-cost 

fighter, to establish what such an aircraft can do, and to evaluate its possible operational feasibility.”67 The LWF effort was focused on demonstrating technologies that provided a direct contribution to performance, 

were of moderate risk (but sufficiently advanced to require prototyping 

to reduce risk), and helped hold both procurement and operating costs 

down. Two companies, General Dynamics (GD) and Northrop, were 

selected, and each was given a contract to build two flight-test prototypes. 

These would be known as the YF-16 and the YF-17. In its YF-16 design, 

GD chose to use an analog-computer-based quadruplex fly-by-wire flight 

control system with no mechanical backup. The aircraft had been 

designed with a negative longitudinal static stability margin of between 

7 percent and 10 percent in subsonic flight—this indicated that its center 

of gravity was aft of the aerodynamic center by a distance of 7 to 10 

percent of the mean aerodynamic chord of the wing. A high-speed, 

computer-controlled fly-by-wire flight control system was essential to provide the artificial stability that made the YF-16 flyable. The aircraft also 

incorporated electronically activated and electrically actuated leading edge 

maneuvering laps that were automatically configured by the flight control system to optimize lift-to-drag ratio based on angle of attack, Mach 

number, and aircraft pitch rate. A side stick controller was used in place 

of a conventional control column.68

67. “General Dynamics and Northrop to Build Lightweight Fighter Prototypes,” Interavia, July 1972, 

p. 693.

68. C. Droste and J. Walker, “The General Dynamics Case Study on the F-16 Fly-By-Wire Flight 

Control System,” AIAA Professional Study Series, AIAA, New York, June 1998.
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Following an exceptionally rapid development effort, the first of 

the two highly maneuverable YF-16 technology demonstrator aircraft 

(USAF serial No. 72-1567) had officially first flown in February 1974, 

piloted by General Dynamics test pilot Phil Oestricher. However, an unintentional first flight had actually occurred several weeks earlier, an event 

that is discussed in a following section as it relates to developmental 

issues with the YF-16 fly-by-wire flight control system. During its development, NASA had provided major assistance to GD and the Air Force 

on the YF-16 in many technical areas. Fly-by-wire technology and the 

side stick controller concept originally developed by NASA were incorporated in the YF-16 design. The NASA Dryden DFBW F-8 was used as 

a flight testbed to validate the YF-16 side stick controller design. NASA 

Langley also helped solve numerous developmental challenges involving 

aerodynamics and control laws for the fly-by-wire flight control system. 

The aerodynamic configuration had been in development by GD since 

1968. Initially, a sharp-edged strake fuselage forebody had been eliminated from consideration because it led to flow separation; however, 

rounded forward fuselage cross sections caused significant directional 

instability at high angles of attack. NASA aerodynamicists conducted 

wind tunnel tests at NASA Langley that showed the vortexes generated 

by sharp forebody strakes produced a more stable flow pattern with 

increased lift and improved directional stability. This and NASA research 

into leading – and trailing-edge flaps were used by GD in the development of the final YF-16 configuration, which was intensively tested in 

the Langley Full-Scale Wind Tunnel at high angle-of-attack conditions.69

During NASA wind tunnel tests, deficiencies in stability and control, 

deep stall, and spin recovery were identified even though GD had predicted the configuration to be controllable at angles of attack up to 36 

degrees. NASA wind tunnel testing revealed serious loss of directional 

stability at angles of attack higher than 25 degrees. As a result, an automatic angle of attack limiter was incorporated into the YF-16 flight control system along with other changes designed to address deep stall and 

spin issues. Ensuring adequate controllability at higher angles of attack 

also required further research on the ability of the YF-16’s fly-by-wire 

flight control system to automatically limit certain other flight parameters during energetic air combat maneuvering. The YF-16’s all-moving 

69. Chambers, Partners in Freedom.
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horizontal tails provided pitch control and also were designed to operate differentially to assist the wing flaperons in rolling the aircraft. The 

ability of the horizontal tails and longitudinal control system to limit 

the aircraft’s angle of attack during maneuvers with high roll rates at 

low airspeeds was critically important. Rapid rolling maneuvers at low 

airspeeds and high angles of attack were found to create large nose-up 

trim changes because of inertial effects at the same time that the aerodynamic effectiveness of the horizontal tails was reduced.70

An important aspect of NASA’s support to the YF-16 flight control system development involved piloted simulator studies in the NASA Langley 

Differential Maneuvering Simulator (DMS). The DMS provided a realistic means of simulating two aircraft or spacecraft operating with (or 

against) each other (for example, spacecraft conducting docking maneuvers or fighters engaged in aerial combat against each other). The DMS 

consisted of two identical fixed-base cockpits and projection systems, 

each housed inside a 40-foot-diameter spherical projection screen. Each 

projection system consisted of a sky-Earth projector to provide a horizon reference and a system for target-image generation and projection. 

The projectors and image generators were gimbaled to allow visual simulation with completely unrestricted freedom of motion. The cockpits 

contained typical fighter cockpit instruments, a programmable buffet 

mechanism, and programmable control forces, plus a g-suit that activated automatically during maneuvering.71 Extensive evaluations of the 

YF-16 flight control system were conducted in the DMS using pilots 

from NASA, GD, and the Air Force, including those who would later fly 

the aircraft. These studies verified the effectiveness of the YF-16 fly-bywire flight control system and helped to identify critical flight control 

system components, timing schedules, and feedback gains necessary 

to stabilize the aircraft during high angle-of-attack maneuvering. As a 

result, gains in the flight control system were modified, and new con70. Later, during high angle-of-attack flight-testing of an early production F-16, the aircraft entered a 

stabilized deep-stall condition following a series of rolls in a vertical climbing maneuver. The test pilot 

was unable to recover with normal aerodynamic controls and used the anti-spin parachute installed for 

high angle-of-attack testing. NASA worked with the Air Force and the contractor to develop a fix that 

involved a “pitch rocker” technique to force the aircraft out of the deep stall. The approach was incorporated into the production F-16 flight control system as a pilot selectable emergency recovery mode. In 

addition, the horizontal tail area of the production F-16 was increased by about 25 percent. Ibid.

71. B.R. Ashworth and William M. Kahlbaum, Jr., “Description and Performance of the Langley Differential Maneuvering Simulator,” NASA TN-D-7304, NASA Langley Research Center (June 1973).
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trol elements—such as a yaw rate limiter, a rudder command fadeout, 

and a roll rate limiter—were developed and evaluated.72

Despite the use of the DMS and the somewhat similar GD Fort 

Worth domed simulator to develop and refine the YF-16 flight control 

system, nearly all flight control functions, including roll stick force gradient, were initially too sensitive. This contributed to the unintentional 

YF-16 first flight by Phil Oestricher at Edwards AFB on January 20, 

1974. The intent of the scheduled test mission on that day was to evaluate the aircraft’s pretakeoff handling characteristics. Oestricher rotated 

the YF-16 to a nose-up attitude of about 10 degrees when he reached 

130 knots, with the airplane still accelerating slightly. He made small 

lateral stick inputs to get a feel for the roll response but initially got no 

response, presumably because the main gear were still on the ground. 

At that point, he slightly increased angle of attack, and the YF-16 lifted 

off the ground. The left wing then dropped rather rapidly. After a right 

roll command was applied, it went into a high-frequency pilot-induced 

oscillation. Before the roll oscillation could be stopped, the aft fin of 

the inert AIM-9 missile on the left wingtip lightly touched the runway, 

the right horizontal tail struck the ground, and the aircraft bounced on 

its landing gear several times, resulting in the YF-16 heading toward 

the edge of the runway. Oestricher decided to take off, believing it 

impossible to stay on the runway. He touched down 6 minutes later 

and reported: “The roll control was too sensitive, too much roll rate as 

a function of stick force. Every time I tried to correct the oscillation, 

I got full-rate roll.” The roll control sensitivity problem was corrected 

with adjustments to the control gain logic. Stick force gradients and 

control gains continued to be refined during the flight-test program, 

with the YF-16 subsequently demonstrating generally excellent control 

characteristics. Oestricher later said that the YF-16 control problem 

would have been discovered before the first flight if better visual 

displays had been available for flight simulators in the early 1970s.73

Lessons from the YF-16 and DFBW F-8 simulation experiences helped 

NASA, the Air Force, and industry refine the way that preflight 

simulation was structured to support new fly-by-wire flight control 

systems development. Another flight control issue that arose during 

72. Chambers, Partners in Freedom.

73. Joe Stout, “What a Wonderful Airplane: YF-16 First Flight,” Code One Magazine, General 

Dynamics, July 1992.
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the YF-16 flight-test program involved an instability caused by interaction of the active fly-by-wire flight control system with the aeroelastic properties of the airframe. Flutter analysis had not accounted for 

the effects of active flight control. Closed loop control systems testing on the ground had used simulated aircraft dynamics based on a 

rigid airframe modeling assumption. In flight, the roll sensors detected 

aeroelastic vibrations in the wings, and the active flight control system 

attempted to apply corrective roll commands. However, at times these 

actually amplified the airframe vibrations. This problem was corrected 

by reducing the gain in the roll control loop and adding a filter in the 

feedback patch that suppressed the high-frequency signals from structural vibrations. The fact that this problem was also rapidly corrected 

added confidence in the ability of the fly-by-wire flight control system 

to be reconfigured. Another change made as a result of flight test was 

to fit a modified side stick controller that provided the pilot with some 

small degree of motion (although the control inputs to the flight control system were still determined by the amount of force being exerted 

on the side stick, not by its position).74

Three days after its first official flight on February 2, 1974, the YF-16 

demonstrated supersonic windup turns at Mach 1.2. By March 11, it 

had flown 20 times and achieved Mach 2.0 in an outstanding demonstration of the high systems reliability and excellent performance that 

could be achieved with a fly-by-wire flight control system. By the time 

the 12-month flight-test program ended January 31, 1975, the two YF-16s 

had flown a total of 439 flight hours in 347 flights, with the YF-16 Joint 

Test Force averaging over 30 sorties per month. Open communications 

between NASA, the Air Force, and GD had been critical to the success of 

the YF-16 development program. In highlighting this success, Harry J. 

Hillaker, GD Vice President and Deputy Program Director for the F-16, 

noted the vital importance of the “free exchange of experience from the 

U.S. Air Force Laboratories and McDonnell-Douglas 680J projects on 

the F-4 and from NASA’s F-8 fly-by-wire research program.”75 The YF-16 

would serve as the basis for the extremely successful family of F-16 mul74. Aronstein and Piccirillo, “The Lightweight Fighter Program: A Successful Approach to Fighter 

Technology Transition.”

75. H.J. Hillaker, “The F-16: A Technology Demonstrator, a Prototype, and a Flight Demonstrator,” 

Proceedings of AIAA Aircraft Prototype and Technology Demonstrator Symposium, AIAA, New York, 

1983, pp. 113–120.
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tinational fighters; over 4,400 were delivered from assembly lines in 

five countries by 2009, and production is expected to continue to 2015. 

While initial versions of the production F-16 (the A and B models) used 

analog computers, later versions (starting with the F-16C) incorporated 

digital computers in their flight control systems.76 Fly-by-wire and relaxed 

static stability gave the F-16 a major advantage in air combat capability over conventional fighters when it was introduced, and this technology still makes it a viable competitor today, 35 years after its first flight.

The F-16’s main international competition for sales at the time was 

another statically unstable full fly-by-wire fighter, the French Dassault 

Mirage 2000, which first flew in 1978. Despite the F-16 being selected 

for European coproduction, over 600 Mirage 2000s would also eventually be built and operated by a number of foreign air forces. The other 

technology demonstrator developed under the LWF program was the 

Northrop YF-17. It featured a conventional mechanical/hydraulic flight 

control system and was statically stable. When the Navy decided to build 

the McDonnell-Douglas F/A-18, the YF-17 was scaled up to meet fleet 

requirements. Positive longitudinal static stability was retained, and a primary fly-by-wire flight control system was incorporated into the F/A-18’s 

design. The flight control system also had an electric backup that enabled 

the pilot to transmit control inputs directly to the control surfaces, bypassing the flight control computer but using electrical rather than mechanical transmission of signals. A second backup provided a mechanical 

linkage to the horizontal tails only. These backup systems were possible 

because the F/A-18, like the YF-17, was statically stable about all axes.77

F-16 CCV

By the mid-1970s, the Air Force Flight Dynamics Laboratory had initiated 

a Control Configured Vehicle flight research program to investigate the 

use of nonconventional (often called “decoupled”) movements of aircraft 

flight control surfaces to enable maneuvers in one plane without movement in another. An example of such a maneuver would be a wings-level 

turn without having to bank the aircraft. The very first General Dynamics 

YF-16 technology demonstrator aircraft (USAF serial No. 72-1567) was 

selected for modification. Rebuilt in December 1975 and fitted with twin 

76. Aronstein and Piccirillo, “The Lightweight Fighter Program: A Successful Approach to Fighter 

Technology Transition.”

77. Ibid.
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vertical canards underneath the air intake, it became known as the 

F-16/CCV. Its flight controls were modified to enable the wing trailing edge flaperons to move in combination with the all-moving stabilator. In addition, the fuel system in the YF-16 was modified to enable 

the aircraft center of gravity to be adjusted in flight by transferring fuel 

between tanks, thus allowing the stability of the aircraft to be varied. 

The YF-16/CCV flew for the first time on March 16, 1976, piloted by GD 

test pilot David J. Thigpin. On June 24, 1976, while being flown by David 

Thigpin, it was seriously damaged in a crash landing. Engine power had 

been lost on final approach, and the landing gear collapsed in the subsequent hard landing. Repairs to the aircraft would take over 6 months. 

The F-16 CCV returned to flight in the spring of 1977. It would complete 

87 flights and 125 flying hours before the research program ended, with 

the last flight F-16 CCV flight on June 30, 1977.78

The F-16 CCV anticipated CCV flight-test approaches that were soon 

undertaken by a number of foreign countries. Flight research projects in 

Germany, the U.K., and Japan converted existing military aircraft into 

CCV testbeds. Fitted with computer-controlled fly-by-wire flight control 

systems, these projects provided experience and insights that enabled 

these countries to incorporate fly-by-wire flight control into their next 

generations of advanced civil and military aircraft. These foreign CCV 

projects are discussed in a separate section of this report. Experience 

gained with the F-16 CCV served as the basis for the subsequent Flight 

Dynamics Laboratory AFTI/F-16 program, which would yield valuable 

insights into many issues associated with developing advanced DFBW 

flight control systems and result in significantly improved capabilities 

being incorporated into U.S. military aircraft.

YC-14

The Air Force Boeing YC-14 Short Take-Off and Landing (STOL) jet 

transport technology demonstrator flew for the first time on August 

9, 1976, from Boeing Field in Seattle, WA, during the period between 

Phase I and Phase II of the NASA DFBW F-8 program. Two prototypes were built with the second aircraft flying in October 1976. The 

YC-14 is noteworthy in that it was the first aircraft to fly with a faulttolerant multichannel redundant digital fly-by-wire flight control system. 

78. Stout, “What a Wonderful Airplane: YF-16 First Flight.” Joseph F. Baugher, “General Dynamics 

YF-16/CCV,” American Military Aircraft, Mar. 31, 2000.
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A mechanical backup flight control capability was retained. The full 

authority triply redundant digital fly-by-wire flight control system, 

designed by the British Marconi Company, performed computational 

commands for pitch, roll, and yaw that were used to control the elevator, aileron, and rudder actuation systems. The reconfigurable computer 

architecture divided the basic control path into three subfunctional elements with these elements replicated to provide fault tolerance. The internal element redundancy management function was intended to detect 

and isolate faulty elements and perform the necessary reconfiguration. 

The input signal selection methodology was intended to guarantee that 

all three computers used the same numbers and thus produced identical output values. During normal operation, the overall system output 

value was selected as the midvalue of the three individual values. The 

system would continue to operate in the event of a failure of one computer by taking the average of the output of the two remaining computers. If they disagreed, both were disabled and the aircraft reverted to 

the backup manual control system.79

The YC-14 was also noteworthy in that it used optical data links to 

exchange data between the triply redundant computers. The optical communications medium was chosen to eliminate electromagnetic interference effects, electrical grounding loop problems, and the potential 

propagation of electrical malfunctions between channels. Optical coupling was used to maintain interchannel integrity. Each sensor’s output was coupled to the other channels so that each computer had data 

from each of the other sensors. Identical algorithms in each computer 

were used. They consolidated the data, enabling equalization and fault 

detection/isolation of the inputs. The computers were synchronized to 

avoid sampling time differences and to assure that all computers were 

receiving identical data inputs.80

An important observation involving redundant computer-controlled 

fly-by-wire flight control systems was derived from the YC-14 flight-test 

experience. As noted above, the system was designed to ensure that 

all computers used the same sensor input values and should therefore 

produce identical outputs. However, a significant fault in the digital 

79. L. Martin and D. Gangsaas, “Testing the YC-14 Flight Control System Software,” AIAA Journal 

of Guidance and Control, July–Aug. 1978.

80. H.A. Rediess and E.C. Buckley, “Technology Review of Flight Crucial Flight Control Systems 

(Application of Optical Technology),” NASA CR-172332 (Supplement 1) (Sept. 1984).
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flight control software was encountered during flight-testing that had 

not been detected during ground laboratory testing. The software fault 

resulted in incorrect tracking of control law computations in each of 

the three flight control channels, with each channel performing signal 

selections on a different set of values. This resulted in different input 

data for the three channels. Although the discrepancies between each 

channel’s inputs were small, the cumulative effect led to large tracking 

errors between flight control channels when airborne.81

Following cancellation of the Air Force YC-14A program in 1979, 

the two prototypes were placed in storage at Air Force’s Aerospace 

Maintenance and Regeneration Group (AMARC) at Davis-Monthan 

AFB, AZ, in April 1980. The first prototype is now displayed at the Pima 

Air and Space Museum in Tucson, AZ.

Advanced Fighter Technology Integration F-16 Program

The USAF Flight Dynamics Laboratory began the Advanced Fighter 

Technology Integration program in the late 1970s. Overall objectives of 

this joint Air Force and NASA research program were to develop and 

demonstrate technologies and assess alternative approaches for use in 

future aircraft design. In December 1978, the F-16 was selected for modification as the AFTI/F-16. General Dynamics began conversion of the 

sixth preproduction F-16A (USAF serial No. 75-0750) at its Forth Worth, 

TX, factory in March 1980. The aircraft had originally been built in 1978 

for the F-16 full-scale development effort. GD built on earlier experience 

with its F-16 CCV program. The twin canted movable canard ventral fins 

from the F-16 CCV were installed under the inlet of the AFTI/F-16. In 

addition, a dorsal fairing was fitted to the top of the fuselage to accommodate extra avionics equipment. A triply redundant, asynchronous, 

multimode, digital flight control system with an analog backup was 

installed in the aircraft. The DFCS was integrated with improved avionics and had different control modes optimized for air-to-air combat 

and air-to-ground attack. The Stores Management System (SMS) 

was responsible for signaling requests for mode change to the DFCS. 

Other modifications included provision for a six-degree-of-freedom 

Automated Maneuvering Attack System (AMAS), a 256-word-capacity 

Voice-Controlled Interactive Device (VCID) to control the avionics 

81. Martin, et al., “Testing the YC-14 Flight Control System Software.”
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During Phase I, five test pilots from NASA, the Air Force, and the Navy flew the AFTI/F-16 at 

NASA Dryden in California. NASA.

suite, and a helmet-mounted target designation sight that could automatically slave the forward-looking infrared (FLIR) device and the radar 

to the pilot’s head movements.82 First flight of the modified aircraft in 

the AFTI/F-16 configuration occurred on July 10, 1982, from Carlswell 

AFB, TX, with GD test pilot Alex V. Wolfe at the controls. Following contractor testing, the aircraft was flown to Edwards AFB for AFTI/F-16 test 

effort. This was organized into two phases; Phase I was a 2-year effort 

focused on evaluating the DFCS, with a follow-on Phase II oriented to 

assessing the AMAS and other technologies.

AFTI Phase I Testing

Phase I flight-testing was conducted by the AFTI/F-16 Joint Test Force 

from the NASA Dryden Flight Research Facility at Edwards AFB, CA, from 

July 10, 1982, through July 30, 1983. During this phase, five test pilots 

from NASA, the Air Force, and the U.S. Navy flew the aircraft. Initial 

flights checked out the aircraft’s stability and control systems. Handling 

qualities were assessed in air-to-air and air-to-ground scenarios, as well 

82. Stephen D. Ishmael and Donald R. McMonagle, “AFTI/F-16 Flight Test Results and Lessons,” 

NASA TM-84920 (Oct. 1983).
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as in-formation flight and during approach and landing. The Voice 

Command System allowed the pilot to change switch positions, display 

formats, and modes simply by saying the correct word. Initial tests were 

of the system’s ability to recognize words, with later testing conducted 

under increasing levels of noise, vibrations, and g-forces. Five pilots flew 

a total of 87 test sorties with the Voice Command System, with a general success rate approaching 90 percent. A prototype helmet-mounted 

sight was also evaluated. On July 30, 1983, the AFTI/F-16 aircraft was 

flown back to the General Dynamics facility at Fort Worth, TX, for modification for Phase II. During the Phase I test effort, 118 flight-test sorties were flown, totaling about 177 flight hours. In addition to evaluating 

the DFCS, the potential operational utility of task-tailored flight modes 

(that included decoupling of aircraft attitude and flight path) was also 

assessed. During these unconventional maneuvers, the AFTI/F-16 demonstrated that it could alter its nose position without changing flight path 

and change its flight path without changing aircraft attitude. The aircraft also performed coordinated horizontal turns without banking or 

sideslip.83 NASA test pilot Bill Dana recounted: “In Phase I we evaluated 

non-classic flight control modes. By deflecting the elevators and flaps in 

various relationships, it was possible to translate the aircraft vertically 

without changing pitch attitude or to pitch-point the airplane without 

changing your altitude. You could also translate laterally without using 

bank and yaw-point without translating the aircraft, by using rudder and 

canard inputs programmed together in the flight control computer.”84

Phase II Testing

From mid-1983 through mid-1984, components for the Automated 

Maneuvering Attack System and related avionics systems were installed 

into the AFTI/F-16 at GD in Fort Worth in preparation for the Phase II 

effort. Precision electrical-optical tracking pods were installed in the wing 

root area on both sides of the aircraft. First flight of the AFTI/F-16 in the 

AMAS configuration was on July 31, 1984, with Phase II flight-testing at 

Edwards beginning shortly after the aircraft returned to Dryden on August 

6, 1984. Beginning in September 1984 and continuing through April 1987, 

improved sensors, integrated fire and flight control, and enhancements 

83. Ibid.

84. Gary Creech, “AFTI/F-16 Retires After 22 Years,” The Dryden Express, Dryden Flight Research 

Center, vol. 43, issue 2, Feb. 23, 2001.
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in pilot-vehicle interface were evaluated. During Phase II testing, the 

system demonstrated automatic gun tracking of airborne targets and 

accurate delivery of unguided bombs during 5-g curvilinear toss bomb 

maneuvers from altitudes as low as 200 feet. An all-attitude automatic 

ground collision avoidance capability was demonstrated,85 as was the 

Voice Command System (for interfacing with the avionics system), a 

helmet-mounted sight (used for high off bore sight target cueing), and 

a digital terrain system with color moving map.86 The sortie rate during Phase II was very high. From the start of the AMAS tests in August 

1984 to the completion of Phase II in early 1987, 226 flights were accomplished, with 160 sorties being flown during 1986. To manage this high 

sortie rate, the ground maintenance crews worked a two-shift operation.

Follow-On AFTI/F-16 Testing

Following Phase II in 1987, the forward fuselage-mounted ventral fins 

were removed and the AFTI/F-16 was flown in support of other test efforts 

and new aircraft programs, such as evaluating strike technologies proposed for use in the next generation ground attack aircraft, which eventually evolved into the Joint Strike Fighter (JSF) program.

Strike Technology Testbed

In the summer of 1991, a flight-test effort oriented to close air support 

and battlefield air interdiction began. The focus was to demonstrate technologies to locate and destroy ground targets day or night, good weather 

or bad, while maneuvering at low altitudes. The AFTI/F-16 was modified with two forward-looking infrared sensors mounted in turrets on 

the upper fuselage ahead of the canopy. The pilot was equipped with a 

helmet-mounted sight that was integrated with the infrared sensors. As 

he moved his head, they followed his line of sight and transmitted their 

images to eyepieces mounted in his helmet. The nose-mounted canards 

used in earlier AFTI/F-16 testing were removed. Testing emphasized giving pilots the capability to fly their aircraft and attack targets in darkness 

or bad weather. To assist in this task, a digital terrain map was stored 

85. James Blaylock, Donald Swihart, and William Urshel, “Integration of Advanced Safety Enhancements for F-16 Terrain Following,” AIAA-1987-2906.

86. Charles A. Baird and Franklin B. Snyder, with introduction by Lt. Mark Bierele, AFTI/F-16 

Program Office, “Terrain-Aided Altitude Computations on the AFTI/F-16,” Harris Corporation, 

Melbourne, FL, Aug. 1990.
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in the aircraft computer. Advanced terrain following was also evaluated. This used the AFTI/F-16’s radar to scan terrain ahead of the aircraft and automatically fly over or around obstacles. The pilot could 

select minimum altitudes for his mission. The system would automatically calculate that the aircraft was about to descend below this altitude and initiate a 5 g pullup maneuver. The advanced terrain following 

system was connected to the Automated Maneuvering Attack System, 

enabling the pilot to delivery weapons from altitudes as low as 500 

feet in a 5 g turn. An automatic Pilot Activated Recovery System was 

integrated with the flight control system. If the pilot became disoriented at night or in bad weather, he could activate a switch on his side 

controller. This caused the flight control computer to automatically 

recover the aircraft putting it into a wings-level climb. Many of these 

technologies have subsequently transitioned into upgrades to existing 

fighter/attack aircraft.87

The final incarnation of this unique aircraft would be as the 

AFTI/F-16 power-by-wire flight technology demonstrator.

Power-By-Wire Testbed

During 1997, NASA Dryden had evaluated a single electrohydrostatic 

actuator installation on the NASA F-18 Systems Research Aircraft (SRA), 

with the primary goal being the flight demonstration of power-by-wire 

technology on a single primary flight control surface. The electrohydrostatic actuator, provided by the Air Force, replaced the F-18’s standard 

left aileron actuator and was evaluated throughout the aircraft’s flight 

envelope out to speeds of Mach 1.6. Numerous mission profiles were 

accomplished that included a full series of aerobatic maneuvers. The 

electrohydrostatic actuator accumulated 23.5 hours of flight time on 

the F-18 SRA between January and July 1997. It performed as well as 

the standard F-18 actuator and was shown to have more load capability than required by the aileron actuator specification for the aircraft.88

At about the same time, a Joint Strike Fighter/Integrated Subsystems 

Technology program had been formed to reduce the risk of selected 

87. Finley Barfield, Duke Browning, and Judith Probert, “All Terrain Ground Collision Avoidance 

and Maneuver Terrain Following for Automated Low Level Night Attack,” IEEE AES Systems Magazine, Mar. 1993.

88. Robert Navarro, “Performance of an Electro-Hydrostatic Actuator on the F-18 Systems Research 

Aircraft,” NASA TM-97-206224, NASA Dryden Flight Research Center (1997).
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technology candidates, in particular the power-by-wire approach that 

was intended to replace cumbersome hydraulic actuation systems with 

all-electrical systems for flight surface actuation. A key to this effort was 

the AFTI F-16, which was modified to replace all of the standard hydraulic actuators on the primary flight control surfaces with electrohydrostatic actuators (EHAs) to operate the flaperons, horizontal tails, and 

rudder. Each electrohydrostatic actuator uses an internal electric motor 

to drive an integral hydraulic pump, thus it relies on local hydraulics 

for force transmission (similar to the approach used with the Powered 

Flight Control Units on the Vickers VC10 aircraft discussed earlier).89

In a conventional F-16, the digital fly-by-wire flight control system 

sends out electrical command signals to each of the flight control actuators. These electrical signals drive the control valves (located with the 

actuators) that schedule the fluid from the high-pressure hydraulic pump 

to position the flight control surfaces. Dual engine-driven 3,000 pounds 

per square inch (psi) hydraulic systems power each primary control surface actuator to drive the control surfaces to the desired position. The 

standard F-16 hydraulic actuators operate continuously at 3,000 psi, 

and power is dumped into the actuators, whether it is needed or not.90

In straight and level flight (where most aircraft operate most of their 

time, including even high-performance fighters), the actual electrical 

power requirement of the actuation system is low (only about 500 watts 

per actuator), and excess energy is dissipated as heat and is transferred 

into the fuel system.91

With the electrohydrostatic power design tested in the AFTI/F-16, 

the standard fly-by-wire flight control system was relatively unchanged. 

However, the existing F-16 hydraulic power system was removed and 

replaced by a new power-by-wire system, consisting of an engine-driven 

Hamilton Sundstrand dual 270-volt direct current (DC) electrical power 

generation system (to provide redundancy) and Parker Aerospace electrohydrostatic actuators on the flaperons, rudder, and horizontal stabilizer. The new electrical system powers five dual power electronics 

units, one for each flight control surface actuator. Each power electronics unit regulates the DC electrical power that drives dual motor/pumps 

that are self-contained in each electrohydrostatic actuator. The dual 

89. James W. Ramsey, “Power-by-Wire,” Avionics Magazine, May 1, 2001.

90. Droste, et al., “The General Dynamics Case Study on the F-16 Fly-By-Wire Flight Control System.”

91. Ramsey, “Power-by-Wire.”
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motor/pumps convert electrical power into hydraulic power, allowing 

the piston on the actuators to move the control surfaces. The electrohydrostatic actuators operate at pressures ranging from 300 to 3,000 psi, 

providing power only on demand and generating much less heat. An 

electrical distribution and electrical actuation system simplifies secondary power and thermal management systems, because the need to provide secondary and emergency backup sources of hydraulic power for 

the flight control surfaces is eliminated. The electrohydrostatic system 

also provides more thermal margin, which can be applied to cooling 

other high-demand systems (such as avionics and electronic warfare), 

or, alternatively, the thermal management system weight and volume 

can be reduced making new aircraft designs smaller, lighter, and more 

affordable. Highly integrated electrical subsystems, including power-bywire, reportedly could reduce takeoff weight by 6 percent, vulnerable 

area by 15 percent, procurement cost by 5 percent, and total life-cycle 

cost by 2 to 3 percent, compared with current fighters based on Air Force 

and industry studies. The power-by-wire approach is now being used 

in the Lockheed Martin F-35 Lightning II, with the company estimating a reduction in aircraft weight of as much as 700 pounds because 

of weight reductions in the hydraulic system, the secondary power system, and the thermal management system, made possible because the 

electrical power-by-wire system produces less heat than the traditional 

hydraulic system that it replaces.92

The modified power-by-wire AFTI/F-16 was the first piloted aircraft 

of any type to fly with a totally electric control surface actuation system 

with no hydraulic or mechanical backup flight control capability of any 

kind. It was designed to have the same flight control system responses 

as an unmodified F-16. After the first power-by-wire AFTI/F-16 flight 

on October 24, 2000, at Fort Worth, Lockheed Martin test pilot Steve 

Barter stated aircraft handling qualities with the power-by-wire modifications were indistinguishable from that of the unmodified AFTI/F-16. 

The aircraft was subsequently flown about 10 times, with flight control 

effectiveness of the power-by-wire system demonstrated during supersonic flight. Test pilots executed various flying quality maneuvers, including high-g turns, control pulses (in pitch, roll, and yaw), doublet inputs, 

and sideslips. The tests also included simulated low-altitude attack 

92. Ibid.
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missions and an evaluation of the electrostatic actuator and generator 

subsystems and their thermal behavior under mission loads.93

NASA Dryden hosted the AFTI/F-16 program for 16 years, from 1982 

to 1998. During that time, personnel from Dryden composed 50 percent 

of the AFTI joint test team. Dryden pilots who flew the AFTI/F-16 included 

Bill Dana, Dana Purifoy, Jim Smolka, Rogers Smith, and Steve Ishmael. 

Dryden responsibilities, in addition to its host role, included flight safety, 

operations, and maintenance. Mark Skoog, who served as the USAF 

AFTI/F-16 project manager for many years and later became a NASA 

test pilot, commented: “AFTI had the highest F-16 sortie success rate on 

base, due to Dryden maintenance personnel having tremendous expertise in tailoring their operations to the uniqueness of the vehicle. That 

includes all the other F-16s based at Edwards during those years, none of 

which were nearly as heavily modified as the AFTI.”94 A good summary of 

the AFTI/F-16’s accomplishments was provided by NASA test pilot Dana 

Purifoy: “Flying AFTI was a tremendous opportunity. The aircraft pineered 

many important technologies including glass cockpit human factors, 

automated ground collision avoidance, integrated night vision capability 

and on-board data link operations. All of these technologies are currently being implemented to improve the next generation of both civil and 

military aircraft.”95 The AFTI F-16’s last flight at Dryden was on November 

4, 1997. Over a period of 15 years, it made over 750 flights and was flown 

by 23 pilots from the U.S. Air Force, NASA, the U.S. Marine Corps, and 

the Swedish Air Force. The AFTI F-16 then served as an Air Force technology testbed. Experience and lessons learned were used to help develop the 

production DFBW flight control system used in the F-16. The F-16, the 

F-22, and the F-35, in particular, directly benefited from AFTI/F-16 research 

and technology maturation efforts. After 22 years as a research aircraft 

for NASA and the Air Force, the AFTI F-16 was flown to Wright-Patterson 

AFB, OH, on January 9, 2001, for display at the Air Force Museum.96

Technology Transfer and Lessons Learned

Flight-test results from the AFTI/F-16 program were exceptionally 

well-documented by NASA and widely published in technical papers, 

93. Ramsey, “Power-by-Wire.”

94. Creech, “AFTI/F-16 Retires After 22 Years.”

95. Ibid.

96. “AFTI F-16 Fact Sheet,” National Museum of the U.S. Air Force, 2001.
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memorandums, and presentations.97 These provide invaluable insights 

into the problems, issues, and achievements encountered in this relatively early attempt to integrate an asynchronous digital flight control 

system into a high-performance military jet fighter. As the definition 

implies, in an asynchronous flight control system design, the redundant channels run autonomously. Each computer samples sensors and 

evaluates flight control laws independently. Each separately sends command signals to an averaging or selection device that is used to drive the 

flight control actuators. In this DFCS implementation, the unsynchronized individual computers can sample the sensors at slightly different 

times. Thus, they can obtain readings that may differ quite appreciably 

from one another, especially if the aircraft is maneuvering aggressively. 

Flight control law gains can further amplify these input differences, 

causing even larger differences between the results that are submitted 

to the output selection algorithm.98

During ground qualification of the AFTI/F-16, it was found that 

these differences sometimes resulted in a channel being declared failed 

when no real failure had occurred.99 An even more serious shortcoming of asynchronous flight control systems can occur when the control 

laws contain decision points. Sensor noise and sampling variations may 

cause independent channels within the DFCS to take different paths 

at the decision points and to produce widely divergent outputs.100 This 

occurred on AFTI/F-16 flight No. 44. Two channels in the DFCS declared 

each other failed; the analog backup was not selected because simultaneous failure of two DFCS channels had not been anticipated. The 

pilot could not reset the system, and the aircraft was flown home on 

the single remaining DFCS channel. In this case, all protective redun97. Stephen D. Ishmael, Dale A. Mackall, and Victoria A. Regenie, “Design Implications From 

AFTI/F16 Flight Test,” NASA TM-86026, NASA Ames Research Center, Dryden Flight Research 

Facility, Edwards, CA (1984); Ishmael, et al., “AFTI/F-16 Flight Test Results and Lessons”; Mackall, 

“Development and Flight Test Experiences with a Flight-Crucial Digital Control System,” NASA TP2857, NASA Ames Research Center, Dryden Flight Research Facility, Edwards, CA (Nov. 1988); 

Mackall, “AFTI/F-16 Digital Flight Control System Experience,” in Gary P. Beasley, ed., NASA Aircraft Controls Research 1983, pp. 469–487, NASA CP-2296 (1984), Proceedings of Workshop 

Held at NASA Langley Research Center, Oct. 25–27, 1983.

98. John Rushby, “Formal Methods and the Digital Systems Validation for Airborne Systems,” NASA 

CR-4551, NASA Langley Research Center (Dec. 1, 1993).

99. Mackall, “Development and Flight Test Experiences with a Flight-Crucial Digital Control System.”

100. Rushby, “Formal Methods and the Digital Systems Validation for Airborne Systems.”
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dancy had been lost, yet an actual hardware failure had not occurred. 

Several other difficulties and failure indications were observed during the flight-test program that were traced to asynchronous operation, 

allowing different channels to take different paths at certain selection 

points. The software was subsequently modified to introduce voting at 

some of these software selection points.101

NASA Observations

NASA observations on some of the more serious issues encountered in 

early testing of the AFTI/F-16 asynchronous digital flight control system are worthy of note. For example, an unknown failure in the Stores 

Management System on flight No. 15 caused it to request DFCS mode 

changes at a rate of 50 times per second. The DFCS could not keep 

up and responded at a rate of 5 mode changes per second. The pilot 

reported that the aircraft felt like it was in severe turbulence. The flight 

was aborted, and the aircraft landed safely. Subsequent analysis showed 

that if the aircraft had been maneuvering at the time, the DFCS would 

have failed. A subsequent software modification improved the DFCS’s 

immunity to this failure mode.102

A highly significant flight control law anomaly was encountered on 

AFTI/F-16 flight No. 36. Following a planned maximum rudder “step 

and hold” input by the pilot, a 3-second departure from controlled flight 

occurred. Sideslip angle exceeded 20 degrees, normal acceleration fluctuated from –4 g to +7 g, angle of attack varied between –10 and +20 

degrees, and the aircraft rolled 360 degrees. Severe structural loads were 

encountered with the vertical tailfin exceeding its design load. During 

the out-of-control situation, all control surfaces were operating at rate 

limits, and failure indications were received from the hydraulics and 

canard actuators. The failures were transient and reset after the pilot 

regained control. The problem was traced to a fault in the programmed 

flight control laws. It was determined that the aerodynamic model used 

to develop the control laws did not accurately model the nonlinear nature 

of yaw stability variations as a function of higher sideslip angles. The 

same inaccurate control laws were also used in the real-time AFTI/F-16 

ground flight simulator. An additional complication was caused when 

the side fuselage-mounted air-data probes were blanked by the canard 

101. Mackall, “Development and Flight Test Experiences with a Flight-Crucial Digital Control System.”

102. Ibid.
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at the high angles of attack and sideslip encountered. This resulted in 

incorrect air data values being passed to the DFCS. Operating asynchronously, the different flight control system channels took different paths 

through the flight control laws. Analysis showed these faults could have 

caused complete failure of the DFCS and reversion to analog backup.103

Subsequently, the canards were removed from the command path to 

prevent the AFTI/F-16 from obtaining higher yaw angles.

AFTI/F-16 flight-testing revealed numerous other flight control problems of a similar nature. These prompted NASA engineer Dale Mackall 

to report: “The asynchronous design of the [AFTI/F-16] DFCS introduced 

a random, unpredictable characteristic into the system. The system 

became untestable in that testing for each of the possible time relationships between the computers was impossible. This random time relationship was a major contributor to the flight test anomalies. Adversely 

affecting testability and having only postulated benefits, asynchronous 

operation of the DFCS demonstrated the need to avoid random, unpredictable, and uncompensated design characteristics.” Mackall also provided additional observations that would prove to be highly valuable in 

developing, validating, and certifying future software-intensive digital 

fly-by-wire flight control system designs. Urging more formal approaches 

and rigorous control over the flight control system software design and 

development process, Mackall reported:

The criticality and number of anomalies discovered in flight and 

ground tests owing to design oversights are more significant than 

those anomalies caused by actual hardware failures or software 

errors. . . . As the operational requirements of avionics systems 

increase, complexity increases. . . . If the complexity is required, 

a method to make system designs more understandable, more 

visible, is needed . . . qualification of such a complex system as 

this, to some given level of reliability, is difficult . . . the number 

of test conditions becomes so large that conventional testing 

methods would require a decade for completion. The faulttolerant design can also affect overall system reliability by being 

made too complex and by adding characteristics which are random in nature, creating an untestable design.104

103. Ibid.

104. Ibid.
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Aircraft Certification Contributions

Certification of new aircraft with digital fly-by-wire flight control systems, 

especially for civilian airline service, requires software designs that provide highly reliable, predictable, and repeatable performance. For this 

reason, NASA experts concluded that a comprehensive understanding of 

all possible software system behaviors is essential, especially in the case 

of highly complex systems. This knowledge base must be formally documented and accurately communicated for both design and system certification purposes. This was highlighted in a 1993 research paper sponsored 

by NASA and the Federal Aviation Administration (FAA) that noted:

This formal documentation process would prove to be a tremendously difficult and challenging task. It was only feasible 

if the underlying software was rationally designed using principles of abstraction, layering, information-hiding, and any 

other technique that can advance the intellectual manageability of the task. This calls strongly for an architecture that 

promotes separation of concerns (whose lack seems to be the 

main weakness of asynchronous designs), and for a method 

of description that exposes the rationale for design decisions 

and that allows, in principle, the behavior of the system to be 

calculated (i.e., predicted or, in the limit, proved) . . . formal 

methods can make their strongest contribution to quality assurance for ultra-dependable systems: they address (as nothing 

else does) [NASA engineer Dale] Mackall’s plea for ‘a method 

to make system designs more understandable, more visible.’105

Formal software development methodologies for critical aeronautical 

and space systems developments have been implemented within NASA 

and are contained in certification guidebooks and other documents for 

use by those involved in mission critical computer and software systems.106 

Designed to help transition Formal Methods from experimental use into 

105. Rushby, “Formal Methods and the Certification of Critical Systems,” SRI CSL-93-07, SRI 

International (Nov. 1993).

106. “Formal Methods Specification and Verification Guidebook for Software and Computer 

Systems, Volume I: Planning and Technology Insertion,” NASA TP-98-208193 (1998); “Formal 

Methods Specification and Analysis Guidebook for the Verification of Software and Computer 

Systems, Volume II: A Practitioner’s Companion,” NASA GB-001-97 (1997).
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practical application for critical software requirements and systems 

design within NASA, they discuss technical issues involved in applying 

Formal Methods techniques to aerospace and avionics software systems. 

Dryden’s flight-test experience and the observations obtained from flighttesting of such systems were exceptionally well-documented and would 

prove to be highly relevant to NASA, the FAA, and military service programs oriented to developing Formal Methods and structured approaches 

in the design, development, verification, validation, testing, and certification of aircraft with advanced digital flight control systems.107 The NASA 

DFBW F-8 and AFTI/F-16 experiences (among many others) were also 

used as background by Government and industry experts tasked with 

preparing the FAA Digital Systems Validation Handbook. Today, the FAA 

uses Formal Methods in the specification and verification of software 

and hardware requirements, designs, and implementations; in the 

identification of the benefits, weaknesses, and difficulties in applying these 

Formal Methods to digital systems used in critical applications; and in 

support of aircraft software systems certification.

YA-7D DIGITAC

Digital Flight Control for Tactical Aircraft (DIGITAC) was a joint program 

between the Air Force Flight Dynamics Laboratory (AFFDL) at WrightPatterson AFB, OH, and the USAF Test Pilot School (TPS) at Edwards 

AFB. Its purpose was to develop and demonstrate digital flight control 

technology for potential use in future tactical fighter and attack aircraft, 

including the feasibility of using digital flight control computer technology 

to optimize an airplane’s tracking and handling qualities for a full range 

of weapons delivery tasks. The second prototype LTV YA-7D (USAF serial 

No. 67-14583) was selected for modification as the DIGITAC testbed by 

replacing the analog computer of the YA-7D Automated Flight Control 

System (AFCS) with the DIGITAC digital multimode flight control system 

that was developed by the AFFDL. The mechanical flight control system 

in the YA-7D was unchanged and was retained as a backup capability.

The YA-7D’s flight control system was eventually upgraded to DIGITAC 

II configuration. DIGITAC II used military standard data buses and 

transferred critical flight control data between individual computers 

and between computers and remote terminals. The data buses used 

107. Szalai, et al., “Digital Fly-By-Wire Flight Control Validation Experience.”
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were dual channel wire and dual channel fiber optic and were selectable 

in the cockpit by the pilot to allow him to either fly-by-wire or fly-bylight. Alternately, for flight-test purposes, the pilot was able to implement 

one wire channel and one fiber optic channel. During early testing, the 

channel with the multifiber cables (consisting of 210 individual fibers) 

encountered numerous fiber breakage problems during normal ground 

maintenance. The multifiber cable design was replaced by single-fiber 

cables with tough protective shields, a move that improved data transmission qualities and nearly eliminated breakage issues. The DIGITAC 

fly-by-light system flew 290 flights during a 3-year period, performing 

flawlessly with virtually no maintenance. It was so reliable that it was used 

to fly the aircraft on all routine test missions. The system performance 

and reliability was considered outstanding, with the technical approach 

assessed as ready for consideration for use in production aircraft.108

The DIGITAC YA-7D provided the TPS with a variable stability 

testbed aircraft for use in projects involving assessments of advanced 

aircraft flying qualities. Results obtained from these projects contributed to the flying qualities database in many areas, including degradedmode flight control cross-coupling, control law design, pro versus 

adverse yaw studies, and roll-subsistence versus roll-time-delay studies. Under a TPS project known as Have Coupling, the YA-7D DIGITAC 

aircraft was used to investigate degradation to aircraft handling qualities that would occur in flight when a single pitch control surface (such 

as one side of the horizontal stabilizer) was damaged or impaired. An 

asymmetric flight control situation would result when a pure pitch 

motion was commanded by the pilot, with roll and yaw cross-coupling 

motions being produced. For the Have Coupling tests, various levels of 

cross-coupling were programmed into the DIGITAC aircraft. The resulting data provided a valuable contribution to the degraded flight control 

mode handling qualities body of knowledge. This included the interesting finding that with exactly the same amounts of cross-coupling 

present, pilot ratings of aircraft handling qualities in flight-testing were 

significantly different compared with those rating obtained in the 

ground-based simulator.109

108. Rediess, et al., “Technology Review of Flight Crucial Flight Control Systems (Application of 

Optical Technology).”

109. E. Robert Lemble, “DIGITAC—A Unique Digital Flight Control Testbed Aircraft,” AIAA-1990-

1288, SFTE, DGLR, and SETP Biannual Flight Test Conference, 5th, Ontario, CA, May 22–24, 1990.
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The TPS operated the YA-7D DIGITAC aircraft for over 15 years, beginning in 1976. It made significant contributions to advances in flight control 

technology during investigations involving improved directional control, 

the effect of depressed roll axis on air-to-air tracking, and airborne verification of computer-simulated flying qualities. The DIGITAC aircraft was 

used to conduct the first Air Force flight tests of a digital fight control system, and it was also used to flight-test the first fiber-optical fly-by-light 

DFCS. Other flight-test firsts included the integration of a dynamic gun 

sight and the flight control system and demonstrations of task-tailored 

multimode flight control laws.110 The DIGITAC YA-7D is now on display 

at the Air Force Flight Test Center Museum at Edwards AFB.

Highly Maneuverable Aircraft Technology

The Highly Maneuverable Aircraft Technology (HiMAT) program provides an interesting perspective on the use of unmanned research aircraft equipped with digital fly-by-wire flight control systems, one that 

is perhaps most relevant to today’s rapidly expanding fleet of unpiloted 

aircraft whose use has proliferated throughout the military services 

over the past decade. HiMAT research at Dryden was conducted jointly 

by NASA and the Air Force Flight Dynamics Laboratory at NASA Dryden 

between 1979 and 1983. The project began in 1973, and, in August 

1975, Rockwell International was awarded a contract to construct two 

HiMAT vehicles based on the use of advanced technologies applicable 

to future highly maneuverable fighter aircraft. Designed to provide a 

level of maneuverability that would enable a sustained 8 g turn at 0.9 

Mach at an altitude of 25,000 feet, the HiMAT vehicles were approximately half the size of an F-16. Wingspan was about 16 feet, and length 

was 23.5 feet. A GE J85 turbojet that produced 5,000 pounds of static 

thrust at sea level powered the vehicle that could attain about Mach 

1.4. Launched from the NASA B-52 carrier aircraft, the HiMAT weighed 

about 4,000 pounds, including 660 pounds of fuel. About 30 percent of 

the airframe consisted of experimental composite materials, mainly 

fiberglass and graphite epoxy. Rear-mounted swept wings, a digital flight 

control system, and controllable forward canards enabled exceptional 

maneuverability with a turn radius about half of a conventional piloted 

fighter. For example, at Mach 0.9 at 25,000 feet, the HiMAT could 

110. Dennis R. Furman, “USAF Test Pilot School Use of DIGITAC in Systems Testing,” Society of 

Automotive Engineers Document No. 851827, Oct. 1985.
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Research on the HiMAT remotely piloted test vehicle was conducted by NASA and the Air Force 

Flight Dynamics Laboratory between 1979 and 1983. NASA.

sustain an 8-g turn, while F-16 capability under the same conditions is 

about 4.5 g.111 

Ground-based, digital fly-by-wire control systems, developed at 

Dryden on programs such as the DFBW F-8, were vital to success of 

the HiMAT remotely piloted research vehicle approach. NASA Ames 

Research Center and Dryden worked closely with Rockwell International 

in design and development of the two HiMAT vehicles and their ground 

control system, rapidly bringing the test vehicles to flight status. Many 

tests that would have been required for a more conventional piloted 

research aircraft were eliminated, an approach largely made possible 

by extensive use of computational aerodynamic design tools developed 

at Ames. This resulted in drastic reductions in wind tunnel testing but 

caused the need to devote several HiMAT flights to obtain stability and 

control data needed for refinements to the digital flight control system.112

The HiMAT flight-test maneuver autopilot was based on a design 

developed by Teledyne Ryan Aeronautical, then a well-known man111. Robert W. Kempel and Michael R. Earls, “Flight Control Systems Development and Flight Test 

Experience with the HiMAT Research Vehicles,” NASA TP-2822 (June 1988).

112. “HiMAT: Highly Maneuverable Aircraft Technology,” NASA Dryden FS-2002-06-025-DFRC 

(2002).
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ufacturer of target drones and remotely piloted aircraft. Teledyne 

also developed the backup flight control system.113 Refining the vehicle 

control laws was an extremely challenging task. Dryden engineers and test 

pilots evaluated the contractor-developed flight control laws in a 

ground simulation facility and then tested them in flight, making adjustments until the flight control system performed properly. The HiMAT 

flight-test maneuver autopilot provided precise, repeatable control, 

enabling large quantities of reliable test data to be quickly gathered. 

It proved to be a broadly applicable technique for use in future flight 

research programs.114

Launched from the NASA B-52 at 45,000 feet at Mach 0.68, the 

HiMAT vehicle was remotely controlled by a NASA research pilot in a 

ground station at Dryden, using control techniques similar to those in 

conventional aircraft. The flight control system used a ground-based 

computer interlinked with the HiMAT vehicle through an uplink and 

downlink telemetry system. The pilot used proportional stick and rudder inputs to command the computer in the primary flight control system. A television camera mounted in the cockpit provided visual cues 

to the pilot. A two-seat Lockheed TF-104G aircraft was used to chase 

each HiMAT mission. The F-104G was equipped with remote control 

capability, and it could take control of the HiMAT vehicle if problems 

developed at the ground control site. A set of retractable skids was 

deployed for landing, which was accomplished on the dry lakebed adjacent to Dryden. Stopping distance was about 4,500 feet. During one of 

the HiMAT flight tests, a problem was encountered that resulted in a 

landing with the skids retracted. A timing change had been made in 

the ground-based HiMAT control system and in the onboard software 

that used the uplinked landing gear deployment command to extend 

the skids. Additionally, an onboard failure of one uplink receiver contributed to cause the anomaly. The timing change had been thoroughly 

tested with the onboard flight software. However, subsequent testing 

determined that the flight software operated differently when an uplink 

failure was present.115

113. Kempel, “Flight Experience with a Backup Flight Control System for the HiMAT Research 

Vehicle,” AIAA Paper 82-1541 (Aug. 1982).

114. E.L. Duke, F.P. Jones, and R.B Roncoli, “Development of a Flight Test Maneuver Autopilot for a 

Highly Maneuverable Aircraft,” NASA TP-2218 (1986).

115. Mackall, “Development and Flight Test Experiences with a Flight-Crucial Digital Control System,” p. 112.
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HiMAT research also brought about advances in digital flight control systems used to monitor and automatically reconfigure aircraft 

flight control surfaces to compensate for in-flight failures. HiMAT provided valuable information on a number of other advanced design features. These included integrated computerized flight control systems, 

aeroelastic tailoring, close-coupled canards and winglets, new composite airframe materials, and a digital integrated propulsion control system. Most importantly, the complex interactions of this set of then-new 

technologies to enhance overall vehicle performance were closely evaluated. The first HiMAT flight occurred July 27, 1979. The research program ended in January 1983, with the two vehicles completing a total 

of 26 flights, during which 11 hours of flying time were recorded.116 

The two HiMAT research vehicles are today on exhibit at the NASA 

Ames Research Center and the Smithsonian Institution National Air 

and Space Museum.

International CCV Flight Research Efforts

As we have seen earlier, as far back as the Second World War and 

continuing through the 1950s and 1960s, the Europeans in particular were very active in exploiting the benefits to be gained from the 

use of fly-by-wire flight control systems in aircraft and missile systems. 

Experimental fly-by-wire research aircraft programs in Europe and 

Japan rapidly followed, sometimes nearly paralleled, and even occasionally led NASA and Air Force fly-by-wire research programs, often 

with the assistance of U.S. flight control system companies. As with U.S. 

programs, foreign efforts focused on the application of digital fly-bywire flight control systems in conjunction with modifications to existing 

service aircraft to create unstable CCV testbeds. Foreign CCV research 

efforts conclusively validated the benefits attainable from integration of 

digital computers into fly-by-wire flight control systems and provided 

experience and confidence in their use in new aircraft designs that have 

increasingly become multinational.

German CCV F-104G

Capitalizing on their earlier experience with analog fly-by-wire flight 

control research, by early 1975 the Germans had begun a flight research 

116. Dwain A. Deets, V. Michael DeAngelis, and David P. Lux, “HiMAT Flight Program: Test Results 

and Program Assessment Overview,” NASA TM-86725 (June 1986).
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program to investigate the flying qualities of a highly unstable highperformance aircraft equipped with digital flight controls. For this 

purpose, they modified a Luftwaffe Lockheed F-104G to incorporate a 

quadruplex digital flight control system. Known as the CCV F-104G, it 

featured a canard (consisting of another F-104G horizontal tail) mounted 

at a fixed negative incidence angle of 4 degrees, on the upper fuselage 

behind the cockpit and a large jettisonable weight carried under the aft 

fuselage. These features, in conjunction with internal fuel transfer, were 

capable of moving the aircraft’s center of gravity rearward to create a negative stability margin of up to 20 percent. The CCV F-104G flew for the 

first time in 1977 from the German flight research center at Manching, 

with flight-testing of the aircraft in the canard configuration beginning 

in 1980. The CCV F-104G test program ended in 1984 after 176 flights.117

U.K. Jaguar ACT

In the U.K., the Royal Aircraft Establishment began an effort oriented to producing a CCV testbed in 1977. For this purpose, an AngloFrench Jaguar strike fighter was modified by British Aerospace (BAe) 

to prove the feasibility of active control technology. Known as the Jaguar 

Active Control Technology (ACT), the aircraft’s mechanical flight control 

system was entirely removed and replaced with a quad-redundant 

digital fly-by-wire control system that used electrical channels to relay 

instructions to the flight control surfaces. The initial flight of the Jaguar 

ACT with the digital FBW system was in October 1981. As with the 

CCV F-104G, ballast was added to the aft fuselage to move the 

center of gravity aft and destabilize the aircraft. In 1984, the Jaguar ACT 

was fitted with rounded oversized leading-edge strakes to 

move the center of lift of the aircraft forward, further contributing to 

pitch instability. It first flew in this configuration in March 

1984. Marconi developed the Jaguar ACT flight control system. It 

included an optically coupled data transmission link that was 

essentially similar to the one that they had developed for the U.S. 

Air Force YC-14 program (an interesting example of the rapid 

proliferation of advanced aerspace technology between nations).118 

117. Butter, “Control, Navigation, Avionics, Cockpit,” Aeronautical Research in Germany (From 

Lilienthal to Today), Springer-Verlag, Berlin Heidelberg, 2004.

118. Rediess, et al., “Technology Review of Flight Crucial Flight Control Systems (Application of 

Optical Technology).”
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Flight-testing began in 1981, with the test program ending in 1984 

after 96 flights.119

French Mirage NG

Although not intended purely as a research aircraft, the French Dassault 

Mirage 3NG (Nouvelle Generation) was a greatly modified Mirage 

IIIE single-engine jet fighter that was used to demonstrate the improved 

air combat performance advantages made possible using relaxed 

static stability and fly-by-wire. One prototype was built by Dassault; 

modifications included destabilizing canards, extended wing root leading edges, an analog-computer-controlled fly-by-wire flight control system (based on that used in the production Mirage 2000 fighter), and the 

improved Atar 9K-50 engine. The Mirage 3NG first flew in December 

1982, demonstrating significant performance improvements over the 

standard operational Mirage IIIE. These were claimed to include a 

20–25-percent reduction in takeoff distance, a 40-percent improvement 

in time to reach combat altitude, a nearly 10,000-foot increase in supersonic ceiling, and similarly impressive gains in acceleration, instantaneous turn rate, and combat air patrol time.

Japanese CCV T-2

In Japan, the CCV approach that was taken involved modification of a 

Mitsubishi T-2 jet training aircraft. Horizontal canards were fitted to 

reduce static stability, and an all-movable vertical surface was added 

to the forward fuselage to enable direct side force control investigations. The existing wing-mounted flaps were modified to enable direct 

lift control and maneuver load control studies. A triply redundant digital fly-by-wire flight control system was installed with quadruplex 

pilot force sensors used to sense stick and rudder pedal forces and aircraft motion sensors. Aircraft motion sensors (such as pitch, roll, and 

yaw rate gyros, and vertical and lateral acceleration sensors) were also 

quadruplex. The original mechanical flight control system was retained 

as a backup mode. Three identical digital computers processed 

sensor signals, and the resultant command signals were used to control the horizontal stabilizer, leading and trailing edge flaps, rudder, 

and vertical canard. Electrohydraulic actuators converted electrical 

119. C.J. Yeo, “The Fly-by-Wire Jaguar,” Society of Experimental Test Pilots (SETP) 27th Symposium 

Proceedings, Beverly Hills, CA, Sept. 28–Oct. 1, 1983, pp. 193–214.
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signals into mechanical inputs for the control surface actuators. The 

CCV T-2 first flew in August 1983. After 24 flights by Mitsubishi, the 

aircraft was delivered to the Japanese Technical Research Development 

Institute (TRDI) at Gifu Air Base in March 1984 for government flighttesting, which was completed in March 1986.120

These research programs (along with the Soviet Projekt 100LDU testbed 

discussed earlier) provided invaluable hands-on experience with stateof-the-art flight control technologies. Data from the Jaguar ACT and 

the CCV F-104G supported the Experimental Aircraft Program (EAP) 

and contributed to the technology base for the Anglo-German-ItalianSpanish Eurofighter multirole fighter, now known as the Typhoon. Many 

other advanced aircraft development programs, including the French 

Rafale, the Mitsubishi F-2 fighter, the Russian Su-27 family of fighters 

and attack aircraft, and the entire family of Airbus airliners, were the 

beneficiaries of these research efforts. In addition, the importance of the 

infusion of technology made possible by open dissemination of NASA 

technical publications should not be underestimated.

X-29

The Grumman X-29 research aircraft played a very interesting role in 

the evolution of modern fly-by-wire flight control systems. Exotic in 

appearance, with its forward-swept wings and large movable canard 

control surfaces, the X-29 was highly unstable about the longitudinal 

axis with a static stability margin of –35 percent. This level of instability probably indicates that the X-29 represents the most unstable 

piloted aircraft that has ever been successfully flown. Not only did it 

fly, but it also demonstrated good controllability at very high angles of 

attack. This degree of success was only possible through the use of a very 

advanced fly-by-wire flight control system that employed a combination 

of both digital computers in the primary system and analog computers 

in the backup system. The program began in 1977, with the Defense 

Advanced Research Projects Agency (DARPA) and the Air Force Flight 

Dynamics Laboratory jointly soliciting industry proposals for a research 

aircraft designed to investigate the forward-swept wing concept in a highperformance aircraft application. In December 1981, Grumman Aircraft 

120. Katsuhei Shibata and Hideaki Ohmiya, “The T-2 Control Configured Vehicle Development, Integration and Flight Test,” AIAA-1988-3882, published in Technical Papers, pt. 1 (A89-18051 05-06) 

AIAA/IEEE Digital Avionics Systems Conference, San Jose, CA, Oct. 17–20, 1988, pp. 177–184.
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The X-29 represents the most unstable piloted aircraft that has ever successfully flown. NASA.

was selected to build two aircraft, which were designated X-29. The 

most unique and visually obvious aspect of the design was the forwardswept wings that incorporated a thin supercritical airfoil, but there were 

many other areas of the X-29 design that embodied advanced technology. The aircraft used advanced composite materials and unusual 

construction approaches. Its control system made use of variable camber wing surfaces, aft fuselage-mounted strake flaps, fully movable 

canards mounted on the sides of the engine inlets, and a computerized 

fly-by-wire flight control system to maintain control of the otherwise 

highly unstable aircraft.121

In constructing the X-29s, Grumman used the forward fuselage and 

nose landing gear from two Northrop F-5A fighters. Control surface 

actuators and the main landing gear came from the F-16. The unique 

aspect of the X-29 airframe, its forward-swept wing, was developed by 

Grumman. Because of the major differences between the X-29s and the 

F-5As, the modified aircraft were assigned new USAF serial numbers, 

becoming 82-0003 and 82-0049. The main difference between the two 

121. Terrill W. Putnam, “X-29 Flight-Research Program,” NASA TM-86025, Ames Research CenterDryden Flight Research Facility, Edwards, CA (Jan. 1984).
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X-29s was the emergency spin parachute system mounted at the base of 

the rudder on the second aircraft. The X-29 flight research program was 

conducted by the NASA Dryden Flight Research Center in two phases 

and included participation by the Air Force Flight Test Center and the 

Grumman Corporation.122 The joint NASA–Air Force portion of the X-29 

test program extended from 1984 to 1991; the Air Force conducted a 

follow-on investigation of vortex flow control (VFC) that lasted into 1992.

The X-29’s thin supercritical forward-swept wing presented significant design challenges. The typical stall pattern of an aft-swept wing, 

from wingtip to root, is reversed for a forward-swept wing, which stalls 

from the root to the tip. The aerodynamic lift forces on the outer portions of a forward-swept wing produce a twisting moment that tends to 

force the leading edge further upward. This increases the angle of attack 

at the wingtips, causing even further twisting that, if uncontrolled, can 

lead to structural failure of the wing, a phenomenon known as aeroelastic divergence. To deal with this problem, Grumman made use of 

state-of-the-art composite materials in designing the wing external skin, 

which it laminated in a way that produced an inherent coupling between 

wing bending and torsion loads, a concept known as aeroelastic tailoring. At increasing angles of attack (higher lift), the structural characteristics incorporated into composite laminates were designed to ensure that 

the wing twisted to counter the upward twist produced by aerodynamic 

loads. The key to the design was balancing the aerodynamic aspects of the 

wing’s configuration with the structural characteristics of the composite 

laminate to control potential aeroelastic divergence. The wing substructure and the basic airframe itself were made of aluminum and titanium.123

The X-29 featured an unusual combination of flight control surfaces. These consisted of forward-mounted canards that contributed 

positive lift and provided primary control about the pitch axis. Wing 

flaperons (combination flaps and ailerons) could change the camber 

of the wing and also functioned as ailerons for roll control. The actuators used to control wing camber were mounted externally in streamlined fairings at the trailing edge of the wing because of the thinness 

of the supercritical airfoil. The strake flaps on each side of aft fuselage 

augmented the canards, proving additional pitch control. The control surfaces were electronically linked to a triple-redundant digital fly-by-wire 

122. Ibid.

123. Ibid.
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flight control system (with analog backup) that provided artificial stability necessary for controlling the inherently unstable forward-swept 

wing, close-coupled canard design used on the X-29. Each of the three 

digital flight control computers had an analog backup. If one of the 

digital computers failed, the remaining two took over. If two of the 

digital computers failed, the flight control system switched to the analog mode. If one of the analog computers failed, the two remaining analog computers took over.124

Grumman chief test pilot Charles A. “Chuck” Sewell flew the first 

X-29 at Edwards AFB on December 14, 1984.125 During the Phase I 

research effort, X-29 aircraft No. 1 was used exclusively, flying 242 

times. Its wingtips remained unstalled up to the 21-degree angle of 

attack allowed in Phase I testing. This limitation was due to the fact that 

an anti-spin parachute was not installed on the aircraft. The aeroelastic tailored wing prevented structural divergence of the wing, and the 

digital flight control system functioned safely and reliably. Flight control laws and control surface effectiveness combined to provide good 

pilot handling qualities during maneuvering flight. The aircraft’s 

supercritical airfoil contributed to enhanced cruise and maneuver 

performance in the transonic regime.126 However, overall drag reduction was not as great as had been predicted for the configuration. On 

December 13, 1985, with NASA test pilot Steve Ishmael at the controls, the X-29 became the first aircraft with a forward-swept wing to 

fly beyond the speed of sound, reaching Mach 1.03 in level flight.127 

Other test pilots who flew the X-29 during Phase I of the joint test 

program were NASA test pilot Rogers Smith, Lt. Col. Theodore “Ted” 

Wierzbanowski and Maj. Harry Walker from the Air Force, and Navy 

Cdr. Ray Craig.

124. Robert Clarke, John J. Burken, John T. Bosworth, and Jeffery E. Bauer, “X-29 Flight Control Lessons Learned,” NASA TM-4598 (June 1984).

125. Sewell was later killed in the crash of a World War II-era Grumman TBM Avenger torpedo 

bomber during a takeoff from Danielston, CT, on Aug. 4, 1988. With Grumman since 1969, he 

had flown with the USMC in Korea and Vietnam and had over 10,000 flying hours on 140 different aircraft types.

126. Developed by NASA and originally tested on the NASA F-8 Super Critical Wing (SCW) 

aircraft at Dryden in the 1970s, a supercritical airfoil has a flattened upper surface compared to a 

conventional airfoil shape. This shape delays the onset of shock waves on the upper wing surface 

and reduces their strength, theoretically resulting in a decrease in overall drag.

127. “X-29 Fact Sheet,” NASA Dryden Flight Research Center, Mar. 1, 2008.
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The second X-29 aircraft, modified to incorporate an anti-spin parachute and its deployment mechanism, was used during Phase II testing to investigate the aircraft’s high-angle-of-attack characteristics and 

the potential usefulness of the forward-swept wing and canard configuration on military fighter plane designs. First flown on May 23, 1989, 

it would eventually fly 120 research flights and demonstrate control 

and maneuvering qualities that were better in many cases than the predictions derived from computational methods and simulation models. 

NASA, Air Force, and Grumman project pilots reported that the X-29 

had excellent control response up to angles of attack of 45 degrees, with 

limited controllability still available at up to 67 degrees angle of attack. 

Phase II flight-testing defined an allowable X-29 flight envelope that 

extended to Mach 1.48, an altitude of just over 50,000 feet, an angle of 

attack of up to 50 degrees at 1 g and 35 degrees at airspeeds up to 300 

knots. Much of the X-29’s high-angle-of-attack capability was attributed 

to the quality of the flight control laws that were cooperatively developed by NASA and the Air Force. These had initially been developed 

using results obtained from extensive wind tunnel testing and predictions derived from radio-controlled flight tests of a 22-percent scale drop 

model at NASA Langley Research Center.128 Flight control system engineers at NASA Dryden and the Air Force Flight Test Center at Edwards 

used these as the basis for detailed flight control system design. This 

design used a combination of pitch rate and angle of attack in developing the longitudinal control laws. Selectable gain was included in the 

flight control system design, and this was used by the X-29 test pilots 

during flying qualities assessments and evaluations of the effects of 

control law gain changes at higher angles of attack. Prior to the start 

of flight-testing, wing rock was estimated to restrict the available angle 

of attack to less than about 35 degrees. However in flight-testing, wing 

rock amplitude was found to be less than half of what had been predicted, allowing the roll rate to aileron gain to be lowered to one-fourth 

of the value that had been derived from preflight data using the subscale free flight model. The available flight envelope was extended to 

67 degrees angle of attack at 1 g. Maneuvering flight about all axes was 

cleared up to an angle of attack of 45 degrees in 1 g flight. The reduced 

wing rock that had been observed in flight was apparently due to higher 

128. Chambers, Partners in Freedom.
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roll damping and increased aileron control power for large 

aileron deflections.129

Preflight predictions of the X-29’s pitch capabilities matched flighttest results up through 40 degrees angle of attack. Differences in noseup pitching moment above an angle of attack of 40 degrees were found 

to require more canard deflection than predicted. Large yaw asymmetries led to several instances in which the aircraft tended to stabilize at very 

high nose-up pitch angles during maneuvers at angles of attack above 50 

degrees, with the aircraft at an aft center of gravity. Modifications to provide additional nose-down pitch authority were not possible because of 

physical limits on canard deflection. Maximum pitch rates were limited by 

the high level of static instability inherent in the X-29 design and control 

surface rate limits. New actuators with at least a 50-percent higher actuation rate would have been required to achieve pitch rates comparable to 

those of an operational fighter like the F/A-18. The full wingspan flaperons were found to provide good roll control that was not affected by the 

fact that the X-29 did not use wing leading-edge maneuvering flaps. Pilotselectable variable gain capability was used during examination of airplane stability and maneuverability. Basic fighter maneuvers were flown, 

and roll and yaw gains were increased to improve roll performance. A gain 

that provided maximum rudder authority resulted in the best pilot comments. Roll coordination was better than anticipated, with rudder effectiveness also higher than preflight predictions at angles of attack between 

20 and 40 degrees. Yaw asymmetries developed above 40 degrees angle of 

attack. Diminished aileron and rudder power was not sufficient to overpower these asymmetries. Increasing gain further produced rudder saturation, resulting in uncoordinated turns, a result that was disliked by 

test pilots, even though this actually resulted in better roll performance.130

Flight-test data from Phase II, the high-angle-of-attack and military utility phase of the X-29 program, satisfied the program’s primary 

objective. The technologies demonstrated in the program had potential to improve future fighter aircraft mission performance, and the 

forward-swept wings/movable canard configuration provided excellent control response at up to 45 degrees angle of attack. Very importantly, the X-29A program provided a significant pool of knowledge that 

129. Jeffrey E. Bauer, Robert Clarke, and John J. Burken, “Flight Test of the X-29A at High Angle of 

Attack: Flight Dynamics and Controls,” NASA TP-3537 (Feb. 1995).

130. Ibid.
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was very useful background for fine-tuning ground-based predictive 

techniques for high-angle-of-attack aircraft.131 One significant potential safety issue with the flight control system, the danger of sensor 

selection thresholds being set too wide, was discovered during the X-29 

test program. The flight control system used three sources of air data in 

its computations. These air data sources were the nose probe and two 

probes mounted one on each side of the forward fuselage. The selection 

algorithm in the flight control system used the data from the nose probe 

as the primary source, provided it was within some threshold of the data 

from both side probes. However, the data selection threshold was intentionally large to accommodate known data errors in certain flight modes 

because of the position of the side sensors in the airflow. Long after the 

start of flight-testing, ground simulation revealed that the nose probe 

could, in some circumstances, furnish erroneous information at very low 

flight speeds, causing the X-29 to go out of control. Although this fault 

was successfully identified through ground simulation, 162 flights had 

already been flown before it was detected and corrected.132

In 1992, the Air Force began a follow-on program with X-29 No. 2 

that investigated the use of vortex flow control as a means of providing 

increased aircraft control at very high angles of attack, at which normal 

rudder control is ineffective. Wind tunnel tests had showed that injection of air into the vortexes coming off the nose of the aircraft would 

change the direction of vortex flow. The forces created on the nose of the 

aircraft could be used to control directional (yaw) stability. The second 

X-29 aircraft was modified to incorporate two high-pressure nitrogen 

tanks, related control valves with two small nozzle jets on the forward 

upper portion of the nose. The nozzles injected air into the vortexes, 

which flowed off the nose of the aircraft at high angles of attack. From 

May to August 1992, 60 test flights were flown. Data from these flights 

were used to determine that VFC was more effective than expected in 

generating yaw forces, but it was less successful in providing control 

when sideslip was present, and it did little to decrease roll oscillations.133

The two X-29 aircraft flew a total of 436 flights, 254 by the first, and 182 

by the second. The former is exhibited at the National Museum of the 

131. Ibid.; Chambers, Partners in Freedom.

132. Mackall, “Development and Flight Test Experiences with a Flight-Crucial Digital Control System.”

133. R. Guyton and F. Luria, “Flight Testing of Pneumatic Forebody Vortex Control on the X-29 Technology Demonstrator,” Society of Automotive Engineers (SAE) No. 922008, Oct. 1992.
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The X-31 aircraft, showing thrust vectoring paddles. NASA.

United States Air Force, while the latter remained at Edwards and is on 

exhibit at the Dryden Flight Research Center.134

X-31 Enhanced Fighter Maneuverability Demonstrator

The X-31 was the first international experimental aircraft development 

program in which the U.S. participated. Two X-31 Enhanced Fighter 

Maneuverability (EFM) demonstrator aircraft were designed and constructed by Rockwell International Corporation’s North American 

Aircraft Division and Deutsche Aerospace. Assigned U.S. Navy bureau 

Nos. 164584 and 164585, the aircraft would be used to obtain data that 

could be applied to the design of highly maneuverable next-generation 

fighters. During the conceptual phase of the program, the personnel 

examined the application of EFM technologies and defined the requirements for the demonstrator aircraft. Next, the preliminary design of the 

demonstrator and the manufacturing approach were defined. Technical 

experts from the U.S. Navy, German Federal Ministry of Defense, and 

134. Hallion and Michael H. Gorn, On the Frontier: Experimental Flight at NASA Dryden, (Washington: Smithsonian Books, 2003), Appendix W, X-29 flight test chronology.
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NASA evaluated all aspects of the design. Detail design and fabrication 

followed, with the two aircraft being assembled at the Rockwell 

International (now Boeing) facility at Palmdale, CA. Both aircraft were 

required to fly a limited flight-test program at Rockwell. The first aircraft 

flew its first flight on October 11, 1990, piloted by Rockwell chief test pilot 

Ken Dyson. The second aircraft made its first flight on January 19, 1991, 

with Deutsche Aerospace chief test pilot Dietrich Seeck at the controls.135

The X-31 had a digital fly-by-wire flight control system that included 

four digital flight control computers with no analog or mechanical 

backup. Three synchronous main computers drove the flight control 

surfaces. The fourth computer served as a tiebreaker in case the three 

main computers produced conflicting commands. Three thrust vectoring 

paddles were mounted on the X-31’s aft fuselage adjacent to the engine 

nozzle. Directed by the DFBW flight control system, the paddles were 

moved in and out of the exhaust flow with the resultant thrust vectoring 

augmenting the aerodynamic control surfaces in pitch and yaw control 

to improve maneuverability. Made of an advanced carbon-fiberreinforced composite material, the paddles could sustain temperatures of 

up to 1,500 degrees Celsius. The X-31 also had movable forward canards for 

pitch control. As a result of controllability issues identified during the X-31 

flight-test program, fixed strakes between the trailing edge of the wing and 

the engine exhaust were incorporated. They provided additional nose-down 

pitch control at very high angles of attack. Another fix that was found necessary was the addition of small fixed-nose strakes to help control sideslip.136

During flight-test operations at the Rockwell Aerospace facility, the 

two X-31s flew 108 test missions, validating the use of thrust vectoring 

to compensate for loss of aerodynamic control at high angles of attack 

and expanding the poststall envelope up to 40 degrees angle of attack. 

The poststall envelope refers to the region in which the aircraft demonstrated an ability to maintain controlled flight beyond the normal 

X-31 stall angle of attack of 30 degrees. X-31 flight operations moved to 

NASA Dryden in February 1992, with the first flight under International 

Test Organization (ITO) management occurring in April 1992. The ITO 

initially included about 110 people from NASA, the U.S. Navy, the U.S. 

Air Force, Rockwell Aerospace, the Federal Republic of Germany, and 

135. “X-31 Enhanced Fighter Maneuverability Demonstrator Fact Sheet,” NASA Dryden Flight 

Research Center, 2001.

136. Chambers, Partners in Freedom.
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Daimler-Benz. The ITO staff was eventually reduced to approximately 

60 people. Overall management of the X-31 program came under by 

the Defense Advanced Research Projects Agency, with NASA responsible for flight-test operations, aircraft maintenance, and research engineering after the project moved to Dryden. The ITO director and NASA’s 

X-31 project manager at Dryden was Gary Trippensee. Pilots included 

NASA pilot Rogers Smith, U.S. Navy Cdr. Al Groves, German pilots Karl 

Lang and Dietrich Seeck, Rockwell International pilot Fred Knox, and 

Air Force Flight Test Center pilot Lt. Col. Jim Wisneski. By July 1992 

the X-31 flight envelope was being expanded in preparation for military utility evaluations that would fly the aircraft against nonthrust vectored fighters to evaluate effectiveness in simulated air combat. Thrust 

vectoring effectiveness at supersonic speed was evaluated out to Mach 

1.28 at an altitude of 35,000 feet.

In early flight-testing, the X-31 flight control system went into a 

reversionary mode four times in the first nine flights because of disagreement between the two air data sources.137 The X-31 was very sensitive to sideslip. This caused difficulties for the flight control system 

at higher angles of attack. Below 30 degrees, the nose boom updated 

the inertial navigation unit with air data. Above angles of attack of 30 

degrees, the inertial navigation unit began calculating erroneous sideslip angles as a result of changes in the relative wind vector. To resolve 

this problem, a so-called Kiel probe replaced the standard NASA Pitot 

tube to calculate airflow. The Kiel probe was bent 10 degrees downward 

from the standard pitot configuration. In addition, the sideslip vane was 

rotated downward 20 degrees relative to the nose boom to compensate 

for a yawing oscillation that occurred at an angle of attack of 62 degrees. 

These changes resulted in accurate air data being provided to the inertial navigation unit throughout the X-31 flight envelope with false sideslip readings at high angles of attack eliminated.138

Throughout the process of envelope expansion, many modifications 

to the flight control laws were required because actual aerodynamics 

of the aircraft were somewhat different from wind tunnel predictions. 

When the pilots started flying at angles of attack above 50 degrees, they 

137. Rushby, “Formal Methods and the Certification of Critical Systems,” SRI-CSL-93-07, SRI 

International, Nov. 1993. Rushby, “Formal Methods and the Digital Systems Validation for Airborne 

Systems,” NASA CR-4551 (Dec. 1, 1993). 

138. “X-31 Enhanced Fighter Maneuverability Demonstrator Fact Sheet.”
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encountered erratic lateral lurching movements. In an attempt to counter this phenomenon, narrow, 1/4-inch-wide strips of grit were attached 

to the sides of the nose boom and the radome. These effectively changed 

the vortex flow across the forward fuselage of the aircraft, reducing the 

randomness of the lurches and enabling expansion of the flight envelope 

to the design angle of attack limit of 70 degrees at 1 g. However, pilots 

encountered unintentional departures from controlled flight as the aircraft approached poststall angles of attack of 60 degrees during Split-S 

maneuvers.139 The asymmetric yawing moment encountered during 

this maneuver was beyond the capability of the thrust vectoring system 

to maintain adequate control.140 Testing in the Langley full-scale wind 

tunnel resulted in nose strakes and a modified slightly blunter nose tip 

design that were fitted to the two aircraft, allowing resumption of the 

flight-test program. The nose strakes were 6/10 of an inch wide and 20 

inches long and forced more symmetric transition of forebody vortexes. 

The blunted nose tip reduced yaw asymmetries.141

Poststall pitch control effectiveness, especially with the X-31 center 

of gravity at the aft allowable design location, was initially marginal.142 In 

these high-angle-of-attack conditions, test pilots rated aircraft response 

as unsatisfactory. NASA Langley conducted wind tunnel tests of various 

approaches intended to provide increased nose-down pitch control at 

high angles of attack. Sixteen different modifications were rapidly tested 

in the full-scale wind tunnel, with Langley recommending that a pair of 

strakes 6 inches wide and 65 inches long be mounted along the sides of 

the aft fuselage to assist in nose-down recovery. These were incorporated 

on the X-31, with subsequent flight-testing confirming greatly improved 

nose-down pitch control.143 Positive control at 70 degrees angle of attack 

with a controlled roll around the aircraft velocity vector was demonstrated 

November 6, 1992. On April 29, 1993, a minimum radius 180-degree post139. A Split-S maneuver consists of a 180-degree roll to inverted flight followed by a pull-through 

to erect level flight with the aircraft ending up positioned 180 degrees from its initial heading at a 

lower altitude.

140. Patrick C. Stoliker and Bosworth, “Evaluation of High-Angle-of-Attack Handling Qualities for 

the X-31A Using Standard Evaluation Maneuvers,” NASA TM-104322 (1996).

141. Chambers, Partners in Freedom.

142. Al Groves, Fred Knox, Rogers Smith, and Jim Wisneski, “X-31 Flight Test Update,” 37th Symposium Proceedings, Society of Experimental Test Pilots (SETP), Lancaster, CA, 1993, pp. 100–116.

143. Chambers, Partners in Freedom.
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stall “Herbst Maneuver” was accomplished for the first time.144 During the 

final phase of evaluation, the X-31s engaged in simulated air combat scenarios against F/A-18s. During these scenarios, the X-31s were able to 

outmaneuver the F/A-18s purely through use of poststall maneuvers and 

without use of thrust vectoring. X-31 test pilots did not support trading 

off basic fighter characteristics to acquire poststall maneuvering capabilities but concluded that improved pitch pointing and velocity vector 

maneuvering possible with thrust vector control did provide additional 

options during close-in combat. Thrust vectoring, combined with fully 

controllable poststall maneuvering, enabled X-31 pilots to position their 

aircraft in ways that adversary pilots could not counter, but it had to be 

used selectively and rapidly to be effective.145

In 1994, software was installed in the X-31 to simulate the feasibility of stabilizing a tailless aircraft at both subsonic and supersonic 

speed using thrust vectoring. The aircraft was modified to enable the 

pilot to destabilize the aircraft with the rudder to lower stability levels 

to those that would have been encountered if the aircraft had a reducedsize vertical tail. For this purpose, the rudder control surface was used 

to cancel the stabilizing effects of the vertical tail, and yaw thrust vector commands were applied by the flight control system to restabilize 

and control the aircraft. The X-31 was flown in the quasi-tailless mode 

supersonically at 38,000 feet at Mach 1.2, and maneuvers involving roll 

and yaw doublets, 30-degree bank-to-bank rolls, and windup turns to 2 

g were flown. During subsonic testing, simulated precision carrier landing approaches and ground attack profiles were successfully evaluated. 

The quasi-tailless flight-test experiment demonstrated the feasibility 

of tailless and reduced-tail highly maneuverable fighter/attack aircraft 

designs. Such designs could have reduced drag and lower weight as well 

as reduced radar and visual detectability. It determined that thrust vectoring is a viable flight control effector that can replace the functions 

provided by a vertical tail and rudder control surface. Potential disad144. A nonconventional poststall turning maneuver named in honor of Dr.-Ing. Wolfgang Herbst, 

who, as director of advanced design and technology at the Military Aircraft Division of MesserschmittBölkow-Blohm in Germany, had emphasized designs capable of poststall maneuvering. The author of 

numerous technical articles and papers dealing with preliminary aircraft design, aerodynamics, and 

maneuverability, Herbst died in the crash of a replica FW 190 fighter that he was flying in 1993.

145. Dave Canter, “X-31 Post-Stall Envelope Expansion and Tactical Utility Testing,” Fourth High Alpha 

Conference, NASA Dryden Flight Research Center, July 12–14, 1994 (NASA CP-10143), vol. 2.
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vantages include the added weight, complexity, and reliability issues 

associated with a thrust vectoring system. Additionally, flight conditions that require lower engine thrust settings (such as approach and 

landing) may necessitate provision of additional aerodynamic highdrag devices to enable high-thrust settings to be maintained, ensuring 

adequate thrust vectoring control. Early integration of such considerations into the overall design process, along with an increased level of 

interaction between propulsion and flight control systems, is required 

in order to derive the maximum benefit from reduced or tailless aircraft 

that rely on thrust vectoring for stability and control.146

The No. 1 X-31 aircraft was lost on its 292nd flight on January 19, 

1995. German test pilot Karl Lang had just finished a series of test maneuvers and was in the process of recovering back to a landing at Edwards. 

At an altitude of 20,000 feet, he observed discrepancies in the air data displays along with a master caution light. The aircraft then began a series 

of diverging pitch oscillations and became uncontrollable. Lang ejected 

safely at an altitude of 18,000 feet, and the aircraft crashed in an unpopulated desert area just north of Edwards. The crash was determined to 

have resulted from an unanticipated single-point failure in the nosemounted Kiel probe that provided critical airspeed and altitude data to 

the aircraft flight control system computers. These data were critical to 

safe flight, yet the Kiel probe did not include provision for electrical deicing, presumably because the aircraft would only be flown in clear desert 

weather conditions. However, during descent to recovery back to Edwards, 

ice accumulated in the unheated X-31 pitot tube, resulting in the flight 

control system automatically configuring the aircraft control surfaces for 

what it assumed were lower airspeed conditions. Unanticipated movements of the flight control surfaces caused the aircraft to begin oscillating 

about all axes followed by an uncontrolled pitch-up to an angle of attack 

of over 90 degrees.147 The subsequent X-31 accident investigation board 

recommended that training be conducted on the system safety analysis process, that procedures be implemented to ensure that all test team 

members receive configuration change notices, and that improvements 

be made in the remaining X-31 to prevent similar single-point failures.148

146. Bosworth and P.C. Stoliker, “The X-31A Quasi-Tailless Flight Test Results,” NASA TP-3624 (1996).

147. Don Haley, “Ice Cause of X-31 Crash,” NASA R-5-203, Dryden Flight Research Center (Nov. 

7, 1995).

148. Jay Levine, “X-31’s loss,” Dryden Flight Research Center X-Press, Jan. 2004.
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A panel that included former Dryden Research Center director Ken 

Szalai met at Dryden in early 2004 to review the X-31 accident. The 

panel noted that the primary contributing factor was the installation 

of the unheated Kiel probe in place of the original heated Pitot tube. 

The lack of electrical de-icing capability on the Kiel probe had not been 

considered a safety risk because X-31 mission rules prohibited flight in 

precipitation or clouds. However, there was no stipulation specifically 

restricting flight during potential icing conditions, despite simulations 

that showed icing of the Pitot static system could lead to loss of control.149 Information had been distributed among the X-31’s test pilots 

and flight-test engineers explaining the Pitot tube change, but a formal 

process was not in place to ensure that everyone fully understood the 

implications of the change. Test pilot Lang had noticed anomalies on his 

cockpit instrumentation and, assuming the presence of icing, told the 

control room that he was switching on Pitot heat. Shortly afterward, he 

advised that he was leaving the Pitot heat on for descent and approach to 

landing. The ground controller then told Lang that the pitot heat switch 

in the cockpit was not functional. Discrepancies between the X-31’s airspeed and altitude readouts were being observed in the control room, 

but that information was not shared with the entire control room staff. 

There was a redundant source of air data and a pilot-selectable alternative control mode that could have saved the aircraft if better communications had existed. Dryden X-31 project manager Gary Trippensee 

noted that complacency is the enemy of success in flight research; prior 

to the accident, 523 successful X-31 research missions had been flown.150

In 2000, the remaining X-31 was brought back from long-term 

storage at NASA Dryden, where it had been since 1995, and reconfigured for another round of flight-testing for the Vectoring, Extremely 

Short Takeoff and Landing Control and Tailless Operation Research 

(VECTOR) program. This program would explore the use of thrust vectoring for extremely short takeoff and landing (ESTOL), with a focus 

on the aircraft carrier environment. An international Cooperative 

Test Organization was created for the VECTOR program. U.S. participants/partners were the Navy, Boeing, General Electric, and NASA.151 

149. The Kiel probe was also found to be more susceptible to ice accumulation, a fact discovered after 

the accident through wind tunnel testing.

150. Ibid.

151. “X-31 VECTOR Program Phase I Begins,” NASA Dryden PR-98-09 (Mar. 9, 1998)
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The Swedish government was represented by Volvo and SAAB, with 

the German Ministry of Defense and DASA (Daimler-Benz consortium) from Germany. The X-31 aircraft was modified to incorporate a 

Swedish RM-1 engine, the same powerplant used in the Saab JAS-39 

Gripen fighter.152 On February 24, 2001, flown by U.S. Navy Cdr. Vivian 

Ragusa, the upgraded X-31 took to the air for the first time from Naval 

Air Station (NAS) Patuxent River.153 German test pilot Rüdiger “Rudy” 

Knöpfel, U.S. Marine Corps Maj. Cody Allee, and Navy Lt. J.R. Hansen 

would fly most of the subsequent ESTOL test program.154 The VECTOR 

X-31 went on to accomplish over 2 years’ of flight-testing, culminating 

in the final ESTOL flight by Maj. Allee on April 29, 2003.

From April 22 to 29, 2003, the VECTOR X-31 flew 11 test flights, 

during which fully automated, high-angle-of-attack approaches to landing were conducted. The automated flight control system utilized inputs 

from a special Global Positioning System (GPS)-based navigation system to maneuver the aircraft to a precise spot above the runway. Known 

as the Integrity Beacon Landing System (IBLS), it was supplemented 

by two virtual satellites, or “pseudolites,” on both sides of the runway. 

Precise spatial position and flight attitude data were inputs for the automatic approach control and landing system used in the VECTOR X-31. 

An ESTOL approach began with the pilot flying into the area covered 

by the pseudolites; after entering an engagement box, the automatic 

approach and landing system was activated. The aircraft then assumed 

a high-angle-of-attack approach attitude and followed a curvilinear 

path to the touchdown point. Just before touchdown, with the thrust 

vectoring paddles less than 2 feet above the runway, the X-31A automatically reduced its attitude back down to the normal 12-degree angle 

of attack for landing. An autothrottle system from an F/A-18 and a 

special autopilot developed by the VECTOR team were coupled with 

the flight control system to provide the integrated flight and propulsion control capability used to automatically derotate the aircraft 

from its steep final approach attitude to touchdown attitude at 2 feet 

above the runway.

152. The RM-1 was license manufactured by Volvo in Sweden and was a derivative of the General 

Electric GE F404 engine used in the F/A-18.

153. James Darcy, “Successful First Flight for VECTOR X-31 at Pax,” NAS Patuxent River Release 

01-040 (Feb. 28, 2001).

154. “EADS/Boeing X-31,” Flugrevue Magazine, Dec. 28, 2002.
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On the final flight of the VECTOR program, the angle of attack during 

landing approach was 24 degrees (twice the angle of attack on a normal 

landing approach). Approach airspeed was 121 knots, or about 30 percent lower than the normal 175 knots, and the resultant landing distance 

was only 1,700 feet, compared to the normal landing distance of nearly 

8,000 feet. Maj. Allee commented on the experience of riding along on a 

VECTOR X-31 automatic approach and landing: “There are no g forces 

and you sit leaning somewhat backwards in the ejection seat while the 

nose is pointing sharply upwards. . . . At angle of attacks greater than 15 

degrees the pilot cannot see the runway except on the screen on the righthand side of the instrument panel. . . . Whereas on a normal landing the 

landscape flashes by, now everything takes place as if in slow motion.”155

Another technical accomplishment demonstrated during the VECTOR 

X-31 program was the successful test of an advanced Flush Air Data 

System (FADS). Based on data collected by a dozen sensors located 

around the nose of the aircraft, the FADS provided accurate air data, 

including airspeed, altitude, angle of attack, and yaw angle, to the flight 

control system at angles of attack up to 70 degrees all the way out to 

supersonic speed.156

The two X-31 aircraft completed a total of 580 flights, a record for 

an X-plane program. Of these, 559 were research missions and 21 were 

flown in Europe in support of the 1995 Paris Air Show. Fourteen pilots 

from NASA, the U.S. Navy, the U.S. Marine Corps, the U.S. Air Force, 

the German Air Force, Rockwell International, and Deutsche Aerospace 

flew the aircraft during the original program at Palmdale and Dryden, 

with two U.S. pilots (one Navy and one Marine Corps) and a German 

pilot flying the VECTOR X-31 test program at Patuxent River. The surviving X-31, U.S. Navy Bureau No. 164585, flew 288 times, making its last 

flight on April 29, 2003. This aircraft is now on display at the Deutsches 

Museum annex at Oberschleißheim, near Munich, and it will eventually 

be returned to the United States. The other X-31, bureau No. 164584, 

had flown 292 times before it was lost on January 19, 1995.

Digital Electronic Engine Control

NASA pioneered in the development and validation of advanced computer-controlled electronic systems to optimize engine performance 

155. Karl Schwartz, “Successful ESTOL Landings with the X-31A,” Flugreview, July 2003.

156. J.R. Wilson, “X-31 Finds a Shorter Path to Success,” Aerospace America, Aug. 2003.
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The YF-12C used twin Pratt & Whitney J58 afterburning engines. NASA.

across the full flight envelope while also improving reliability. One such 

system was the Digital Electronic Engine Control (DEEC), whose genesis can be traced back to NASA Dryden work on the integrated flight and 

engine control system developed and evaluated in a joint NASA–Air Force 

program that used two Mach 3+ Lockheed YF-12C aircraft. The YF-12C 

was a cousin of the SR-71 strategic reconnaissance aircraft, and both aircraft used twin Pratt & Whitney J58 afterburning engines. As the SR-71 

neared Mach 3, a significant portion of the engine thrust was produced 

from the supersonic shock wave that was captured within each engine 

inlet and exited through the engine nozzle. A serious issue with the operational SR-71 fleet was so-called engine inlet unstarts. These occurred 

when the airflow into the inlet was not properly matched to that of the 

engine. This caused the standing shock wave normally located in the 

inlet to be expelled out the front of the SR-71’s inlet, causing insufficient pressure and airflow for normal engine operations. The result was 

a sudden loss of thrust on the affected engine. The resulting imbalance 

in thrust between the two SR-71 engines caused violent yawing, along 

with pitching and rolling motions. Studies showed that strong vortexes 

produced by each of the forward fuselage chines passed directly into the 
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inlets during the yawing motion produced by an unstart. NASA efforts supported development of a computerized automatic inlet sensing and cone 

control system and helped to optimize the ratio of air passing through 

the engine to that leaving the inlet through the forward bypass doors. 

Dryden successfully integrated the engine inlet control, auto-throttle, 

air data, and navigation functions to improve overall performance, with 

aircraft range being increased 7 percent. Handling qualities were also 

improved, and the frequency of engine inlet unstarts was greatly reduced. 

Pratt & Whitney and the Air Force incorporated the improvements demonstrated by Dryden into the entire SR-71 fleet in 1983.157 The Dryden 

YF-12C made its last NASA flight on October 31, 1979. On November 7, 

1979, it was ferried to the Air Force Museum at Wright-Patterson AFB, 

OH, where it is now on display.158

The broad objective of the DEEC program, conducted by NASA 

Dryden between 1981 and 1983, was to demonstrate and evaluate the 

system on a turbofan engine in a high-performance fighter across its full 

flight envelope. The program was a joint effort between Dryden, Pratt 

& Whitney, the Air Force, and NASA Lewis Research Center (now the 

NASA Glenn Research Center). The DEEC had been commercially developed by Pratt & Whitney based on its experience with the J58 engine 

during the NASA YF-12 flight research program. It integrated a variety of engine functions to improve performance and extend engine life. 

The DEEC system was tested on an F100 engine mounted in the left 

engine bay of a NASA Dryden McDonnell-Douglas F-15 fighter. Enginemounted and fuel-cooled, the DEEC was a single-channel digital controller. Engine inputs to the DEEC included compressor face static pressure 

and temperature, fan and core rotation speed, burner pressure, turbine 

inlet temperature, turbine discharge pressure, throttle position, afterburner fuel flow, and fan and compressor speeds. Using these inputs, 

the DEEC computer set the variable vanes, positioned the compressor 

air bleed, controlled gas-generator and augmentor fuel flows, adjusted 

the augmentor segment-sequence valve, and controlled the exhaust 

nozzle position. Thirty test missions that accumulated 35.5 flight hours 

were flown during the 2-year test program, which covered the operational 

157. “YF-12 Flight Research Program,” NASA Dryden Technology Fact Sheet TF-2004-17, NASA 

Dryden Flight Research Center (2004).

158. “Lockheed YF-12 NASA Fact Sheet,” FS-2002-09-047 DFRC, NASA Dryden Flight Research 

Center (2002).
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envelope of the F-15 at speeds up to Mach 2.36 and altitudes up to 60,000 

feet. The DEEC evaluation included nearly 1,300 throttle and afterburner transients, more than 150 air starts, maximum accelerations 

and climbs, and the full spectrum of flight maneuvers. An engine nozzle instability that caused stalls and blowouts was encountered when 

operating in afterburner at high altitudes. This instability had not been 

predicted in previous computer simulations or during ground-testing in 

NASA high-altitude test facilities. The instability problem was eventually 

resolved, and stall-free engine operation was demonstrated across the 

entire F-15 flight envelope. Faster throttle response, improved engine airstart capability, and an increase of more than 10,000 feet in the altitude 

that could be attained in afterburner without pilot restrictions on throttle use were achieved.159

DEEC-equipped engines were then installed on several operational 

USAF F-15s for service testing, during which they showed major improvements in reliability and maintainability. Mean time between failures was 

doubled, and unscheduled engine removals were reduced by a factor 

of nine. As a result, DEEC-equipped F100 engines were installed in all 

USAF F-15 and F-16 aircraft. The DEEC was a major event in the history of jet engine propulsion control and represented a significant transition from hydromechanical to digital-computer-based engine control. 

Performance improvements made possible by the DEEC included faster 

throttle responses, improved air-start capability, and an altitude increase 

of over 10,000 feet in afterburner without pilot restrictions on throttle 

use. Following the successful NASA test program, the DEEC went into 

standard use on F100 engines in the Boeing F-15 and the Lockheed F-16. 

Pratt & Whitney also incorporated digital engine control technology in 

turbofan engines used on some Boeing commercial jetliners. The lineage of similar digital engine control units used on other engines can 

be traced to the results of NASA’s DEEC test and evaluation program.160

Highly Integrated Digital Electronic Control

The Highly Integrated Digital Electronic Control (HIDEC) evolved from 

the earlier DEEC research effort. Major elements of the HIDEC were 

159. L.K. Myers, D.A. Mackall, and F.W. Burcham, Jr., “Flight Test Results of a Digital Electronic 

Engine Control System in an F-15 Airplane,” AIAA Paper 82-1080 (June 1982).

160. “Digital Electronic Engine Control,” NASA Dryden Technology Fact Sheet TF-2004-03, 

Dryden Flight Research Center (2004).
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a Digital Electronic Flight Control System (DEFCS), engine-mounted 

DEECs, an onboard general-purpose computer, and an integrated architecture that provided connectivity between components. The HIDEC 

F-15A (USAF serial No. 71-0287) was modified to incorporate DEECequipped F100 engine model derivative (EMD) engines. A dual channel Digital Electronic Flight Control System augmented the standard 

hydromechanical flight control system in the F-15A and replaced its analog control augmentation system. The DEFCS was linked to the aircraft 

data buses to tie together all other electronic systems, including the aircraft’s variable geometry engine inlet control system.161 Over a span of 

about 15 years, the HIDEC F-15 would be used to develop several modes 

of integrated propulsion and flight control systems. These integrated 

modes were Adaptive Engine Control System, Performance Seeking 

Control, Self-Repairing Flight Control System, and the Propulsion-Only 

Flight Control System. They are discussed separately in the following 

sections.162

Adaptive Engine Control System

The Adaptive Engine Control System (ADECS) improved engine performance by exploiting the excess stall margin originally designed into the 

engines using capabilities made possible with the integrated computerized flight and engine control systems.163 ADECS used airframe and 

engine data to allow the engine to operate at higher performance levels 

at times when inlet distortion was low and the full engine stall margin 

is not needed. Initial engineering work on ADECS began in 1983, with 

research flights beginning in 1986. Test results showed thrust improvements of between 8 and 10 percent depending on altitude. Fuel flow 

reductions of between 7 and 17 percent at maximum afterburning thrust 

at an altitude of 30,000 feet were recorded. Rate of climb increased 14 

percent at 40,000 feet. Time required to climb from 10,000 feet to 40,000 

feet dropped 13 percent. Acceleration improved between 5 and 24 percent at intermediate and maximum power settings, depending on altitude. 

No unintentional engine stalls were encountered in the test program. 

161. Frank W. Burcham, Lawrence P. Myers, and Ronald J. Ray, “Predicted Performance Benefits of 

an Adaptive Digital Engine Control System on an F-15 Airplane,” NASA TM-85916 (Jan. 1985).

162. “The HIDEC Program,” NASA Dryden Research Center Technology Fact Sheet, 1999.

163. Engine stall margin is a measure of the amount that engine operating pressure is reduced to 

provide a margin of safety to help prevent engine compressor stall due to excessive pressure.
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ADECS technology has been incorporated into the Pratt & Whitney F119 

engine used on the Air Force F-22 Raptor.164

Performance Seeking Control

The Performance Seeking Control (PSC) effort followed the Adaptive 

Electronic Control System project. Previous engine control modes utilized on the HIDEC aircraft used stored schedules of optimum engine 

pressure ratios based an average engine on a normal standard day. Using 

digital flight control, inlet control, and engine control systems, PSC used 

highly advanced computational techniques and control laws to identify 

the actual condition of the engine components and optimize the overall 

propulsion system for best efficiency based on actual engine and flight 

conditions that the aircraft was encountering, ensuring the highest 

engine and maneuvering performance in all flight environments. PSC 

testing with the HIDEC aircraft began in 1990. Results of flight-testing 

with PSC included increased fuel efficiency, improved engine thrust during accelerations and climbs, and increased engine service life achieved 

by reductions in turbine inlet temperature. Flight-testing demonstrated 

turbine inlet temperature reductions of more than 160 °F. Such large 

operating temperature reductions can significantly extend the life of jet 

engines. Additionally, improvements in thrust of between 9 percent and 

15 percent were observed in various flight conditions, including acceleration and climb.165 PSC also included the development of methodologies within the digital engine control system designed to detect engine 

wear and impending failure of certain engine components. Such information, coupled with normal preventative maintenance, could assist 

in implementing future fail-safe propulsion systems.166 The flight demonstration and evaluation of the PSC system at NASA Dryden directly 

contributed to the rapid transition of the technology into operational 

use. For example, PSC technology has been applied to the F100 engine 

164. Jennifer L. Baer-Riedhart and Robert J. Landy, “Highly Integrated Digital Electronic Control-Digital 

Flight Control, Aircraft Model Identification and Adaptive Engine Control,” NASA TM-86793 (Mar. 

1987); “The HIDEC Program,” NASA Dryden Flight Research Center Technology Fact Sheet, 1999.

165. Glen B. Gilyard and John S. Orme, “Subsonic Flight Test Evaluation of a Performance Seeking 

Algoritm on an F-15 Airplane,” NASA TM-4400 (Aug. 1992).

166. Baer-Riedhart, et al., “Highly Integrated Digital Electronic Control-Digital Flight Control, Aircraft 

Model Identification and Adaptive Engine Control”; “The HIDEC Program,” NASA Dryden Flight 

Research Center Technology Fact Sheet, 1999.
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used in the F-15 Eagle, the F119 engine in the F-22 Raptor, and the F135 

engine for the F-35 Lightning II.

Self-Repairing Flight Control System

The Self-Repairing Flight Control System (SRFCS) consists of software 

integrated into an aircraft’s digital flight control system that is used to 

detect failures or damage to the aircraft control surfaces. In the event 

of control surface damage, the remaining control surfaces are automatically reconfigured to maintain control, enabling pilots to complete their 

mission and land safely. The program, sponsored by the U.S. Air Force, 

demonstrated the ability of a flight control system to identify the failure 

of a control surface and reconfigure commands to other control devices, 

such as ailerons, rudders, elevators, and flaps, to continue the aircraft’s 

mission or allow it to be landed safely. As an example, if the horizontal 

elevator were damaged or failed in flight, the SRFCS would diagnose the 

failure and determine how the remaining flight control surfaces could 

be repositioned to compensate for the damaged or inoperable control 

surface. A visual warning to the pilot was used to explain the type of failure that occurred. It also provided revised aircraft flight limits, such as 

reduced airspeed, angle of attack, and maneuvering loads. The SRFCS 

also had the capability of identifying failures in electrical, hydraulic, 

and mechanical systems. Built-in test and sensor data provided a diagnostic capability and identified failed components or system faults for 

subsequent ground maintenance repair. System malfunctions on an aircraft with a SRFCS can be identified and isolated at the time they occur 

and then repaired as soon as the aircraft is on the ground, eliminating 

lengthy postflight maintenance troubleshooting.167

The SRFCS was flown 25 times on the HIDEC F-15 at NASA Dryden 

between December 1989 and March 1990, with somewhat mixed results. 

The maintenance diagnostics aspect of the system was a general success, but there were frequent failures with the SRFCS. Simulated control system failures were induced, with the SRFCS correctly identifying 

every failure that it detected. However, it only sensed induced control 

system failures 61 percent of the time. The overall conclusion was 

that the SRFCS concept was promising, but it needed more development if it was to be successfully implemented into production aircraft. 

167. “F-15 Flight Research Facility,” NASA Dryden Flight Research Center Fact Sheet FS-022, 1999.
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NASA test pilot Jim Smolka flew the first SRFCS flight, on December 

12, 1989, with test engineer Gerard Schkolnik in the rear cockpit; other 

SRFCS test pilots were Bill Dana and Tom McMurtry.168

Propulsion Controlled Aircraft System

Initiated in 1989, the Propulsion Controlled Aircraft (PCA) system was 

developed and flight-tested at NASA Dryden, with the goal being to help 

pilots land safely in the event that flight control components were disabled. PCA automatically provides computer-controlled variations in 

engine thrust that give pilots adequate pitch, yaw, and roll authority to 

fly their aircraft. The PCA system was tested and initially demonstrated 

on the HIDEC F-15. In simulator studies, NASA demonstrated the PCA 

concept on more than a dozen other commercial and military aircraft. 

The PCA system integrates the aircraft flight control and engine control computers to manage engine thrust and ensure adequate aircraft 

control. When the PCA system is activated, moving the control column 

aft causes the engine thrust to be automatically increased, and the aircraft begins to climb. Forward movement of the control column results 

in reduced thrust, and descent begins. Right or left movements of the 

control column produce differential engine thrust, resulting in the aircraft yawing in the direction of the desired turn. Flight-testing with the 

HIDEC F-15 was carried out at landing approach speeds of 150 knots 

with the flaps down and between 170 and 190 knots with the flaps 

retracted. At the conclusion of testing, the HIDEC F-15 accomplished a 

successful landing using the PCA system on April 21, 1993, after a flight 

in which the pilot used only engine power to turn, climb, and descend for 

approach to the runway.169

The NASA Dryden F-15A HIDEC testbed had originally been obtained 

from the Air Force in January 1976. During its career with NASA, it 

was involved in more than 25 advanced research projects involving 

aerodynamics, performance, propulsion control, systems integration, 

instrumentation development, human factors, and flight-test techniques 

before its last flight at Dryden, in October 1993.170

168. Tomayko, “The Story of Self-Repairing Flight Control Systems,” Dryden Historical Study, No. 1, 

Christian Gelzer, ed., 2003.

169. “F-15 Flight Research Facility.”

170. “F-15 HIDEC (Highly Integrated Digital Electronic Control),” NASA Dryden Flight Research 

Center Fact Sheet, 1999.
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A similar propulsion controlled aircraft approach was later evaluated 

and publicly demonstrated using a modified three-engine McDonnellDouglas MD-11 jet airliner in a cooperative program between NASA, 

McDonnell-Douglas, Pratt & Whitney, and Honeywell. Pratt & Whitney 

modified the engine control software, and Honeywell designed the software for the MD-11 flight control computer. Standard autopilot controls 

already present in the aircraft were used along with the Honeywell PCA 

software in the reprogrammed MD-11 flight control computers. NASA 

Ames performed computer simulations in support of the PCA program. 

On August 29, 1995, NASA Ames test pilot Gordon Fullerton successfully 

landed the PCA-modified MD-11 at Edwards AFB with an engine out 

after activating the aircraft’s auto-land system.171 Simulator testing of a 

PCA system continued using a NASA Ames–FAA motion-based Boeing 

747 simulator, with pilots making about 50 landings in the simulator. 

Additional simulation research by Ames resulted in further tests of the 

PCA system on B-747, B-757, MD-11, and C-17 aircraft. NASA Dryden 

test pilots flew simulated tests of the system in August 1998 in the NASA 

Ames Advanced Concepts Simulator. Ten pilots were involved in these 

tests, with 20 out of 20 attempted landings successfully accomplished. 

PCA technology can be used on current or future aircraft equipped with 

digital flight control systems.172

NF-15B Advanced Control Technology: Air Force S/MTD

NASA Dryden used an NF-15B research aircraft on various research projects from 1993 through early 2009. Originally designated the TF-15, it 

was the first two-seat F-15 Eagle built by McDonnell-Douglas, the sixth 

F-15 off the assembly line, and the oldest F-15 flying up to its retirement. First flown in July 1973, the aircraft was initially used for F-15 

developmental testing and evaluation as part of the F-15 combined test 

force at Edwards AFB in the 1970s. In the 1980s, the aircraft was extensively modified for the Air Force’s Short Takeoff and Landing Maneuver 

Technology Demonstrator (S/MTD) program. Modifications included 

the integration of a digital fly-by-wire control system, canards mounted 

on the engine inlets ahead of the wings,173 and two-dimensional thrustvectoring, thrust-reversing nozzles. The vectoring nozzles redirected 

171. Tomayko, “The Story of Self-Repairing Flight Control Systems.”

172. “Propulsion Controlled Aircraft,” NASA Dryden Fact Sheet FS-041, 1999.

173. The canards used on the NF-15B were based on the F/A-18 horizontal stabilators.
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The F-15 ACTIVE during yaw vectoring tests in 1996. NASA.

engine exhaust either up or down, giving greater pitch control and additional aerodynamic braking capability. Designated NF-15B to reflect 

its status as a highly modified research aircraft, the aircraft was used 

in the S/MTD program from 1988 until 1993. During Air Force S/MTD 

testing, a 25-percent reduction in takeoff roll was demonstrated with 

thrust-reversing, enabling the aircraft to stop in just 1,650 feet. Takeoffs 

using thrust-vectoring produced nose rotation speeds as low as 40 knots, 

resulting in greatly reduced takeoff distances. Additionally, thrustreversing produced extremely rapid in-flight decelerations, a feature 

valuable during close-in combat.174

NASA Advanced Control Technology for Integrated Vehicles

In 1994, after the conclusion of Air Force S/MTD testing, the aircraft was 

transferred to NASA Dryden for the NASA Advanced Control Technology 

for Integrated Vehicles (ACTIVE) research project. ACTIVE was oriented 

to determining if axisymmetric vectored thrust could contribute to drag 

reduction and increased fuel economy and range compared with conventional aerodynamic controls. The project was a collaborative effort 

between NASA, the Air Force Research Laboratory, Pratt & Whitney, 

174. Chambers, Partners in Freedom.
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and Boeing (formerly McDonnell-Douglas). An advanced digital flight 

fly-by-wire control system was integrated into the NF-15B, which was 

given NASA tail No. 837. Higher-thrust versions of the Pratt & Whitney 

F100 engine with newly developed axisymmetric thrust-vectoring engine 

exhaust nozzles were installed. The nozzles could deflect engine exhaust 

up to 20 degrees off centerline. This allowed variable thrust control in 

both pitch and yaw, or combinations of the two axes. An integrated propulsion and flight control system controlled both aerodynamic flight 

control surfaces and the engines. New cockpit controls and electronics from an F-15E aircraft were also installed in the NF-15B. The first 

supersonic flight using yaw vectoring occurred in early 1996. Pitch and 

yaw thrust vectoring were demonstrated at speeds up to Mach 2.0, and 

yaw vectoring was used at angles of attack up to 30 degrees. An adaptive 

performance software program was developed and successfully tested 

in the NF-15B flight control computer. It automatically determined the 

optimal setting or trim for the thrust-vectoring nozzles and the aerodynamic control surfaces to minimize aircraft drag. An improvement 

of Mach 0.1 in level flight was achieved at Mach 1.3 at 30,000 feet with 

no increase in engine thrust. The ACTIVE NF-15B continued investigations of integrated flight and propulsion control with thrust-vectoring 

during 1997 and 1998, including an experiment that combined thrust 

vectoring with aerodynamic controls during simulated ground attack 

missions. Following completion of the ACTIVE project, the NF-15B was 

used as a testbed for several other NASA Dryden research experiments, 

which included the efforts described below.175

High Stability Engine Control

NASA Lewis (now Glenn) Research Center evaluated an automated computerized engine control system that sensed and responded to high levels of engine inlet airflow turbulence to prevent sudden in-flight engine 

compressor stalls and potential engine failures. Known as High Stability 

Engine Control (HISTEC), the system used a high-speed digital processor 

to evaluate airflow data from engine sensors. The technology involved in 

the HISTEC approach was intended to control distortion at the engine 

face. The HISTEC system included two major functional subelements: a 

Distortion Estimation System (DES) and a Stability Management Control 

175. “NASA NF-15B Research Aircraft Fact Sheet,” Dryden Flight Research Center, Mar. 11, 2009.
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(SMC). The DES is an aircraft-mounted, high-speed computer processor. It uses state-of-the-art algorithms to estimate the amount and type 

of distortion present at the engine face based on measurements from 

pressure sensors in the engine inlet near the fan. Maneuver information from the digital flight control system and predictive angle-of-attack 

and angle-of-yaw algorithms are used to provide estimates of the type 

and extent of airflow distortion likely to be encountered by the engine. 

From these inputs, the DES calculates the effects of the engine face distortion on the overall propulsion system and determines appropriate 

fan and compressor pressure ratio commands. These are then passed 

to the SMC as inputs. The SMC performs an engine stability assessment 

using embedded stall margin control laws. It then issues actuator commands to the engine to best accommodate the estimated distortion.176

A dozen flights were flown on the ACTIVE F-15 aircraft at Dryden 

from July 15 to August 26, 1997, to validate the HISTEC concept, during which the system successfully directed the engine control computer 

to automatically command engine trim changes to adjust for changes 

in inlet turbulence level. The result was improved engine stability when 

inlet airflow was turbulent and increased engine performance when the 

airflow was stable.177

Intelligent Flight Control System

Beginning in 1999, the NF-15B supported the Intelligent Flight Control 

System (IFCS) neural network project. This was oriented to developing a 

flight control system that could identify aircraft characteristics through 

the use of neural network technology in order to optimize performance 

and compensate for in-flight failures by automatically reconfiguring 

the flight control system. IFCS is an extension of the digital fly-by-wire 

flight control system and is intended to maintain positive aircraft control under certain failure conditions that would normally lead to loss of 

control. IFCS would automatically vary engine thrust and reconfigure 

flight control surfaces to compensate for in-flight failures. This is accomplished through the use of upgrades to the digital flight control system 

software that incorporate self-learning neural network technology. A 

176. John C. DeLaat, George W. Gallops, Laura J. Kerr, Robert P. Kielb, John S. Orme, Robert D. 

Southwick, and Mark G. Walsh, “High Stability Engine Control (HISTEC) Flight Test Results,” NASA 

TM-1998-208481 (July 1998).

177. Ibid.
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neural network that could train itself to analyze flight properties of an 

aircraft was developed, integrated into the NASA NF-15B, and evaluated 

in flight testing. The neural network “learns” aircraft flight characteristics in real time, using inputs from the aircraft sensors and from error 

corrections provided by the primary flight control computer. It uses this 

information to create different aircraft flight characteristic models. The 

neural network learns to recognize when the aircraft is in a stable flight 

condition. If one of the flight control surfaces becomes damaged or nonresponsive, the IFCS detects this fault and changes the flight characteristic model for the aircraft. The neural network then drives the error 

between the reference model and the actual aircraft state to zero. Dryden 

test pilot Jim Smolka flew the first IFCS test mission on March 19, 1999, 

with test engineer Gerard Schkolnik in the rear cockpit.178

The NF-15B IFCS test program provided the opportunity for a limited 

flight evaluation of a direct adaptive neural network-based flight control 

system.179 This effort was led by the Dryden Flight Research Center, with 

collaboration from the Ames Research Center, Boeing, the Institute for 

Scientific Research at West Virginia University, and the Georgia Institute 

of Technology.180 John Bosworth was the NASA Dryden IFCS chief engineer. Flight-testing of the direct adaptive neural network-based flight 

control system began in 2003 and evaluated the outputs of the neural 

network. The neural network had been pretrained using flight characteristics obtained for the F-15 S/MTD aircraft from wind tunnel testing. 

During this phase of testing, the neural network did not actually provide any flight control inputs in-flight. The outputs of the neural network 

were run directly to instrumentation for data collection purposes only.

In 2005, a fully integrated direct adaptive neural-network-based 

flight control system demonstrated that it could continuously provide 

error corrections and measure the effects of these corrections in order to 

learn new flight models or adjust existing ones. To measure the aircraft 

state, the neural network took a large number of inputs from the roll, 

pitch, and yaw axes and the aircraft’s control surfaces. If differences were 

detected between the measured aircraft state and the flight model, the 

neural network adjusted the outputs from the primary flight computer 

178. Tomayko, “The Story of Self-Repairing Flight Control Systems.”

179. Peggy S. Williams-Hayes, “Flight Test Implementation of a Second Generation Intelligent Flight 

Control System,” NASA TM-2005-213669 (Nov. 2005).

180. “Intelligent Flight Control System,” NASA Dryden Fact Sheet FS-076, Feb. 13, 2006.
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to bring the differences to zero before they were sent to the actuator 

control electronics that moved the control surfaces.181 IFCS software 

evaluations with the NF-15B included aircraft handling qualities maneuvers, envelope boundary maneuvers, control surface excitations for 

real-time parameter identification that included pitch, roll, and yaw 

doublets, and neural network performance assessments.182 During 

NF-15B flight-testing, a simulated failure was introduced into the right 

horizontal stabilizer that simulated a frozen pitch control surface. 

Handling qualities were evaluated with and without neural network 

adaptation. The performance of the adaptation system was assessed in 

terms of its ability to decouple roll and pitch response and reestablish 

good onboard model tracking. Flight-testing with the simulated stabilator 

failure and the adaptive neural network flight control system adaptation 

showed general improvement in pitch response. However, a tendency 

for pilot-induced roll oscillations was encountered.183

Concurrent with NF-15B IFCS flight-testing, NASA Ames conducted 

a similar neutral network flight research program using a remotely controlled Experimental Air Vehicle (EAV) equipped with an Intelligent 

Flight Controller (IFC). Aerodynamically, the EAV was a one-quarter-scale 

model of the widely used Cessna 182 Skylane general aviation aircraft. 

The EAV was equipped with two electrical power supplies, one for the 

digital flight control system that incorporated the neural-network IFC 

capability and one for the avionics installation that included three video 

cameras to assist the pilots with situation awareness. Several pilots flew 

the EAV during the test program. Differences in individual pilot control 

techniques were found to have a noticeable effect on the performance 

of the Intelligent Flight Controller. Interestingly, IFCS flight-testing with 

the NF-15B aircraft uncovered many of the same issues related to the 

controller that the EAV program found. IFCS was determined to provide increased stability margins in the presence of large destabilizing 

failures. The adaptive system provided better closed-loop behavior with 

181. M.G. Perhinschi, J. Burken, M.R. Napolitano, G. Campa, and M.L. Fravolini, “Performance 

Comparison of Different Neural Augmentation for the NASA Gen - 2 IFCS F-15 Control Laws,” 

Proceedings of the 2004 American Control Conference, Boston, MA, 2004.

182. Creech, “NASA Dryden Neural Network Passes Milestone,” NASA Dryden Press Release, 

Sept. 3, 2003.

183. Bosworth and Williams-Hayes, “Flight Test Results from the NF-15B Intelligent Flight Control System 

(IFCS) Project with Adaptation to a Simulated Stabilator Failure,” AIAA Paper 2007-2818, 2007.
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improved matching of the onboard reference model. However, the convergent properties of the controller were found to require improvement 

because continued maneuvering caused continued adaptation change. 

During ground simulator evaluation of the IFCS, a trained light-plane 

pilot was able to successfully land a heavily damaged large jet airliner 

despite the fact that he had no experience with such an aircraft. Test 

data from the IFCS program provided a basis for analysis and understanding of neural network-based adaptive flight control system technology as an option for implementation into future aircraft.184 

After a 35-year career, during which it had flown with McDonnellDouglas, the Air Force, and NASA, the NF-15B was retired following its 

final flight, on January 30, 2009. During its 14 years at NASA Dryden, 

the aircraft had flown 251 times. The NF-15B will be on permanent display with a group of other retired NASA research aircraft at Dryden.185

The Continuing Legacy of FBW Research in Aircraft Development

Fly-by-wire technology developed by NASA and the Air Force served 

as the basis for flight control systems in several generations of military 

and civilian aircraft. Many of these aircraft featured highly unconventional airframe configurations that would have been unflyable without 

computer-controlled fly-by-wire systems. An interesting example was 

the then highly classified Lockheed Have Blue experimental stealth 

technology flight demonstrator. This very unusual aircraft first flew in 

1977 and was used to validate the concept of using a highly faceted airframe to provide a very low radar signature. Unstable about multiple 

axes, Have Blue was totally dependent on its computer-controlled flyby-wire flight control system that was based on that used in the F-16. 

Its success led to the rapid development and early deployment of the 

stealthy Lockheed F-117 attack aircraft that first flew in 1981 and was 

operational in 1983.186 More advanced digital fly-by-wire flight control 

systems enabled an entirely new family of unstable, aerodynamically 

refined “stealth” combat aircraft to be designed and deployed. These 

184. Cory Ippolito, John Kaneshige, and Yoo-Hsiu Yeh, “Neural Adaptive Flight Control Testing on 

an Unmanned Experimental Aerial Vehicle,” AIAA Paper 2007-2827, 2003; Tomayko, “The Story 

of Self-Repairing Flight Control Systems.”

185. “NASA NF-15B Research Aircraft Fact Sheet.”

186. Aronstein and Piccirillo, “Have Blue and the F-117A, Evolution of the “Stealth Fighter,” American Institute of Aeronautics and Astronautics, Reston, VA, 1997.
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include the Northrop B-2 Spirit flying wing bomber and Lockheed’s F-22 

Raptor and F-35 Lightning II fighters with their highly integrated digital propulsion and flight control systems.

Knowledge of the benefits and confidence in the use of digital fly-bywire technology are today widespread across the international aerospace 

industry. Nearly all new military aircraft—including fighters, bombers, 

and cargo aircraft, as well as commercial airliners, both U.S. and foreign—have reaped immense benefits from the legacy of NASA’s pioneering digital fly-by-wire flight and propulsion control efforts. On the airlift 

side, the Air Force’s Boeing C-17 was designed with a quad-redundant 

digital fly-by-wire flight control system.187 In Europe, Airbus Industrie 

was an early convert to digital fly-by-wire and the increasing use of electronic subsystems. All of its airliners, starting with the A320 in 1987, were 

designed with fully digital fly-by-wire flight control architectures along 

with side stick controllers.188 Reliance on complex and heavy hydraulic 

systems is being reduced as companies increase the emphasis on electrically powered flight controls. With this approach, both electrical and 

self-contained electrohydraulic actuators are controlled by the digital 

flight control system’s computers. The benefits are lower weight, reduced 

maintenance cost, the ability to provide redundant electrical power circuits, and improved integration between the flight control system and 

the aircraft’s avionics and electrical subsystems. Electric flight control 

technology reportedly resulted in a weight reduction of 3,300 pounds in 

the A380 compared with a conventional hydromechanical flight control 

system.189 Boeing introduced fly-by-wire with its 777, which was certified 

for commercial airline service in 1995. It has been in routine airline service with its reliable digital fly-by-wire flight control system ever since. 

In addition to a digital fly-by-wire flight control system, the next Boeing 

airliner, the 787, incorporates some electrically powered and operated 

flight control elements (the spoilers and horizontal stabilizers). These 

are designed to remain functional in the event of either total hydraulic 

systems failure or flight control computer failure, allowing the pilots 

to maintain control in pitch, roll, and yaw and safely land the aircraft.

187. Brian W. Kowal, Carl J. Scherz, and Richard Quinliven, “C-17 Flight Control System Overview,” IEEE Aerospace and Electronic Systems Magazine, July 1992.

188. C. Favre, “Fly-by-Wire for Commercial Aircraft: the Airbus Experience,” International Journal of 

Control, vol. 59, no. 1, 1994, pp. 139–157.

189. “All-Electric Aircraft Research Speeds Up,” Aerospace America, Jan. 2009, pp. 4–8.
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Today, the tremendous benefits made possible by the use of digital 

fly-by-wire in vehicle control systems have migrated into a variety of 

applications beyond the traditional definition of aerospace systems. As 

a significant example, digital fly-by-wire ship control systems are now 

operational in the latest U.S. Navy warships, such as the Seawolf and 

Virginia class submarines. NASA experts, along with those from the FAA 

and military and civil aviation agencies, supported the Navy in developing its fly-by-wire ship control system certification program.190 Thus, the 

vision of early advocates of digital fly-by-wire technology within NASA 

has been fully validated. Safe and efficient, digital fly-by-wire technology is today universally accepted with its benefits available to the military services, airline travelers, and the general public on a daily basis.

190. William Palmer, “Submarine Fly-by-Wire Ship Control System Certification Program Leverages 

Aviation Industry Best Practice,” Wavelengths Online, The Official Publication of the Naval Surface 

Warfare Center, Carderock Division, Oct. 1, 2004.
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Recommended Additional Readings

Reports, Papers, Articles, and Presentations:

“All-Electric Aircraft Research Speeds Up,” Aerospace America, Jan. 2009.

B.R. Ashworth and William M. Kahlbaum, Jr., “Description and 

Performance of the Langley Differential Maneuvering Simulator,” 

NASA TN-D-7304 (June 1973).

“Automatic Flight Control System Sought,” Aviation Week, Aug. 6, 1956.

Jennifer L. Baer-Riedhart and Robert J. Landy, “Highly Integrated Digital 

Electronic Control-Digital Flight Control, Aircraft Model Identification 

and Adaptive Engine Control,” NASA TM-86793 (Mar. 1987).

R.E. Bailey, M.F. Schaeffer, R.E. Smith, and J.F. Stewart, “Flight Test 

Experience With Pilot-Induced-Oscillation Suppression Filters,” 

NASA TM-86028 (Jan. 1984).

Charles A. Baird and Franklin B. Snyder, with introduction by Lt. Mark 

Bierele, “Terrain-Aided Altitude Computations on the AFTI/F-16,” 

Harris Corporation, Melbourne, FL, Aug. 1990.

Finley Barfield, Duke Browning, and Judith Probert, “All Terrain Ground 

Collision Avoidance and Maneuver Terrain Following for Automated 

Low Level Night Attack,” IEEE AES Systems Magazine, March 1993.

Joseph F. Baugher, “General Dynamics YF-16/CCV,” American Military 

Aircraft, Mar. 31, 2000.

Jeffrey E. Bauer, Robert Clarke, and John J. Burken, “Flight Test of the 

X-29A at High Angle of Attack: Flight Dynamics and Controls,” NASA 

TP-3537 (Feb. 1995).

James Blaylock, Donald Swihart, and William Urshel, “Integration of 

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Case 10 | Fly-By-Wire: Making the Electric Jet

725

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John T. Bosworth and P.C. Stoliker, “The X-31A Quasi-Tailless Flight 

Test Results,” NASA TP-3624 (1996).

John T. Bosworth and Peggy S. Williams-Hayes, “Flight Test Results 

from the NF-15B Intelligent Flight Control System (IFCS) Project 

with Adaptation to a Simulated Stabilator Failure,” AIAA Paper 

2007-2818 (2007).

Frank W. Burcham, Lawrence P. Myers, and Ronald J. Ray, “Predicted 

Performance Benefits of an Adaptive Digital Engine Control System 

on an F-15 Airplane,” NASA TM-85916 (Jan. 1985).

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vol. 43, issue 2, Feb. 23, 2001.

Gary Creech, “NASA Dryden Neural Network Passes Milestone,” NASA 

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NASA’s Contributions to Aeronautics

726

10

John C. DeLaat, George W. Gallops, Laura J. Kerr, Robert P. Kielb, John 

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Dennis R. Furman, “USAF Test Pilot School Use of DIGITAC in Systems 

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851827, Oct. 1985.

“General Dynamics and Northrop to Build Lightweight Fighter 

Prototypes,” Interavia, July 1972.
Case 10 | Fly-By-Wire: Making the Electric Jet

727

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Glen B. Gilyard and John S. Orme, “Subsonic Flight Test Evaluation 

of a Performance Seeking Algoritm on an F-15 Airplane,” NASA 

TM-4400 (Aug. 1992).

Yefim Gordon, “An Industry of Prototypes: Sukhoi T-4, Russia’s Mach 

3 Bomber,” Wings of Fame, vol. 9, Aerospace Publishing Limited, 

London (1997).

Al Groves, Fred Knox, Rogers Smith, and Jim Wisneski, “X-31 Flight 

Test Update,” 37th Symposium Proceedings, Society of Experimental 

Test Pilots (SETP), Lancaster, CA, 1993.

R. Guyton and F. Luria, “Flight Testing of Pneumatic Forebody 

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Don Haley, “Ice Cause of X-31 Crash,” NASA Press Release 5-203, Dryden 

Flight Research Center, Nov. 7, 1995.

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a Flight Demonstrator,” Proceedings of AIAA Aircraft Prototype and 

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J.K.B. Illingworth and H.W. Chinn, “Variable Stability and Control Tests 

on the S.C.1 Aircraft in Jet-Borne Flight, with Particular Emphasis 

on Reference to Desirable VTOL Flying Qualities,” Royal Aircraft 

Establishment, Bedford, U.K., Her Majesty’s Stationery Office, London, 

1969.

Cory Ippolito, John Kaneshige, and Yoo-Hsiu Yeh, “Neural Adaptive 

Flight Control Testing on an Unmanned Experimental Aerial Vehicle,” 

AIAA Paper 2007-2827, 2007.

Stephen D. Ishmael, Dale A. Mackall, and Victoria A. Regenie, “Design 

Implications From AFTI/F-16 Flight Test,” NASA TM-86026, 1984.
NASA’s Contributions to Aeronautics

728

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Stephen D. Ishmael and Donald R. McMonagle, “AFTI/F-16 Flight Test 

Results and Lessons,” NASA TM-84920 (Oct. 1983).

Gavin D. Jenny, “JB-47E Fly-By-Wire Flight Test Program (Phase I),” 

AFFDL TR-69-40, Air Force Flight Dynamics Laboratory, WrightPatterson AFB, OH, Sept. 1969.

Robert W. Kempel, “Flight Experience with a Backup Flight Control System 

for the HiMAT Research Vehicle,” AIAA Paper 82-1541 (Aug. 1982).

Robert W. Kempel and Michael R. Earls, “Flight Control Systems 

Development and Flight Test Experience with the HiMAT Research 

Vehicles,” NASA TP-2822 (June 1988).

Brian W. Kowal, Carl J. Scherz, and Richard Quinliven, “C-17 Flight 

Control System Overview,” IEEE Aerospace and Electronic Systems 

Magazine, July 1992.

D. Lean and H.W. Chinn, “Review of General Operating Experience 

with a Jet-Lift VTOL Research Aircraft (Short S.C.1),” Aeronautical 

Research Council Current Paper No. 832, Her Majesty’s Stationary 

Office, London (1965).

Robert E. Lemble, “DIGITAC—A Unique Digital Flight Control Testbed 

Aircraft,” AIAA-1990-1288, SFTE, DGLR, and SETP Biannual Flight 

Test Conference, 5th, Ontario, CA, May 22–24, 1990.

Jay Levine, “X-31’s loss,” Dryden Flight Research Center X-Press, Jan. 2004.

Dale A. Mackall, “AFTI/F-16 Digital Flight Control System Experience,”

in Gary P. Beasley, ed., NASA Aircraft Controls Research 1983, pp. 

469–487, NASA CP-2296, 1984; Proceedings of Workshop Held at 

NASA Langley Research Center, Oct. 25–27, 1983.

Dale A. Mackall, “Development and Flight Test Experiences with a FlightCrucial Digital Control System,” NASA TP-2857 (Nov. 1988).

D.L. Martin and D. Gangsaas, “Testing the YC-14 Flight Control System 

Software,” AIAA Journal of Guidance and Control, July–Aug. 1978.
Case 10 | Fly-By-Wire: Making the Electric Jet

729

10

Duane McRuer and Dunstan Graham, “A Flight Control Century: 

Triumphs of the Systems Approach,” Journal of Guidance, Control 

and Dynamics, vol. 27, no. 2, pp. 161–173, AIAA, 2003.

L.K. Myers, D.A. Mackall, and F.W. Burcham, Jr., “Flight Test Results 

of a Digital Electronic Engine Control System in an F-15 Airplane,” 

AIAA Paper 82-1080 (June 1982).

Robert Navarro, “Performance of an Electro-Hydrostatic Actuator on 

the F-18 Systems Research Aircraft,” NASA TM-97-206224 (1997).

Molly Neal, “VC10: Vickers-Armstrongs’ Long-range Airliner,” Flight 

International, May 10, 1962.

William Palmer, “Submarine Fly-by-Wire Ship Control System 

Certification Program Leverages Aviation Industry Best Practice,”

Wavelengths Online, Naval Surface Warfare Center, Carderock 

Division, Oct. 1, 2004.

M.G. Perhinschi, J. Burken, M.R. Napolitano, G. Campa, and M.L. Fravolini, 

“Performance Comparison of Different Neural Augmentation for the 

NASA Gen – 2 IFCS F-15 Control Laws,” Proceedings of the 2004 

American Control Conference, Boston, MA, 2004.

Terrill W. Putnam, “X-29 Flight-Research Program,” NASA TM-86025 

(Jan. 1984).

RAE Electric Hunter,” Flight International, June 28, 1973.

James W. Ramsey, “Power-by-Wire,” Avionics Magazine, May 1, 2001.

H.A. Rediess and E.C. Buckley, “Technology Review of Flight Crucial 

Flight Control Systems (Application of Optical Technology),” NASA 

CR-172332 (Supplement 1), Sept. 1984.

John Rushby, “Formal Methods and the Certification of Critical Systems,” 

SRI-CSL-93-07, SRI International, Nov. 1993.

“
NASA’s Contributions to Aeronautics

730

10

John Rushby, “Formal Methods and the Digital Systems Validation for 

Airborne Systems,” NASA CR-4551 (Dec. 1, 1993).

Karl Schwartz, “Successful ESTOL Landings with the X-31A,” Flugreview, 

July 2003.

Katsuhei Shibata and Hideaki Ohmiya, “The T-2 Control Configured 

Vehicle Development, Integration and Flight Test,” AIAA-1988-3882, 

published in Technical Papers, Part 1 (A89-18051 05-06) AIAA/IEEE 

Digital Avionics Systems Conference, San Jose, CA, Oct. 17–20, 1988.

S.A. Sjoberg, “Some Experience With Side Controllers,” Research Airplane 

Committee Report on the Progress of the X-15 Project, Conference 

held at NACA Langley Aeronautical Laboratory, Langley Field, VA, 

Oct. 25–26, 1956.

Patrick C. Stoliker and John T. Bosworth, “Evaluation of High-Angle-ofAttack Handling Qualities for the X-31A Using Standard Evaluation 

Maneuvers,” NASA TM-104322 (1996).

Joe Stout, “What a Wonderful Airplane: YF-16 First Flight,” Code One 

Magazine, July 1992.

Kenneth J. Szalai, Calvin R. Jarvis, Gary E. Krier, Vincent A. Megna, 

Larry D. Brock, and Robert N. O’Donnell, “Digital Fly-by-Wire Flight 

Control Validation Experience,” NASA TM-72860 (Dec. 1978).

L.W. Taylor, Jr., and E.J. Adkins, “Adaptive Flight Control Systems—Pro 

and Con,” NASA TM-X-56008 (1964).

L.W. Taylor, Jr., and G.B. Merrick, “X-15 Airplane Stability Augmentation 

System,” NASA TN-D-1157 (1962).

L.W. Taylor, Jr., and J.W. Smith, “An Analysis of the Limit-Cycle 

and Structural-Resonance Characteristics of the X-15 Stability 

Augmentation System,” NASA TN-D-4287 (1967).

James E. Tomayko, “Blind Faith: The United States Air Force and the 

Development of Fly-By-Wire Technology,” Technology and the Air 
Case 10 | Fly-By-Wire: Making the Electric Jet

731

10

Force: A Retrospective Assessment, Air Force History and Museums 

Program, Washington, DC, 1997.

Robert A. Tremant, “Operational Experience and Characteristics of the 

X-15 Flight Control System,” NASA TN-D-1402 (1962).

Robert B. Voas, “Manned Control of Mercury Spacecraft,” Astronautics, 

vol. 7, no. 3, Mar. 1962.

Kyrill Von Gersdorff, “Transfer of German Aeronautical Knowledge 

After 1945,” Aeronautical Research in Germany (From Lilienthal to 

Today), Springer-Verlag, Berlin Heidelberg, 2004.

Norman C. Weingarten, “History of In-Flight Simulation & Flying 

Qualities Research at CALSPAN,” AIAA Journal of Aircraft, vol. 42, 

no. 2, Mar./Apr. 2005.

M.J. Wendle, G.G. Grose, J.L. Porter, and V.R. Pruitt, Flight/Propulsion 

Control Integration Aspects of Energy Management, SAE Paper 

Number 740480, 1974.

Peggy S. Williams-Hayes, “Flight Test Implementation of a Second 

Generation Intelligent Flight Control System,” NASA TM-2005-

213669, Nov. 2005.

J.R. Wilson, “X-31 Finds a Shorter Path to Success,” Aerospace America, 

Aug. 2003.

B.S. Wolfe, “The Concorde Automatic Flight Control System: A 

Description of the Automatic Flight Control System for the AngloFrench Supersonic Transport and its Development to Date,” Aircraft 

Engineering and Aerospace Technology, vol. 39, issue 5, 1967.

“X-31 VECTOR Program Phase I Begins,” NASA Dryden Press Release 

98-09, Mar. 9, 1998.

Michael L. Yaffee, “Survivable Controls Gain Emphasis,” Aviation Week, 

Feb. 2, 1970.
NASA’s Contributions to Aeronautics

732

10

C.J. Yeo, “The Fly-by-Wire Jaguar,” Society of Experimental Test Pilots 

(SETP) 27th Symposium Proceedings, Beverly Hills, CA, Sept. 28–

Oct. 1, 1983.

Books and Monographs:

David C. Aronstein, Michael J. Hirschberg and Albert C. Piccirillo, 

Advanced Tactical Fighter to F-22 Raptor (Reston: AIAA, 1998).

David C. Aronstein and Albert C. Piccirillo, Have Blue and the F-117A, 

Evolution of the Stealth Fighter (Reston: AIAA, 1997).

David C. Aronstein and Albert C. Piccirillo, The Lightweight Fighter 

Program: A Successful Approach to Fighter Technology Transition 

(Reston: AIAA, 1996).

Robert E. Ball, The Fundamentals of Aircraft Combat Survivability 

Analysis and Design (New York: AIAA, 1985).

Paul F. Borchers, James A. Franklin, and Jay W. Fletcher, Flight Research 

at Ames, 1940–1997: Fifty Seven Years of Development and Validation 

of Aeronautical Technology, NASA SP-3300 (Washington: NASA, 1998).

Joseph R. Chambers, Partners in Freedom: Contributions of the Langley 

Research Center to U.S. Military Aircraft of the 1990s, SP-2000-4519 

(Washington: NASA, 2000).

C. Droste and J. Walker, The General Dynamics Case Study on the F-16 

Fly-By-Wire Flight Control System (New York: AIAA, undated).

Richard P. Hallion, On the Frontier: Flight Research at Dryden, 1946–

1981, SP-4303 (Washington: NASA, 1984).

E.H. Herschel, H. Prem, and G. Madelung, Aeronautical Research in 

Germany (From Lilienthal to Today) (Berlin Heidelberg: SpringerVerlag, 2004).

Dennis R. Jenkins, X-15: Extending the Frontiers of Flight, SP-2007-562 

(Washington: NASA, 2007).
Case 10 | Fly-By-Wire: Making the Electric Jet

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Gavin D. Jenny, James W. Morris, and Vernon R. Schmitt, Fly-by-Wire, A 

Historical and Design Perspective (Warrendale: SAE, 1998).

Lawrence K. Loftin, Quest for Performance: The Evolution of Modern 

Aircraft, Part II: The Jet Age, Chapter 12: Jet Bomber and Attack Aircraft, 

NASA SP-468 (Washington: NASA, 1985).

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Pilot’s Odyssey, NASA SP-4108 (Washington: NASA, 2003).

NASA Office of Safety and Mission Assurance, Formal Methods 

Specification and Verification Guidebook for Software and Computer 

Systems, Volume I: Planning and Technology Insertion,” P-98-208193 

(Washington: NASA, 1998).

NASA Office of Safety and Mission Assurance, Formal Methods 

Specification and Analysis Guidebook for the Verification of Software 

and Computer Systems, Volume II: A Practitioner’s Companion, NASA 

GB-001-97 (Washington: NASA, 1997).

Michael J. Neufeld, The Rocket and the Reich; Peenemünde and the Coming 

of the Ballistic Missile Era (Washington: Smithsonian Institution, 1995).

Albert C. Piccirillo, German V/STOL Fighter Program, (Reston: AIAA, 1997).

Rowland F. Pocock, German Guided Missiles of the Second World War, 

(New York: Arco publishing company, 1967)

Ray Sturtivant, British Research and Development Aircraft: Seventy Years 

at the Leading Edge (UK: Haynes/Foulis, 1990).

James E. Tomayko, Computers Take Flight: A History of NASA’s Pioneering 

Digital Fly-by Wire Project, NASA SP-2000-4224 (Washington: NASA, 

2000).

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Dryden Historical Study No.1, Christian Gelzer, ed. (Washington: 

NASA, 2003).
734

An NACA cowling on a Northrop A-17A at the Agency’s Langley Laboratory in 1939. The NACA 

cowling was its signal contribution to integrating propulsive and aerodynamic design. NASA.
735

11

Advancing Propulsive 

Technology

James Banke

CASE

11

Ensuring proper aircraft propulsion has been a powerful stimulus. In the 

interwar years, the NACA researched propellers, fuels, engine cooling, supercharging, and nacelle and cowling design. In the postwar 

years, the Agency refined gas turbine propulsion technology. NASA 

now leads research in advancing environmentally friendly and fuelconserving propulsion, thanks to the Agency’s strengths in aerodynamic 

and thermodynamic analysis, composite structures, and other areas.

E

ACH DAY, OUR SKIES FILL with general aviation aircraft, business jets, 

and commercial airliners. Every 24 hours, some 2 million passengers worldwide are moved from one airport to the next, almost all 

of them propelled by relatively quiet, fuel-efficient, and safe jet engines.1

And no matter if the driving force moving these vehicles through the 

air comes from piston-driven propellers, turboprops, turbojets, turbofans—

even rocket engines or scramjets—the National Aeronautics and Space 

Administration (NASA) during the past 50 years has played a significant 

role in advancing that propulsion technology the public counts on every day.

Many of the advances seen in today’s aircraft powerplants can trace 

their origins to NASA programs that began during the 1960s, when the 

Agency responded to public demand that the Government apply major 

resources to tackling the problems of noise pollution near major airports. 

Highlights of some of the more noteworthy research programs to reduce 

noise and other pollution, prolong engine life, and increase fuel efficiency 

will be described in this case study.

But efforts to improve engine efficiency and curb unwanted noise actually predate NASA’s origins in 1958, when its predecessor, the National 

Advisory Committee for Aeronautics (NACA), served as the Nation’s preeminent laboratory for aviation research. It was during the 1920s that 

1. William H. More, ed., National Transportation Statistics (Washington, DC: U.S. Department of 

Transportation, 2009), p. 72.
NASA’s Contributions to Aeronautics

736

11

Air pollution is evident as this Boeing B-47B takes off in 1954 with the help of its General Electric 

J47 jet engines and Rocket Assisted Take Off solid rocket motors. U.S. Air Force.

the NACA invented a cowling to surround the front of an airplane and its 

radial engine, smoothing the aerodynamic flow around the aircraft while 

also helping to keep the engine cool. In 1929, the NACA won its first Collier 

Trophy for the breakthrough in engine and aerodynamic technology.2

During World War II, the NACA produced new ways to fix problems 

discovered in higher-powered piston engines being mass-produced for 

wartime bombers. NACA research into centrifugal superchargers was 

particularly useful, especially on the R-1820 Cyclone engines intended 

for use on the Boeing B-17 Flying Fortress, and later with the Wright 

R-3350 Duplex Cyclone engines that powered the B-29.

Basic research on aircraft engine noise was conducted by NACA 

engineers, who reported their findings in a paper presented in 1956 to 

the 51st Meeting of the Acoustical Society of America in Cambridge, 

MA. It would seem that measurements backed up the prediction that 

the noise level of the spinning propeller depended on several variables, 

2. Roger E. Bilstein, Orders of Magnitude: A History of the NACA and NASA, 1915–1990, 

NASA SP-4406 (Washington, DC: NASA, 1989), p. 9.
Case 11 | Advancing Propulsive Technology

737

11

including the propeller diameter, how fast it is turning, and how far 

away the recording device is from the engine.3

As the jet engine made its way from Europe to the United States and 

designs for the basic turboprop, turbojet, and turbofan were refined, the 

NACA during the early 1950s began one of the earliest noise-reduction 

programs, installing multitube nozzles of increasing complexity at the 

back of the engines to, in effect, act as mufflers. These engines were tested 

in a wind tunnel at Langley Research Center in Hampton, VA. But the 

effort was not effective enough to prevent a growing public sentiment 

that commercial jet airliners should be seen and not heard.

In fact, a 1952 Presidential commission chaired by the legendary pilot 

James H. Doolittle predicted that aircraft noise would soon turn into a 

problem for airport managers and planners. The NACA’s response was 

to form a Special Subcommittee on Aircraft Noise and pursue a threepart program to understand better what makes a jet noisy, how to quiet it, 

and what, if any, impact the noise might have on the aircraft’s structure.4

As the NACA on September 30, 1958, turned overnight into the 

National Aeronautics and Space Administration on October 1, the new 

space agency soon found itself with more work to do than just beating 

the Soviet Union to the Moon.

Noise Pollution Forces Engine Improvements

Fast-forward a few years, to a time when Americans embraced the promise that technology would solve the world’s problems, raced the Soviet 

Union to the Moon, and looked forward to owning personal family hovercraft, just like they saw on the TV show The Jetsons. And during that 

same decade of the 1960s, the American public became more and more 

comfortable flying aboard commercial airliners equipped with the modern marvel of turbojet engines. Boeing 707s and McDonnell-Douglas 

DC-8s, each with four engines bolted to their wings, were not only a 

common sight in the skies over major cities, but their presence could 

also easily be heard by anyone living next to or near where the planes 

took off and landed. Boeing 727s and 737s soon followed. At the same 

3. Edward M. Kerwin, Jr., “Procedures for Estimating the Near Field Noise of Rotating Aircraft 

Propellers,” presented at the Fifty-First Meeting of the Acoustical Society of America, Cambridge, 

MA, June 17–23, 1956.

4. J.H. Doolittle, The Airport and Its Neighbors, The Report of the President’s Airport Commission

(Washington, DC: U.S. Government Printing Office, 1952), p. 45.
NASA’s Contributions to Aeronautics

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11

A jet engine is prepared for a test in 1967 as part of an early noise research program at Lewis 

Research Center. NASA.

time that commercial aviation exploded, people moved away from the 

metropolis to embrace the suburban lifestyle. Neighborhoods began 

to spring up immediately adjacent to airports that originally were 

built far from the city, and the new neighbors didn’t like the sound 

of what they hearing.5

5. Alain Depitre, “Aircraft Noise Certification History/Development,” presented at the ICAO Noise 

Certification Workshop, Montreal, 2004, p. 3.
Case 11 | Advancing Propulsive Technology

739

11

By 1966, the problem of aircraft noise pollution had grown to the 

point of attracting the attention of President Lyndon Johnson, who 

then directed the U.S. Office of Science and Technology to set a new 

national policy that said:

The FAA and/or NASA, using qualified contractors as necessary, (should) establish and fund . . . an urgent program for 

conducting the physical, psycho-acoustical, sociological, and 

other research results needed to provide the basis for quantitative noise evaluation techniques which can be used . . . for 

hardware and operational specifications.6

As a result, NASA began dedicating resources to aggressively 

address aircraft noise and sought to contract much of the work to 

industry, with the goals of advancing technology and conducting 

research to provide lawmakers with the information they needed to 

make informed regulatory decisions.7

During 1968, the Federal Aviation Administration (FAA) was given 

authority to implement aircraft noise standards for the airline industry. Within a year, the new standards were adopted and called for all 

new designs of subsonic jet aircraft to meet certain criteria. Aircraft 

that met these standards were called Stage 2 aircraft, while the older 

planes that did not meet the standards were called Stage 1 aircraft. 

Stage 1 aircraft over 75,000 pounds were banned from flying to or from 

U.S. airports as of January 1, 1985. The cycle repeated itself with the 

establishment of Stage 3 aircraft in 1977, with Stage 2 aircraft needing to be phased out by the end of 1999. (Some of the Stage 2 aircraft 

engines were modified to meet Stage 3 aircraft standards.) In 2005, 

the FAA adopted an even stricter noise standard, which is Stage 4. 

All new aircraft designs submitted to the FAA on or after July 5, 2005, 

must meet Stage 4 requirements. As of this writing, there is no timetable for the mandatory phaseout of Stage 3 aircraft.8

6. Alleviation of Jet Aircraft Noise Near Airports (Washington, DC: U.S. Office of Science and 

Technology, 1966), p. 8.

7. Newell D. Sanders, Aircraft Engine Noise Reduction, NASA SP-311 (Washington, DC: NASA, 

1972), p. 2.

8. David M. Bearden, Noise Abatement and Control: An Overview of Federal Standards and 

Regulations (Washington, DC: Congressional Research Service, 2006), p. 3.
NASA’s Contributions to Aeronautics

740

11

With every new set of regulations, the airline industry required 

upgrades to its jet engines, if not wholesale new designs. So having 

already helped establish reliable working versions of each of the major 

types of jet engines—i.e., turboprop, turbojet, and turbofan—NASA 

and its industry partners began what has turned out to be a continuing 

50-year-long challenge to constantly improve the design of jet engines 

to prolong their life, make them more fuel efficient, and reduce their 

environmental impact in terms of air and noise pollution. With this new 

direction, NASA set in motion three initial programs.9

NASA’s first major new program was the Acoustically Treated Nacelle 

program, managed by the Langley Research Center. Engines flying on 

Douglas DC-8 and Boeing 707 aircraft were outfitted with experimental mufflers, which reduced noise during approach and landing but had 

negligible effect on noise pollution during takeoff, according to program 

results reported during a 1969 conference at Langley.10

The second was the Quiet Engine program, which was managed 

by the Lewis Research Center in Cleveland (Lewis became the Glenn 

Research Center on March 1, 1999). Attention here focused on the interior design of turbojet and turbofan engines to make them quieter by 

as much as 20 decibels. General Electric (GE) was the key industry 

partner in this program, which showed that noise reduction was possible by several methods, including changing the rotational speed of the 

fan, increasing the fan bypass ratio, and adjusting the spacing of rotating and stationary parts.11

The third was the Steep Approach program, which was jointly 

managed by Langley and the Ames Research Center/Dryden Flight 

Research Facility, both in California. This program did not result in 

new engine technology but instead focused on minimizing noise on the 

ground by developing techniques for pilots to use in flying steeper and 

faster approaches to airports.12

9. U.S. Government Support of the U.S. Commercial Aircraft Industry, Prepared for the Commission 

of the European Communities (Washington, DC: Arnold and Porter, 1991), pp. 37–43.

10. Sanders, Aircraft Engine Noise Reduction, NASA SP-311 (Washington, DC: NASA, 1972),

p. 2.

11. M.J. Benzakein, S.B. Kazin, and F. Montegani, “NASA/GE Quiet Engine ‘A,’” AIAA Paper 

72-657 (1972).

12. Vicki L. Golich and Thomas E. Pinelli, Knowledge Diffusion in the U.S. Aerospace Industry

(London: Alex Publishing, 1998), p. 61.
Case 11 | Advancing Propulsive Technology

741

11

Quiet Clean Short Haul Experimental Engine

A second wave of engine-improvement programs was initiated in 1969 

and continued throughout the 1970s, as the noise around airports continued to be a social and political issue and the FAA tightened its environmental regulations. Moreover, with the oil crisis and energy shortage later 

in the decade adding to the forces requiring change, the airline industry once again turned to NASA for help in identifying new technology.

At the same time, the airline industry was studying the feasibility 

of introducing a new generation of commuter airliners to fly between 

cities along the Northeast corridor of the United States. To make these 

routes attractive to potential passengers, new airports would have to 

be built as close to the center of cities such as Boston, New York, and 

Philadelphia. For aircraft to fly into such airports, which would have 

shorter runways and strict noise requirements, the airliners would have 

to be capable of making steep climbs after takeoff, quick turns without 

losing control, and steep descents on approach to landing, accommodating short runways and meeting the standards for Stage 2 noise levels.13

In terms of advancing propulsion technology, NASA’s answer to all 

of these requirements was the Quiet Clean Short Haul Experimental 

Engine. Contracts were awarded to GE to design, build, and test two 

types of high-bypass fanjet engines: an over-the-wing engine and an 

under-the-wing engine. Self-descriptive as to their place on the airplane, 

both turbofans were based on the same engine core used in the military 

F-101 fighter jet. Improvements to the design included noise-reduction 

features evolved from the Quiet Engine program; a drive-reduction gear 

to make the fan spin slower than the central shaft; a low-pressure turbine; advanced composite construction for the inlet, fan frame, and fan 

exhaust duct; and a new digital control system that allowed flight computers to monitor and control the jet engine’s operation with more precision and quicker response than a pilot could.14

In addition to those “standard” features on each engine, the underthe-wing engine tried out a variable pitch composite low-pressure fan 

with a 12 to 1 ratio—both features were thought to be valuable in reducing noise, although the variable pitch proved challenging for the GE 

13. Robert V. Garvin, “Starting Something Big: The Commercial Emergence of GE Aircraft Engines,” 

AIAA Paper 72-657 (1999), pp. 162–165.

14. A.P. Adamson, “Quiet Clean Short-Haul Experimental Engine (QCSEE) Design Rationale,” SAE 

Paper 750605 (1975).
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The giant General Electric GE90 jet engine that powers the Boeing 777 benefited from the 

Energy Efficient Engine project. General Electric.

team leading the research. Two pitch change mechanisms were tested, 

one by GE and the other by Hamilton Standard. Both worked well in 

controlled test conditions but would need a lot of work before they could 

go into production.15

The over-the-wing engine incorporated a higher fan pressure and 

a 10 to 1 bypass ratio, a fixed pitch fan, a variable area D-shaped fan 

exhaust nozzle, and low tip speeds on the fans. Both engines directed 

their exhaust along the surface of the wing, which required modifications to handle the hot gas and increase lift performance.16

The under-the-wing engine was test-fired for 153 hours before it was 

delivered to NASA in August of 1978, while the over-the-wing engine 

received 58 hours of testing and was received by NASA during July of 

1977. Results of the tests proved that the technology was sound and, when 

configured to generate 40,000 pounds of thrust, showed a reduction in 

15. Garvin, “Starting Something Big,” pp. 162–165.

16. C.C. Ciepluch, “A Review of the QCSEE Program,” NASA TM-X-71818 (1975).
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noise of 8 to 12 decibels, or about 60- to 75-percent quieter than the 

quietest engines flying on commercial airliners at that time. The new 

technologies also resulted in sharp reductions in emissions of carbon 

monoxide and unburned hydrocarbons.17

Unfortunately, the new generation of Short Take-Off and Landing 

(STOL) commuter airliners and small airports near city centers never 

materialized, so the new engine technology research managed and paid 

for by NASA but conducted mostly by its industry partners never found 

a direct commercial application. But there were many valuable lessons 

learned about the large-diameter turbofans and their nacelles, information that was put to good use by GE years later in the design and fabrication of the GE90 engine that powers the Boeing 777 aircraft.18

Aircraft Energy Efficiency Program

Approved in 1975 and begun in 1976, the Aircraft Energy Efficiency (ACEE) 

program was managed by NASA and funded through 1983, as yet another 

round of research and development activities were put in work to improve 

the state of the art of aircraft structural and propulsion design. And once 

again, the program was aimed at pushing the technological envelope to see 

what might be possible. Then, based on that information, new Government 

regulations could be enacted, and the airline industry could decide if the 

improvements would offer a good return on its investment. The answer, 

as it turned out, was an enthusiastic yes, as the overall results of the program led directly to the introduction of the Boeing 757 and 767.19

Driving this particular program was the rapid increase in fuel costs 

since 1973 and the accompanying energy crisis, which was brought on 

by the Organization of Arab Petroleum Exporting Countries’ decision 

to embargo all shipments of oil to the United States. This action began 

in October 1973 and continued to March 1974. As a result of this and 

other economic influences, the airlines saw their fuel prices as a percentage of direct operating costs rise from 25 percent to as high as 50 

percent within a few weeks. With the U.S. still vulnerable to a future oil 

embargo, along with general concerns about an energy shortage, the 

17. Ciepluch and W.S. Willis, “QCSEE—The Key to Future Short-Haul Air Transport,” ICAO Bulletin 34 (1979).

18. “A Giant Step in Jetliner Propulsion,” Spinoff 1996 (Washington, DC: NASA, 1996), pp. 

56–57.

19. U.S. Government Support of the U.S. Commercial Aircraft Industry (1991).
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Federal Government reacted by ordering NASA to lead an effort to help 

find ways for airlines to become more profitable. Six projects were initiated under the ACEE program, three of which had to do with the aircraft structure and three of which involved advancing engine technology. 

The aircraft projects included Composite Structures, Energy Efficient 

Transport, and Laminar Flow Control. The propulsion technology projects included Engine Component Improvement, Energy Efficient Engine, 

and Advanced Turboprop—all three of which are detailed next.20

Engine Component Improvement Project

The Engine Component Improvement project was tasked with enhancing performance and lowering fuel consumption of several existing commercial aircraft jet engines, in particular Pratt & Whitney’s JT8D and 

JT9D engines and GE’s CF6. The specific goals included:

• 

• 

• 

Improving the current versions of the engines without 

requiring a brand-new design or engine replacement.

Reducing the amount of fuel a typical jet engine would 

use on any given flight by 5 to 6 percent.

Significantly slowing the pace at which the engine’s 

components would naturally degrade and cause a loss 

of performance over time. 

To do this, researchers tried and tested several ideas, including reducing the clearance between rotating parts, lowering the amount of cooling air that is passed through the engine, and making refinements to the 

aerodynamic design of certain engine parts to raise their efficiency. All 

together, engineers identified 16 concepts to incorporate into the engines.21

Ultimately, as a result of the Engine Component Improvement efforts, 

engine parts were incorporated that could resist erosion and warping, 

better seals were introduced, an improved compressor design was used, 

and ceramic coatings were added to the gas turbine blades to increase 

their performance. Tests of the improvements were so promising that 

many were put into production before the program ended, benefiting the 

20. Peter G. Batterton, “Energy Efficient Engine Program Contributions to Aircraft Fuel Conservation,” 

NASA TM-83741 (1984).

21. Louis J. Williams, Small Transport Aircraft Technology (Honolulu: University Press of the Pacific, 

2001), pp. 37–39.
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The classic Pratt & Whitney JT9D engine interior and its major components: the fan, compressor, combustion chamber, turbine, and nozzle. Pratt & Whitney.

workhorse airliners at the time, namely the McDonnell-Douglas DC-9 

and DC-10, as well as the Boeing 727, 737, and 747.22

Energy Efficient Engine Project

Taking everything learned to date by NASA and the industry about making turbo machinery more fuel efficient, the Energy Efficient Engine (E 

Cubed) project sought to further reduce the airlines’ fuel usage and its effect 

on direct operating costs, while also meeting future FAA regulations and 

Environmental Protection Agency exhaust emission standards for turbofan engines. Research contracts were awarded to GE and Pratt & Whitney, 

which initially focused on the CF6-50C and JT9D-7A engines, respectively. 

The program ran from 1975 to 1983 and cost NASA about $200 million.23

Similar to the goals for the Engine Component Improvement project, the E Cubed goals included a 12-percent reduction in specific fuel 

consumption (SFC), which is a measure of the ratio between the mass 

of fuel used to the output power of the jet engine—much like a miles 

per gallon measurement for automobiles. Other goals of the E Cubed 

effort included a 5-percent reduction in direct operating costs and a 

22. U.S. Government Support of the U.S. Commercial Aircraft Industry.

23. Lawrence. E. Macioce, John W. Schaefer, and Neal T. Saunders, “The Energy Efficient Engine 

Project,” NASA TM-81566 (1980).
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A high-pressure 14 to 1 ratio compressor rotor for a prototype Energy Efficient Engine on 

display in 1984 at Lewis Research Center. NASA.

50-percent reduction in the rate at which the SFC worsens over time 

as the engine ages. In addition to making these immediate improvements, it was hoped that a new generation of fuel-conservative turbofan 

engines could be developed from this work.24

Highlighting that program was development of a new type of compressor core and an advanced combustor made up of a doughnut-shaped 

ring with two zones—or domes—of combustion. During times when 

low power is needed or the engine is idling, only one of the two zones 

is lit up. For higher thrust levels, including full power, both domes 

are ignited. By creating a dual combustion option, the amount of fuel 

being burned can be more carefully controlled, reducing emissions of 

smoke, carbon monoxide, and hydrocarbons by 50 percent, and nitrogen oxides by 35 percent.25

As part of the development of the new compressor in particular, and 

the E Cubed and Engine Component Improvement programs in general, the Lewis Research Center developed first-generation computer programs for use in creating the new engine. The software helped engineers 

24. Saunders, “Advanced Component Technologies for Energy-Efficient Turbofan Engines,” NASA 

TM-81507 (1980).

25. Guy Norris and Mark Wagner, Boeing 777: The Technological Marvel (Osceola, WI: MBI 

Publishing, 2001).
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with conceptualizing the aerodynamic design and visualizing the flow of 

gases through the engine. The computer programs were credited with 

making it possible to design more fuel-efficient compressors with less tip 

and end-wall pressure losses, higher operating pressure ratios, and the 

ability to use fewer blades. The compressors also helped to reduce performance deterioration, surface erosion, and damage from bird strikes.26

History has judged the E Cubed program as being highly successful, 

in that the technology developed from the effort was so promising—and 

proved to meet the objectives for reducing emissions and increasing fuel 

efficiency—that both major U.S. jet engine manufacturers, GE and Pratt 

& Whitney, moved quickly to incorporate the technology into their products. The ultimate legacy of the E Cubed program is found today in the 

GE90 engine, which powers the Boeing 777. The E Cubed technology 

is directly responsible for the engine’s economical fuel burn, reduced 

emissions, and low maintenance cost.27

Advanced Turboprop Project—Yesterday and Today

The third engine-related effort to design a more fuel-efficient powerplant 

during this era did not focus on another idea for a turbojet configuration. Instead, engineers chose to study the feasibility of reintroducing 

a jet-powered propeller to commercial airliners. An initial run of the 

numbers suggested that such an advanced turboprop promised the largest reduction in fuel cost, perhaps by as much as 20 to 30 percent over 

turbofan engines powering aircraft with a similar performance. This 

compared with the goal of a 5-percent increase in fuel efficiency for 

the Engine Component Improvement program and a 10- to 15-percent 

increase in fuel efficiency for the E Cubed program.28

But the implementation of an advanced turboprop was one of NASA’s 

more challenging projects, both in terms of its engineering and in securing public acceptance. For years, the flying public had been conditioned 

to see the fanjet engine as the epitome of aeronautical advancement. 

Now they had to be “retrained” to accept the notion that a turbopropeller 

engine could be every bit as advanced, indeed, even more advanced, than 

the conventional fanjet engine. The idea was to have a jet engine firing 

26. NASA Glenn Research Center at Lewis Field: Achieving the Extraordinary (Cleveland: NASA, 

1999), p. 24.

27. “A Giant Step in Jetliner Propulsion,” Spinoff 1996 (Washington, DC: NASA, 1996).

28. Williams, Small Transport Aircraft Technology, p. 38.
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as usual with air being compressed and ignited with fuel and the exhaust 

expelled after first passing through a turbine. But instead of the turbine 

spinning a shaft that turned a fan at the front of the engine, the turbines 

would be spinning a shaft, which fed into a gearbox that turned another 

shaft that spun a series of unusually shaped propeller blades exterior 

to the engine casing.29

Begun in 1976, the project soon grew into one of the larger NASA 

aeronautics endeavors in the history of the Agency to that point, eventually involving 4 NASA Field Centers, 15 university grants, and more than 

40 industrial contracts.30

Early on in the program, it was recognized that the major areas of 

concern were going to be the efficiency of the propeller at cruise speeds, 

noise both on the ground and within the passenger cabin, the effect of 

the engine on the aerodynamics of the aircraft, and maintenance costs. 

Meeting those challenges were helped once again by the computer-aided, 

three-dimensional design programs created by the Lewis Research Center. 

An original look for an aircraft propeller was devised that changed the 

blade’s sweep, twist, and thickness, giving the propellers the look of a 

series of scimitar-shaped swords sticking out of the jet engine. After much 

development and testing, the NASA-led team eventually found a solution 

to the design challenge and came up with a propeller shape and engine 

configuration that was promising in terms of meeting the fuel-efficiency 

goals and reduced noise by as much as 65 decibels.31

In fact, by 1987, the new design was awarded a patent, and the 

NASA–industry group was awarded the coveted Collier Trophy for creating a new fuel-efficient turboprop propulsion system. Unfortunately, two 

unexpected variables came into play that stymied efforts to put the design 

into production.32

The first had to do with the public’s resistance to the idea of flying 

in an airliner powered by propellers—even though the blades were still 

29. Roy D. Hager and Deborah Vrabel, Advanced Turboprop Project, NASA SP-495 (Washington, 

DC: NASA, 1988), p. 5.

30. Mark D. Bowles and Virginia P. Dawson, “The Advanced Turboprop Project: Radical Innovation 

in a Conservative Environment,” in From Engineering Science to Big Science, The NACA and 

NASA Collier Trophy Research Project Winners, NASA SP-4219 (Washington, DC: NASA, 1998), 

p. 323.

31. Glenn A. Mitchell, “Experimental Aerodynamic Performance of Advanced 40 Degree-Swept, 

10-Blade Propeller Model at Mach 0.6 to 0.85,” NASA TM-88969 (1988).

32.Bowles and Dawson, “The Advanced Turboprop Project,” p. 323.
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A General Electric design for an Unducted Fan engine is tested during the early 1980s. 

General Electric.

being turned by a jet engine. It didn’t matter that a standard turbofan 

jet also derived most of its thrust from a series of blades—which did, in 

fact, look more like a fan than a series of propellers. Surveys showed 

passengers had safety concerns about an exposed blade letting go and 

sending shrapnel into the cabin, right where they were sitting. Many 

passengers also believed an airliner equipped with an advanced turboprop was not as modern or reliable as pure turbojet engine. Jets were 

in; propellers were old fashioned. The second thing that happened was 

that world fuel prices dropped to the lower levels that preceded the oil 

embargo and the very rationale for developing the new turboprop in the 

first place. While fuel-efficient jet engines were still needed, the “extra 

mile” in fuel efficiency the advanced turboprop provided was no longer required. As a result, NASA and its partners shelved the technology 

and waited to use the archived files another day.33

33. Ibid.
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The story of the Advanced Turboprop project had one more twist to 

it. While NASA and its team of contractor engineers were working on 

their new turboprop design, engineers at GE were quietly working on 

their own design, initially without NASA’s knowledge. NASA’s engine was 

distinguished by the fact that it had one row of blades, while GE’s version featured two rows of counter-rotating blades. GE’s design, which 

became known as the Unducted Fan (UDF), was unveiled in 1983 and 

demonstrated at the 1985 Paris Air Show. A summary of the UDF’s technical features is described in a GE-produced report about the program:

The engine system consists of a modified F404 gas generator 

engine and counterrotating propulsor system, mechanically 

decoupled, and aerodynamically integrated through a mixing 

frame structure. Utilization of the existing F404 engine minimized engine hardware, cost, and timing requirements and 

provided an engine within the desired thrust class. The power 

turbine provides direct conversion of the gas generator horsepower into propulsive thrust without the requirement for a 

gearbox and associated hardware. Counterrotation utilizes the 

full propulsive efficiency by recovering the exit swirl between 

blade stages and converting it into thrust.34

Although shelved during the late 1980s, the Alternate Turboprop and 

UDF technology and concept is being explored again as part of programs 

such as the Ultra-High Bypass Turbofan and Pratt & Whitney’s Geared 

Turbofan. Neither engine is routinely flying yet on commercial airliners. But both concepts promise further reductions in noise, increases in 

fuel efficiency, and lower operating costs for the airline—goals the aerospace community is constantly working to improve upon.

Several concepts are under study for an Ultra-High Bypass Turbofan, 

including a modernized version of the Advanced Turboprop that takes 

advantage of lessons learned from GE’s UDF effort. NASA has teamed 

with GE to start testing an open-rotor engine. For the NASA tests 

at Glenn Research Center, GE will run two rows of counter-rotating 

fan blades, with 12 blades in the front row and 10 blades in the back 

row. The composite fan blades are one-fifth subscale in size. Tests in 

34. “Full Scale Technology Demonstration of a Modern Counterrotating Unducted Fan Engine 

Concept: Design Test,” NASA CR-180867 (1987).
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a low-speed wind tunnel will simulate low-altitude aircraft speeds for 

acoustic evaluation, while tests in a high-speed wind tunnel will 

simulate high-altitude cruise conditions in order to evaluate blade 

efficiency and performance.35

“The tests mark a new journey for GE and NASA in the world of open 

rotor technology. These tests will help to tell us how confident we are 

in meeting the technical challenges of an open-rotor architecture. It’s a 

journey driven by a need to sharply reduce fuel consumption in future 

aircraft,” David Joyce, president of GE Aviation, said in a statement.36

In an Ultra-High Bypass Turbofan, the amount of air going through 

the engine casing but not through the core compressor and combustion 

chamber is at least 10 times greater than the air going through the core. 

Such engines promise to be quieter, but there can be tradeoffs. For example, an Ultra-High Bypass Engine might have to operate at a reduced 

thrust or have its fan spin slower. While the engine would meet all the 

goals, it would fly slower, thus making passengers endure longer trips.

In the case of Pratt & Whitney’s Geared Turbofan engine, the idea is 

to have an Ultra-High Bypass Ratio engine, yet spin the fan slower (to 

reduce noise and improve engine efficiency) than the core compressor 

blades and turbines, all of which traditionally spin at the same speed, as 

they are connected to the same central shaft. Pratt & Whitney designed 

a gearbox into the engine to allow for the central shaft to turn at one 

speed yet turn a second shaft connected to the fan at another speed.37

Alan H. Epstein, a Pratt & Whitney vice president, testifying before 

the House Subcommittee on Transportation and Infrastructure in 2007, 

explained the potential benefits the company’s Geared Turbofan might 

bring to the aviation industry:

The Geared Turbofan engine promises a new level of very low 

noise while offering the airlines superior economics and environmental performance. For aircraft of 70 to 150 passenger 

size, the Geared Turbofan engine reduces the fuel burned, 

35. Deb Case and Rick Kennedy, “GE and NASA To Begin Wind-Tunnel Testing This Summer of Open 

Rotor Jet Engine Systems,” GE Aviation News Release (Evendale, OH: General Electric, 2009).

36. Ibid.

37. Jeff Schweitzer, “An Overview of Recent Collaboration Research with NASA in Ultra High Bypass Technology,” presented at the NASA Fundamental Aeronautics 2007 Annual Meeting, New 

Orleans, Oct. 30–Nov. 1, 2007.
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and thus the CO2 produced, by more than 12% compared to 

today’s aircraft, while reducing cumulative noise levels about 

20dB below the current Stage 4 regulations. This noise level, 

which is about half the level of today’s engines, is the equivalent difference between standing near a garbage disposal running and listening to the sound of my voice right now.38

Pratt & Whitney’s PW1000G engine incorporating a geared turbofan is selected to be used on the Bombardier CSeries and Mitsubishi 

Regional Jet airliners beginning in 2013. The engine was first flight-tested 

in 2008, using an Airbus A340-600 airliner out of Toulouse, France.39

Digital Electronic Engine Controls

As one set of NASA and contractor engineers worked on improving the 

design of the various types of jet engines, another set of researchers representing another science discipline were increasingly interested in marrying the computer’s capabilities to the operation of a jet engine, much 

in the same way that fly-by-wire systems already were in use with aircraft flight controls.

Beginning with that first Wright Flyer in 1903, flying an airplane 

meant moving levers and other mechanical contrivances that were 

directly connected by wires and cables to control the operation of the 

rudder, elevator, wing surfaces, instruments, and engine. When Chuck 

Yeager broke the sound barrier in 1947 in the X-1, if he wanted to go 

up, he pulled back on the yoke and cables directly connecting the stick 

to the elevator, which made that aerosurface move to effect a change 

in the aircraft’s attitude. The rockets propelling the X-1 were activated 

with a switch throw that closed an electrical circuit whose wiring led 

directly from the cockpit to the engines. As planes grew bigger, so did 

their control surfaces. Aircraft such as the B-52 bomber had aerosurfaces as big as the entire wings of smaller airplanes—too bulky and heavy 

for a single pilot to move using a simple cable/pulley system. A hydraulic system was required and “inserted” between the pilot’s input on the 

yoke and the control surface needing to be moved. Meanwhile, engine 

38. Alan H. Epstein, Statement Before the Subcommittee on Aviation Committee on Transportation 

and Infrastructure, U.S. House of Representatives Hearing on Aviation and Environment: Noise, 

Washington, DC, Oct. 24, 2007.

39. “Pratt & Whitney Pure Power PW1000G Engines,” Pratt & Whitney S16154.9.08 (2008).
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operation remained more or less “old fashioned,” with all parameters 

such as fuel flow and engine temperatures reported to the cockpit on 

dials the pilot could read, react to, and then make changes by adjusting the throttle or other engine controls.

With the introduction of digital computers and the miniaturization of their circuits—a necessity inspired, in part, by the reduced mass 

requirements of space flight—engineers began to consider how the 

quick-thinking electronic marvels might ease the workload for pilots 

flying increasingly more complex aircraft designs. In fact, as the 1960s 

transitioned to the 1970s, engineers were already considering aircraft 

designs that could do remarkable maneuvers in the sky but were inherently unstable, requiring constant, subtle adjustments to the flight controls to keep the vehicle in the air. The solution—already demonstrated 

for spacecraft applications during Project Apollo—was to insert the 

power of the computer between the cockpit controls and the flight control surfaces—a concept known as fly-by-wire. A pilot using this system 

and wanting to turn left would move the control stick to the left, apply 

a little back pressure, and depress the left rudder pedal. Instead of a 

wire/cable system directly moving the related aerosurfaces, the movement of the controls would be sensed by a computer, which would send 

electronic impulses to the appropriate actuators, which in turn would 

deflect the ailerons, elevator, and rudder.40

Managed by NASA’s Dryden Flight Research Facility, the fly-by-wire 

system was first tested without a backup mechanical system in 1972, 

when a modified F-8C fighter took off from Edwards Air Force Base in 

California. Testing on this aircraft, whose aerodynamics were known 

and considered stable, proved that fly-by-wire could work and be reliable. 

In the years to follow, the system was used to allow pilots to safely fly 

unstable aircraft, including the B-2 bomber, the forward-swept winged 

X-29, the Space Shuttle orbiter, and commercial airliners such as the 

Airbus A320 and Boeing 777.41

As experienced was gained with the digital flight control system and 

computers shrunk in size and grew in power, it didn’t take long for propulsion experts to start thinking about how computers could monitor 

40. C.R. Jarvis, “An Overview of NASA’s Digital Fly-By-Wire Technology Development Program,” 

NASA 75N18246 (1975).

41. James E. Tomayko, Computers Take Flight: A History of NASA’s Pioneering Digital Fly-By-Wire 

Project, NASA SP-4224 (Washington, DC: NASA, 2000), p. vii.
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engine performance and, by making many adjustments in every variable that affects the efficiency of a jet engine, improve the powerplant’s 

overall capabilities.

The first step toward enabling computer control of engine operations 

was taken by Dryden engineers in managing the Integrated Propulsion 

Control System (IPCS) program during the mid-1970s. A joint effort 

with the U.S. Air Force, the IPCS was installed on an F-111E longrange tactical fighter-bomber aircraft. The jet was powered by twin TF30 

afterburning turbofan engines with variable-geometry external compression inlets. The IPCS effort installed a digital computer to control 

the variable inlet and realized significant performance improvements 

in stallfree operations, faster throttle response, increased thrust, and 

improved range flying at supersonic speeds. During this same period, 

results from the IPCS tests were applied to NASA’s YF-12C Blackbird, 

a civilian research version of the famous SR-71 Blackbird spy plane. A 

digital control system installed on the YF-12C successfully tested, monitored, and adjusted the engine inlet control, autothrottle, air data, and 

navigation functions for the Pratt & Whitney-built engines. The results 

gave the aircraft a 7-percent increase in range, improved handling characteristics, and lowered the frequency of inlet unstarts, which happen 

when an engine shock wave moves forward of the inlet and disrupts the 

flow of air into the engine, causing it to shutdown. Seeing how well this 

computer-controlled engine worked, Pratt & Whitney and the U.S. Air 

Force in 1983 chose to incorporate the system into their SR-71 fleet.42

The promising future for more efficient jet engines from developing digitally controlled integrated systems prompted Pratt & Whitney, 

the Air Force, and NASA (involving both Dryden and Lewis) to pursue a more robust system, which became the Digital Electronic Engine 

Control (DEEC) program.

Pratt & Whitney actually started what would become the DEEC program, using its own research and development funds to pay for configuration studies beginning during 1973. Then, in 1978, Lewis engineers tested 

a breadboard version of a computer-controlled system on an engine in 

an altitude chamber. By 1979, the Air Force had approached NASA and 

asked if Dryden could demonstrate and evaluate a DEEC system using 

an F100 engine installed in a NASA F-15, with flight tests beginning in 

42. James F. Stewart, Frank W. Burcham, Jr., and Donald H. Gatlin, “Flight-Determined Benefits of 

Integrated Flight-Propulsion Control Systems,” NASA TM-4393 (1992), pp. 2–4.
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The Digital Electronic Engine Control system was tested on a Pratt & Whitney F100 turbojet, similar to the one shown here undergoing a hot fire on a test stand. Pratt & Whitney.

1981. At every step in the test program, researchers took advantage of lessons learned not only from the IPCS exercise but also from a U.S. Navyfunded effort called the Full Authority Digital Engine Control program, 

which ran concurrently to the IPCS program during the mid-1970s.43

A NASA Dryden fact sheet about the control system does a good job 

of explaining in a concise manner the hardware involved, what it monitored, and the resulting actions it was capable of performing:

The DEEC system tested on the NASA F-15 was an engine 

mounted, fuel-cooled, single-channel digital controller that 

received inputs from the airframe and engine to control a wide 

range of engine functions, such as inlet guide vanes, compressor stators, bleeds, main burner fuel flow, afterburner fuel flow 

and exhaust nozzle vanes.

43. T.W. Putnam, “Digital Electronic Engine Control History,” NASA 86N25344 (1984), p. 2.
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Engine input measurements that led to these computercontrolled functions included static pressure at the compressor face, fan and core RPM, compressor face temperature, 

burner pressure, turbine inlet temperature, turbine discharge 

pressure, throttle position, afterburner fuel flow, fan and compressor speeds and an ultra violet detector in the afterburner 

to check for flame presence.

Functions carried out after input data were processed by the 

DEEC computer included setting the variable vanes, positioning compressor start bleeds, controlling gas-generator and 

augmentation of fuel flows, adjusting the augmenter segmentsequence valve, and controlling the exhaust nozzle position.

These actions, and others, gave the engine—and the pilot—

rapid and stable throttle response, protection from fan and 

compressor stalls, improved thrust, better performance at 

high altitudes, and they kept the engine operating within its 

limits over the full flight envelope.44

When incorporated into the F100 engine, the DEEC provided improvements such as faster throttle responses, more reliable capability to restart 

an engine in flight, an increase of more than 10,000 feet in altitude when 

firing the afterburners, and the capability of providing stallfree operations. 

And with the engine running more efficiently thanks to the DEEC, overall 

engine and aircraft reliability and maintainability were improved as well.45

So successful and promising was this program that even before testing was complete the Air Force approved widespread production of the 

F100 control units for its F-15 and F-16 fighter fleet. Almost at the same 

time, Pratt & Whitney added the digital control technology in its PW2037 

turbofan engines for the then-new Boeing 757 airliner.46

With the DEEC program fully opening the door to computer control 

of key engine functions, and with the continuing understanding of flyby-wire systems for aircraft control—along with steady improvements 

in making computers faster, more capable, and smaller—the next logi44. “The DEEC,” NASA TF-2004-03-DFRC (2004).

45. “Digital Electronic Engine Control (DEEC) Flight Evaluation in an F-15 Airplane,” NASA CP2298 (1984).

46. Christian Gelzer, “60 Years of Cutting-Edge Flight Research Marked at NASA Dryden,” Dryden 

News Release 06-37 (2006).
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cal step was to combine together computer control of engines and flight 

controls. This was done initially with the Adaptive Engine Control System 

(ADECS) program accomplished between 1985 and 1989, followed by the 

Performance Seeking Control (PSC) program that performed 72 flight tests 

between 1990 and 1993. The PSC system was designed to handle multiple 

variables in performance, compared with the single-variable control allowed 

in ADECS. The PSC effort was designed to optimize the engine and flight controls in four modes: minimum fuel flow at constant thrust, minimum turbine 

temperature at constant thrust, maximum thrust, and minimum thrust.47

The next evolution in the combining of computer-controlled flight 

and engine controls— a legacy of the original DEEC program—was 

inspired in large part by the 1989 crash in Sioux City, IA, of a DC-10 

that had lost all three of its hydraulic systems when there was an uncontained failure of the aircraft’s No. 2 engine. With three pilots in the 

cockpit, no working flight controls, and only the thrust levels available 

for the two remaining working engines, the crew was able to steer the jet 

to the airport by using variable thrust. During the landing, the airliner 

broke apart, killing 111 of the 296 people on board.48

Soon thereafter, Dryden managers established a program to thoroughly investigate the idea of a Propulsion Controlled Aircraft (PCA) 

using variable thrust between engines to maintain safe flight control. 

Once again, the NASA F-15 was pressed into service to demonstrate 

the concept. Beginning in 1991 with a general ability to steer, refinements in the procedures were made and tested, allowing for more precise 

maneuvering. Finally, on April 21, 1993, the flight tests of PCA concluded 

with a successful landing using only engine power to climb, descend, and 

maneuver. Research continued using an MD-11 airliner, which successfully demonstrated the technology in 1995.49

Numerical Propulsion System Simulation

NASA and its contractor colleagues soon found another use for computers 

to help improve engine performance. In fact, looking back at the history 

47. John S. Orne, “Performance Seeking Control Program Overview,” NASA 95N33011 (1995), 

p. 32.

48. “Aircraft Accident Report: United Airlines Flight 232, McDonnell Douglas DC-10-10, Sioux 

Gateway Airport, Sioux City, Iowa, July 19, 1989,” NTSB AAR-90-06 (1989).

49. Tom Tucker, Touchdown: The Development of Propulsion Controlled Aircraft at NASA Dryden

(Washington, DC: NASA, 1999).
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A computer system known as Propulsion Controlled Aircraft is tested aboard an MD-11 airliner 

in 1995 at the Dryden Flight Research Center. NASA.

of NASA’s involvement with improving propulsion technology, a trilogy 

of major categories of advances can be suggested based on the development of the computer and its evolution in the role that electronic thinkers have played in our culture.

Part one of this story includes all the improvements NASA and 

its industry partners have made with jet engines before the computer 

came along. Having arrived at a basic operational design for a turbojet 

engine—and its relations, the turboprop and turbofan—engineers sought 

to improve fuel efficiency, reduce noise, decrease wear, and otherwise 

reduce the cost of maintaining the engines. They did this through such 

efforts as the Quiet Clean Short Haul Experimental Engine and Aircraft 

Energy Efficiency program, detailed earlier in this case study. By tinkering with the individual components and testing the engines on the ground 

and in the air for thousands of hours, incremental advances were made.50

Part two of the story introduces the capabilities made available to 

engineers as computers became powerful enough and small enough to 

be incorporated into the engine design. Instead of requiring the pilot 

to manually make occasional adjustments to the engine operation in 

50. “Propulsion/ACEE,” NASA FACTS-93/8-81 (1981).
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flight depending on what the instruments read, a small digital computer 

built into the engine senses thousands of measurements per minute and 

caused an equal number of adjustments to be made to keep the powerplant performing at peak efficiency. With the Digital Electronic Engine 

Control, engines designed years before behaved as though they were 

fresh off the drawing boards, thanks to their increased capabilities.51

Having taken engine designs about as far as it was thought possible, 

the need for even more fuel-efficient, quieter, and capable engines continued. Unfortunately, the cost of developing a new engine from scratch, 

building it, and testing it in flight can cost millions of dollars and take 

years to accomplish. What the aerospace industry needed was a way 

to take advantage of the powerful computers available at the dawn of 

the 21st century to make the engine development process less expensive and timelier. The result was part three of NASA’s overarching story 

of engine development: the Numerical Propulsion System Simulation 

(NPSS) program.52

Working with the aerospace industry and academia, NASA’s Glenn 

Research Center led the collaborative effort to create the NPSS program, which was funded and operated as part of the High Performance 

Computing and Communications program. The idea was to use modern 

simulation techniques and create a virtual engine and test stand within a 

virtual wind tunnel, where new designs could be tried out, adjustments 

made, and the refinements exercised again without costly and timeconsuming tests in the “real” world. As stated in a 1999 industry review 

of the program, the NPSS was built around inclusion of three main elements: “Engineering models that enable multi-disciplinary analysis of 

large subsystems and systems at various levels of detail, a simulation 

environment that maximizes designer productivity and a cost-effective, 

high-performance computing platform.”53

In explaining to the industry the potential value of the program during a 2006 American Society of Mechanical Engineers conference in 

51. Jennifer L. Baer-Riedhart and Robert J. Landy, “Highly Integrated Digital Electronic Control—Digital 

Flight Control, Aircraft Model Identification and Adaptive Engine Control,” NASA TM-86793 (1987).

52. John K. Lytle, “The Numerical Propulsion System Simulation: A Multidisciplinary Design System 

for Aerospace Vehicles,” NASA TM-1999-209194 (1999), p. 1.

53. John Lytle, Greg Follen, Cynthia Naiman, Austin Evans, Joseph Veres, Karl Owen, and Isaac Lopez, 

“Numerical Propulsion System Simulation (NPSS) 1999 Industry Review,” NASA TM-2000-209795 

(2000), p. 7.
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Spain, a NASA briefer from Glenn suggested that if a standard turbojet development program for the military—such as the F100—took 10 

years, $1.5 billion, construction of 14 ground-test engines, 9 flight-test 

engines, and more than 11,000 hours of engine tests, the NPSS program could realize a:

• 

• 

• 

• 

• 

50-percent reduction in tooling cost.

33-percent reduction in the average development 

engine cost.

30-percent reduction in the cost of fabricating, 

assembling, and testing rig hardware.

36-percent reduction in the number of development 

engines.

60-percent reduction in total hardware cost.54

A key—and groundbreaking—feature of NPSS was its ability to integrate simulated tests of different engine components and features, and 

run them as a whole, fully modeling all aspects of a turbojet’s operation. The program did this through the use of the Common Object 

Request Broker Architecture (CORBA), which essentially provided a 

shared language among the objects and disciplines (mechanical, thermodynamics, structures, gas flow, etc.) being tested so the resulting data 

could be analyzed in an “apples to apples” manner. Through the creation of an NPSS developer’s kit, researchers had tools to customize the 

software for individual needs, share secure data, and distribute the simulations for use on multiple computer operating systems. The kit also 

provided for the use of CORBA to “zoom” in on the data to see specific 

information with higher fidelity.55

Begun in 1997, the NPSS team consisted of propulsion experts and 

software engineers from GE, Pratt & Whitney, Boeing, Honeywell, RollsRoyce, Williams International, Teledyne Ryan Aeronautical, Arnold 

Engineering Development Center, Wright-Patterson AFB, and NASA’s 

54. Ann K. Sehra, “The Numerical Propulsion System Simulation: A Vision for Virtual Engine Testing,” 

presented at the American Society of Mechanical Engineers TURBO EXPO, Barcelona, Spain, May 

8–11, 2006.

55. Cynthia G. Naiman and Gregory J. Follen, “Numerical Propulsion System Simulation—A 

Common Tool for Aerospace Propulsion Being Developed,” Research and Technology Report 2000 

(Cleveland: NASA, 2001).
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Glenn Research Center. By the end of the 2000 fiscal year, the NPSS 

team had released Version 1.0.0 on schedule. According to a summary 

of the program produced that year:

(The new software) can be used as an aero-thermodynamic 

zero-dimensional cycle simulation tool. The capabilities 

include text-based input syntax, a sophisticated solver, steadystate and transient operation, report generation, a built-in 

object-oriented programming language for user-definable 

components and functions, support for distributed running 

of external codes via CORBA, test data reduction, interactive 

debug capability and customer deck generation.56

Additional capabilities were added in 2001, including the ability to 

support development of space transportation technologies. At the same 

time, the initial NPSS software quickly found applications in aviation 

safety, ground-based power, and alternative energy devices, such as fuel 

cells. Moreover, project officials at the time suggested that with the further development of the software, other applications could be found for 

the program in the areas of nuclear power, water treatment, biomedicine, chemical processing, and marine propulsion. NPSS proved to be so 

capable and promising of future applications that NASA designated the 

program a cowinner of the NASA Software of the Year Award for 2001.57

Work to improve the capabilities and expand the applications of the 

software continued, and, in 2008, NASA transferred NPSS to a consortium of industry partners, and, through a Space Act Agreement, it is currently offered commercially by Wolverine Ventures, Inc., of Jupiter, FL. 

Now at Version 1.6.5, NPSS’s features include the ability to model all 

types of complex systems, plug-and-play interfaces for fluid properties, 

built-in plotting package, interface to higher fidelity legacy codes, multiple model views, command language interpreter with language sensitive text editor, comprehensive component solver, and variable setup 

controls. It also can operate on Linux, Windows, and UNIX platforms.58

56. Ibid.

57. Laurel J. Strauber and Cynthia G. Naiman, “Numerical Propulsion System Simulation (NPSS): 

An Award Winning Propulsion System Simulation Tool,” Research and Technology Report 2001 

(Cleveland: NASA, 2002).

58. “NPSS User Guide, Software Release: NPSS 1.6.5,” NASA NPSS-User (2008), pp. 1-1 to 1-2.
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Originally begun as a virtual tool for designing new turbojet 

engines, NPSS has since found uses in testing rocket engines, fuel cells, 

analog controls, combined cycle engines, thermal management systems, airframe vehicles preliminary design, and commercial and 

military engines.59

Ultra Efficient Engine Technology Program

With the NPSS tool firmly in place and some four decades of experience 

incrementally improving the design, operation, and maintenance of 

the jet engine, it was time to go for broke and assemble an ultrabright team of engineers to come up with nothing short of the best jet 

engine possible.

Building on the success of technology development programs 

such as the Quiet Clean Short Haul Experimental Engine and Energy 

Efficient Engine project—all of which led directly to the improvements 

and production of turbojet engines now propelling today’s commercial 

airliners—NASA approached the start of the 21st century with plans 

to take jet engine design to accomplish even more impressive feats. In 

1999, the Aeronautics Directorate of NASA began the Ultra Efficient 

Engine Technology (UEET) program—a 5-year, $300-million effort—

with two primary goals. The first was to find ways that would enable 

further improvements in engine efficiency to reduce fuel burn and, as a 

result, carbon dioxide emissions by yet another 15 percent. The second 

was to continue developing new materials and configuration schemes 

in the engine’s combustor to reduce emissions of nitrogen oxides (NOx) 

during takeoff and landings by 70 percent relative to the standards 

detailed in 1996 by the International Civil Aviation Organization.60

NASA’s Glenn Research Center led the program, with participation 

from three other NASA Centers: Ames, Langley, and the Goddard Space 

Flight Center in Greenbelt, MD. Also involved were GE, Pratt & Whitney, 

Honeywell, Allison/Rolls-Royce, Williams International, Boeing, and 

Lockheed Martin.61

59. Edward J. Hall, Joseph Rasche, Todd A. Simons, and Daniel Hoyniak, “NPSS Multidisciplinary 

Integration and Analysis,” NASA CR-2006-213890 (2006).

60. Joe Shaw, “Ultra-Efficient Engine Technology Project Continued to Contribute to Breakthrough 

Technologies,” Research and Technology Report 2002 (Cleveland: NASA, 2003).

61. Lori A. Manthey, “NASA Glenn Research Center UEET (Ultra-Efficient Engine Technology) 

Program: Agenda and Abstracts,” NASA RTOP-714-01-4A (2001).
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The program was comprised of seven major projects, each of which 

addressed particular technology needs and exploitation opportunities.62

The Propulsion Systems Integration and Assessment project examined 

overall component technology issues relevant to the UEET program to 

help furnish overall program guidance and identify technology shortfalls.63 The Emissions Reduction project sought to significantly reduce 

NOx and other emissions, using new combustor concepts and technologies such as lean burning combustors with advanced controls and 

high-temperature ceramic matrix composite materials.64 The Highly 

Loaded Turbomachinery project sought to design lighter-weight, reducedstage cores, low-pressure spools and propulsors for more efficient and 

environmentally friendly engines, and advanced fan concepts for quieter, lighter, and more efficient fans.65 The Materials and Structures for 

High Performance project sought to develop and demonstrate hightemperature material concepts such as ceramic matrix composite 

combustor liners and turbine vanes, advanced disk alloys, turbine airfoil material systems, high-temperature polymer matrix composites, and 

innovative lightweight materials and structures for static engine structures.66 The Propulsion-Airframe Integration project studied propulsion systems and engine locations that could furnish improved engine 

and environmental benefits without compromising the aerodynamic 

performance of the airplane; lowering aircraft drag itself constituted 

a highly desirable means of reducing fuel burn, and, hence, CO2 emissions will develop advanced technologies to yield lower drag propulsion 

system integration with the airframe for a wide range of vehicle classes. 

Decreasing drag improves air vehicle performance and efficiency, which 

62. Manthey, “Ultra-Efficient Engine Technology (UEET) Program,” Research and Technology Report 

2001 (NASA, 2002).

63. Ronald C. Plybon, Allan VanDeWall, Rajiv Sampath, Mahadevan Balasubramaniam, 

Ramakrishna Mallina, and Rohinton Irani, “High Fidelity System Simulation of Multiple Components 

in Support of the UEET Program,” NASA CR-2006-214230 (2006).

64. Kathleen M. Tacina and Changlie Wey, “NASA Glenn High Pressure Low NOx Emissions 

Research,” NASA TM-2008-214974 (2008).

65. Michael T. Tong and Scott M. Jones, “An Updated Assessment of NASA Ultra-Efficient Engine 

Technologies,” presented at 17th International Symposium on Airbreathing Engines, Munich, Germany, Sept. 4–9, 2006.

66. James A. DiCarlo, Hee Mann Yun, Gregory N. Morscher, and Ramakrishna T. Bhatt, “HighPerformance SiC/SiC Ceramic Composite Systems Developed for 1315 C (2400 F) Engine 

Components,” Research and Technology Report 2003 (Cleveland: NASA, 2004).
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reduces fuel burn to accomplish a particular mission, thereby reducing 

the CO2 emissions.67 The Intelligent Propulsion Controls Project sought to 

capitalize upon breakthroughs in electronic control technology to improve 

propulsion system life and enhance flight safety via integrating information, propulsion, and integrated flight propulsion control technologies.68

Finally, the Integrated Component Technology Demonstrations project 

sought to evaluate the benefits of off-the-shelf propulsion systems integration on NASA, Department of Defense, and aeropropulsion industry 

partnership efforts, including both the UEET and the military’s Integrated 

High Performance Turbine Engine Technology (IHPTET) programs.69

By 2003, the 7 project areas had come up with 10 specific technology areas that UEET would investigate and incorporate into an engine 

that would meet the program’s goals for reducing pollution and increasing fuel burn efficiency. The technology goals included:

1. Advanced low-NOx combustor design that would feature 

a lean burning concept.

2. A highly loaded compressor that would lower system 

weight, improve overall performance, and result in lower 

fuel burn and carbon dioxide emissions.

3. A highly loaded, high-pressure turbine that could allow 

a reduction in the number of high-pressure stages, parts 

count, and cooling requirements, all of which could 

improve fuel burn and lower carbon dioxide emissions.

4. A highly loaded, low-pressure turbine and aggressive transition duct that would use flow control techniques that 

would reduce the number of low-pressure stages within 

the engine.

5. Use of a ceramic matrix composite turbine vane that 

would allow high-pressure vanes to operate at a higher 

67. Cecile M. Burg, Geoffrey A. Hill, Sherilyn A. Brown, and Karl A. Geiselhart, “Propulsion Airframe Aeroacoustics Technology Evaluation and Selection Using a Multi-Attribute Decision Making 

Process and Non-Deterministic Design,” AIAA Paper 2004-4436 (2004).

68. Sanjay Garg, “NASA Glenn Research in Controls and Diagnostics for Intelligent Aerospace 

Propulsion Systems,” presented at the Integrated Condition Management 2006 Conference, Anaheim, CA, Nov. 14–16, 2006.

69. Mary Jo Long-Davis, “Integrated Components Technology Demonstrations Overview,” NTRS 

Document ID 200.502.14062 (2001).
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inlet temperature, which would reduce the amount of 

engine cooling necessary and result in lower carbon dioxide emissions.

6. The same ceramic matrix composite material would be 

used to line the combustor walls so it could operate at a 

higher temperature and reduce NOx emissions.

7. Coat the turbine airfoils with a ceramic thermal barrier 

material to allow the turbines to operate at a higher temperature and thus reduce carbon dioxide emissions.

8. Use advanced materials in the construction of the turbine airfoil and disk. Specifically, use a lightweight single 

crystal superalloy to allow the turbine blades and vanes 

to operate at a higher temperature and reduce carbon 

dioxide emissions, as well as a dual microstructure nickelbase superalloy to manufacture turbine disks tailored to 

meet the demands of the higher-temperature environment.

9. Determine advanced materials and structural concepts 

for an improved, lighter-weight impact damage tolerance 

and noise-reducing fan containment case.

10. Develop active tip clearance control technology for use in 

the fan, compressor, and turbine to improve each component’s efficiency and reduce carbon dioxide emissions.70

In 2003, the UEET program was integrated into NASA’s Vehicle 

Systems program to enable the enginework to be coordinated with 

research into improving other areas of overall aircraft technology. But 

in the wake of policy changes associated with the 2004 decision to redirect NASA’s space program to retire the Space Shuttle and return humans 

to the Moon, the Agency was forced to redirect some of its funding to 

Exploration, forcing the Aeronautics Directorate to give up the $21.6 million budgeted for UEET in fiscal year 2005, effectively canceling the biggest and most complicate jet engine research program ever attempted. 

At the same time, NASA was directed to realign its jet engine research to 

concentrate on further reducing noise.71

70. Michael T. Tong and Scott M. Jones, “An Updated Assessment of NASA Ultra-Efficient Engine 

Technologies,” ISABE-2005-1163 (2005), p. 3.

71. John W. Douglass, “NASA Aeronautics Research Funding: The Wrong Direction,” Space 

News, Mar. 28, 2005, opinion page.
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Nevertheless, results from tests of UEET hardware showed promise that a large, subsonic aircraft equipped with some of the technologies 

detailed above would have a “very high probability” of achieving the program goals laid out for reducing emissions of carbon dioxide and other 

pollutants. The data remain for application to future aircraft and engine 

schemes.72

Damage-Tolerant Fan Casing

While most eyes were on the big picture of making major engine advancements through the years, some very specific problems were addressed 

with programs that are just as interesting to consider as the larger 

research endeavors. The casings that surround the jet engine’s turbomachinery are a case in point.

With the 1989 crash of United Airlines Flight 232 at Sioux City, IA, 

aviation safety officials became more interested in finding new materials 

capable of containing the resulting shrapnel created when a jet engine’s 

blade or other component breaks free. In the case of the DC-10 involved 

in this particular crash, the fan disk of the No. 2 engine—the one located 

in the tail—separated from the engine and caused the powerplant to 

explode, creating a rain of shrapnel that could not be contained within the 

engine casing. The sharp metal fragments pierced the body of the aircraft 

and cut lines in all three of the aircraft’s hydraulic systems. As previously 

mentioned in this case study, the pilots on the DC-10 were able to steer 

their aircraft to a nearly controlled landing. The incident inspired NASA 

pilots to refine the idea of using only jet thrust to maneuver an airplane 

and undertake the Propulsion Controlled Aircraft program, which took 

full advantage of the earlier Digital Electronic Engine Control research. 

The Iowa accident also sent structures and materials experts off on a hunt 

to find a way to prevent accidents like this in the future.

The United Flight 232 example notwithstanding, the challenge for 

structures engineers is to design an engine casing that will contain a 

failed fan blade within the engine so that it has no chance to pierce the 

passenger compartment wall and threaten the safety of passengers or 

cause a catastrophic tear in the aircraft wall. Moreover, not only does 

the casing have to be strong enough to withstand any blade or shrapnel 

impacts, it must not lose its structural integrity during an emergency 

72. Tong and Jones, “An Updated Assessment of NASA Ultra-Efficient Engine Technologies,” p. 1.
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engine shutdown in flight. A damaged engine can take some 15 seconds 

to shut down, during which time cracks from the initial blade impacts 

can propagate in the fan case. Should the fan case totally fail, the resulting breakup of the already compromised turbomachinery could be catastrophic to the aircraft and all aboard.73

As engineers considered the use of composite materials, two methods 

for containing blade damage within the engine casing were now available: the new softwall and the traditional hardwall. In the softwall concept, 

the casing was made of a sandwich-type aluminum structure overwound 

with dry aramid fibers. (Aramid fibers were introduced commercially by 

DuPont during the early 1960s and were known by the trade name Nomex.) 

The design allows broken blades and other shrapnel to pass through the 

“soft” aluminum and be stopped and contained within the aramid fiber 

wrap. In the hardwall approach, the casing is made of aluminum only 

and is built as a rigid wall to reflect blade bits and other collateral damage 

back into the casing interior. Of course that vastly increases the risk that 

the shrapnel will be ingested through the engine and cause even greater 

damage, perhaps catastrophic. While that risk exists with the softwall 

design, it is not as substantial. Another benefit of the hardwall is that it 

maintains its structural soundness, or ductility, during a breakup of an 

engine. A softwall also features some amount of ductility, but the energyabsorbing properties of the aramid fibers is the major draw.74

In 1994, NASA engineers at the Lewis Research Center began looking into better understanding engine fan case structures and conducted 

impact tests as part of the Enabling Propulsion Materials program. 

Various metallic materials and new ideas for lightweight fan containment structures were studied. By 1998, the research expanded to include 

investigations into use of polymer composites for engine fan casings. As 

additional composite materials were made available, NASA researchers 

sought to understand their properties and the appropriateness of those 

materials in terms of containment capability, damage tolerance, commercial viability, and understanding any potential risk not yet identified for their use on jet engines.75

73. C.L. Stotler and A.P. Coppa, “Containment of Composite Fan Blades,” NASA CR-159544 

(1979).

74. Bob Griffiths, “Composite Fan Blade Containment Case: Innovative Use of Carbon-Fiber Braid 

Yields a Ductile Structure that Resists Blade Impact,” High Performance Composites (May 1, 2005).

75. Ibid.
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In 2001, NASA awarded a Small Business Innovation Research 

(SBIR) grant to A&P Technology, Inc., of Cincinnati to develop a 

damage-tolerant fan casing for a jet engine. Long before composites 

came along, the company’s expertise was in braiding materials together, 

such as clotheslines and candlewicks. A&P—working together with the 

FAA, Ohio State University, and the University of Akron—was able to 

rapidly develop a prototype composite fan case that could be compared 

to the metal fan case. Computer simulations were key to the effort and 

serendipitously provided an opportunity to grow the industry’s understanding and ability to use those very same simulation capabilities. First, 

well understood metallic casings undergoing a blade-out scenario were 

modeled, and the computer tested the resulting codes to reproduce the 

already-known results. Then came the trick of introducing code that 

would represent A&P’s composite casing and its reaction to a blade-out 

situation. The process was repeated for a composite material wrapped 

with a braided fiber material, and results were very promising.76

The composite casing proposed by A&P used a triaxial carbon braid, 

which has a toughness superior to aluminum but is lighter, which helps 

ease fuel consumption. In tests of debris impact, the braided laminate 

performed better than the metal casing, because in some cases, the composite structure absorbed the energy of the impact as the debris bounced 

off the wall, and in other cases where the shrapnel penetrated the material, the damage to the wall was isolated to the impact point and did 

not spread. In a metal casing that was pierced, the resulting hole would 

instigate several cracks that would continue to propagate along the casing wall, appearing much like the spiderweb of cracks that appear on an 

automobile windshield when it is hit with a small stone on the freeway.

NASA continues to study the use of composite casings to better 

understand the potential effects of aging and/or degradation following 

the constant temperature, vibration, and pressure cycles a jet engine 

experiences during each flight. There also is interest in studying the 

effects of higher operating temperatures on the casing structure for possible use on future supersonic jets. (The effect of composite fan blades 

on casing containment also has been studied.)77

76. “Damage-Tolerant Fan Casings for Jet Engines,” Spinoff 2006 (Washington, DC: NASA, 

2006), p. 14.

77. C.L. Stotler and A.P. Coppa, “Containment of Composite Fan Blades,” NASA CR-159544 

(1979).
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A General Electric GEnx engine with a composite damage-tolerant fan casing is checked out 

before eventual installation on the new Boeing 787. General Electric.

While composites have found many uses in commercial and military 

aviation, the first use of an all-composite engine casing, provided by A&P, 

is set to be used on GE’s GEnx turbojet designed for the Boeing 787. The 

braided casing weighs 350 pounds less per engine, and, when other engine 

installation hardware to handle the lighter powerplants is considered, the 

787 should weigh 800 pounds less than a similarly equipped airliner using 

aluminum casings. The weight reduction also should provide a savings 

in fuel cost, increased payload, and/or a greater range for the aircraft.78

Conclusion and a Look Ahead

For more than 50 years now, NASA has methodically and, for the most 

part, quietly advanced the state of art of propulsion technology. With 

the basic design of the jet engine unchanged since it was invented during World War II, modern jet engines incorporate every lesson learned 

during NASA’s past five decades of research. As a result, jet engines are 

78. “Damage-Tolerant Fan Casings for Jet Engines.”
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now quieter, safer, more fuel-efficient, less expensive to operate, and less 

polluting, while being easier to maintain. And thanks to advancements 

in computers and simulations, new engines can be tested for thousands 

of hours at a time without ever bending one piece of aluminum or braiding a square yard of composite material.

So what’s in store for propulsion technology during the next few 

decades? More improvements with every possible variable of engine 

operations are still possible, with future advances more closely linked 

to new aircraft designs, such as the blended wing and body in which the 

engines may be more fully integrated into the structure of the aircraft.

In a feature story written in April 2009 for NASA’s Aeronautics 

Research Mission Directorate Web site, this author interviewed several key Agency officials who are considering what the future holds for 

engine development and making plans for what the Agency’s approach 

will be for managing the effort. Here is that look ahead.

NASA Researchers Work to Reduce Noise in Future Aircraft Design

It’s a noisy world out there, especially around the Nation’s busiest airports, so NASA is pioneering new technologies and aircraft designs that 

could help quiet things down a bit. Every source of aircraft noise, from 

takeoff to touchdown, is being studied for ways to reduce the racket, 

which is expected to get worse as officials predict that air traffic will 

double in the next decade or so.

“It’s always too noisy. You have to always work on making it quieter,” 

said Edmane Envia, an aerospace engineer at NASA’s Glenn Research 

Center in Cleveland. “You always have to stay a step ahead to fulfill the 

needs and demands of the next generation of air travel.”79

Noise reduction research is part of a broader effort by NASA’s 

Aeronautics Research Mission Directorate in Washington to lay a technological foundation for a new generation of airplanes that are not as 

noisy, fly farther on less fuel, and may operate out of airports with much 

shorter runways than exist today. There are no clear solutions yet to 

these tough challenges, neither is there a shortage of ideas from NASA 

researchers who are confident positive results eventually will come.80

79. Interview of Envia by Jim Banke, Cape Canaveral, Feb. 4, 2009.

80. Jeffrey J. Berton, Envia, and Casey L. Burley, “An Analytical Assessment of NASA’s N1 Subsonic Fixed Wing Project Noise Goal,” NASA LF99-8609 (2009).
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“Our goal is to have the technologies researched and ready, but ultimately it’s the aircraft industry, driven by the market, that makes the decision when to introduce a particular generation of aircraft,” Envia said.

NASA organized its research to look three generations into the future, 

with conceptual aircraft designs that could be introduced 10, 20, or 30 

years from now. The generations are called N+1, N+2, and N+3. Each 

generation represents a design intended to be flown a decade or so later 

than the one before it and is to feature increasingly sophisticated methods for delivering quieter aircraft and jet engines.81

“Think of the Boeing 787 Dreamliner as N and the N+1 as the next 

generation aircraft after that,” Envia said.

The N+1 is an aircraft with familiar parts, including a conventional 

tube-shaped body, wings, and a tail. Its jet engines still are attached to 

the wings, as with an N aircraft, but those engines might be on top of 

the wings, not underneath. Conceptual N+2 designs throw out convention and basically begin with a blank computer screen, with design 

engineers blending the line between the body, wing, and engines into 

a more seamless, hybrid look. What an N+3 aircraft might look like is 

anyone’s guess right now. But with its debut still 30 years away, NASA 

is sponsoring research that will produce a host of ideas for consideration. The Federal Aviation Administration’s current guidelines for 

overall aircraft noise footprints constitute the design baseline for all 

of NASA’s N aircraft concepts. That footprint summarizes in a single 

number, expressed as a decibel, the noise heard on the ground as an 

airplane lands, takes off, and then cuts back on power for noise abatement. The noise footprint extends ahead and behind the aircraft and 

to a certain distance on either side. NASA’s design goal is to make each 

new aircraft generation quieter than today’s airplanes by a set number of 

decibels. The N+1 goal is 32 decibels quieter than a fully noise compliant 

Boeing 737, while the N+2 goal is 42 decibels quieter than a Boeing 777. 

So far, the decibel goal for the N+1 aircraft has been elusive.82

“What makes our job very hard is that we are asked to reduce noise 

but in ways that do not adversely impact how high, far or fast an airplane is capable of flying,” Envia said.

81. Envia, “Progress Toward SFW N+1 Noise Goal,” presented at the NASA Fundamental Aeronautics Program 2nd Annual Meeting, Atlanta, Oct. 7, 2008.

82. Beth Dickey, “NASA Awards Future Vehicle Aircraft Research Contracts,” NASA Contract 

Release C08-60 (2008).
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NASA researchers have studied changes in the operation, shape, 

or materials from which key noise contributors are made. The known 

suspects include the airframe, wing flaps, and slats, along with 

components of the jet engine, such as the fan, turbine, and exhaust nozzle. While some reductions in noise can be realized with some design 

changes in these components, the overall impact still falls short of the 

N+1 goal by about 6 decibels. Envia said that additional work with 

design and operation of the jet engine’s core may make up the difference, 

but that a lot more work needs to be done in the years to come. Meanwhile, 

reaching the N+2 goals may or may not prove easier to achieve.83

“We’re starting from a different aircraft configuration, from a clean 

sheet, that gives you the promise of achieving even more aggressive goals,” 

said Russell Thomas, an aerospace engineer at Langley Research Center. 

“But it also means that a lot of your prior experience is not directly applicable, so the problem gets a lot harder from that point of view. You may 

have to investigate new areas that have not been researched heavily 

in the past.”84

Efforts to reduce noise in the N+2 aircraft have focused on the airframe, which blends the wing and fuselage together, greatly reducing 

the number of parts that extend into the airflow to cause noise. Also, 

according to Thomas, the early thinking on the N+2 aircraft is that the 

jet engines will be on top of the vehicle, using the airplane body to shield 

most of the noise from reaching the ground.

“We’re on course to do much more thorough research to get higher 

quality numbers, better experiments, and better prediction methods so 

we can really understand the acoustics of this new aircraft configuration,” Thomas said.

As for the N+3 aircraft, it remains too early to say how NASA 

researchers will use technology not yet invented to reduce noise levels 

to their lowest ever.

“Clearly significant progress has been made over the years and airplanes are much quieter than they were 20 years ago,” Envia said, noting that further reductions in noise will require whole new approaches 

to aircraft design. “It is a complicated problem and so it is a worthy 

challenge to rise up to.”

83. Don Weir, ed., “Engine Validation of Noise and Emission Reduction Technology Phase 1,” 

NASA CR-2008-215225 (2008).

84. Interview of Russell Thomas by Banke, Cape Canaveral, Feb. 4, 2009.
Case 11 | Advancing Propulsive Technology

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11

Recommended Additional Readings

Technical Reports:

“ACEE Composite Structures Technology: Review of Selected NASA 

Research on Composite Materials and Structures,” NASA CP-2321 

(1984).

“Advanced Aerodynamics, Selected NASA Research,” NASA CP-2208 

(1981).

A.P. Adamson, “Quiet Clean Short-Haul Experimental Engine (QCSEE) 

Design Rationale,” SAE Paper 750605 (1975).

“Aircraft Accident Report: United Airlines Flight 232, McDonnell Douglas 

DC-10-10, Sioux Gateway Airport, Sioux City, Iowa, July 19, 1989,” 

NTSB AAR-90-06 (1989).

R.J. Antl and J.E. Mcaulay, “Improved Components for Engine Fuel 

Savings,” NASA TM-81577 (1980).

Jennifer L. Baer-Riedhart and Robert J. Landy, “Highly Integrated 

Digital Electronic Control—Digital Flight Control, Aircraft Model 

Identification and Adaptive Engine Control,” NASA TM-86793 (1987).

Peter G. Batterton, “Energy Efficient Engine Program Contributions to 

Aircraft Fuel Conservation,” NASA TM-83741 (1984).

R.S. Beitler and G.W. Bennett, “Energy Efficient Engine: Control System 

Component Performance Report,” NASA CR-174651 (1984).

M.J. Benzakein, S.B. Kazin, and F. Montegani, “NASA/GE Quiet Engine 

‘A,’” AIAA Paper 72-657 (1972).

Jeffrey J. Berton, Edmane Envia, and Casey L. Burley, “An Analytical 

Assessment of NASA’s N1 Subsonic Fixed Wing Project Noise Goal,” 

NASA LF99-8609 (2009).
NASA’s Contributions to Aeronautics

774

11

Frank W. Burcham, Jr., L.P. Myers, and K.R. Walsh, “Flight Evaluation 

Results for a Digital Electronic Engine Control in an F-15 Airplane,” 

AIAA Paper 83-2703 (1983).

Frank W. Burcham, Jr., and Ronald J. Ray, “The Value of Early Flight 

Evaluation of Propulsion Concepts Using the NASA F-15 Research 

Airplane,” NASA TM-100408 (1987).

Cecile M. Burg, Geoffrey A. Hill, Sherilyn A. Brown, and Karl A. 

Geiselhart, “Propulsion Airframe Aeroacoustics Technology 

Evaluation and Selection Using a Multi-Attribute Decision Making 

Process and Non-Deterministic Design,” AIAA Paper 2004-4436 

(2004).

T.K. Cho and F.W. Burcham, Jr., “Preliminary Flight Evaluation of F100 

Engine Model Derivative Airstart Capability in an F-15 Airplane,” 

NASA TM-86031 (1984).

C.C. Ciepluch, “A Review of the QCSEE Program,” NASA TM-X-71818 

(1975).

C.C. Ciepluch, “Overview of the QCSEE Program,” NTRS 197.800.16123 

(1976).

C.C. Ciepluch and W.S. Willis, “QCSEE—The Key to Future Short-Haul 

Air Transport,” ICAO Bulletin No. 34 (1979).

D.B. Crawford and F.W. Burcham, Jr., “Effect of Control Logic 

Modifications on Airstart Performance of F100 Engine Model 

Derivative Engines in an F-15 Airplane,” NASA TM-85900 (1984).

Donald Y. Davis and Marshall Stearns, “Energy Efficient Engine: Flight 

Propulsion System Final Design and Analysis,” NASA CR-168219 

(1985).

“The DEEC,” NASA TF-2004-03-DFRC (2004).

“Digital Electronic Engine Control (DEEC) Flight Evaluation in an F-15 

Airplane,” NASA CP-2298 (1984).
Case 11 | Advancing Propulsive Technology

775

11

W.A. Fasching, “The CF6 Engine Performance Improvement,” NASA 

CR-165612 (1982).

“Full Scale Technology Demonstration of a Modern Counterrotating 

Unducted Fan Engine Concept: Design Test,” NASA CR-180867 

(1987) and CR-180868 (1987).

W.O. Gaffin, “The JT8D and JT9D Engine Component Improvement: 

Performance Improvement Program,” NASA CR-167965 (1982).

Robert V. Garvin, “Starting Something Big: The Commercial Emergence 

of GE Aircraft Engines,” AIAA Paper 72-657 (1999).

D.E. Gray and William B. Gardner, “Energy Efficient Engine Program 

Technology Benefit/Cost Study. Volume 1: Executive Summary,” 

NASA CR-174766 (1983).

Edward J. Hall, Joseph Rasche, Todd A. Simons, and Daniel Hoyniak, 

“NPSS Multidisciplinary Integration and Analysis,” NASA CR-2006-

213890 (2006).

Scott M. Jones, “An Introduction to Thermodynamic Performance 

Analysis of Aircraft Gas Turbine Engine Cycles Using the Numerical 

Propulsion System Simulation Code,” NASA TM-2007-214690 (2007).

Theo G. Keith, Jr., “Aeroelastic Stability and Response of Rotating 

Structures,” NASA CR-191803 (1993).

C.R. Jarvis, “An Overview of NASA’s Digital Fly-By-Wire Technology 

Development Program,” NASA 75N18246 (1975).

Anatole P. Kurkov, “Optical Measurement of Unducted Fan Blade 

Deflections,” NASA TM-100966 (1988).

S.J. Licata and F.W. Burcham, Jr., “Airstart Performance of a Digital 

Electronic Engine Control System in an F-15 Airplane,” NASA 

TM-84908 (1983).
NASA’s Contributions to Aeronautics

776

11

Mary Jo Long-Davis, “Integrated Components Technology 

Demonstrations Overview,” NTRS Document ID 200.502.14062 

(2001).

Isaac Lopez, Gregory J. Follen, Richard Guiterrez, Ian Foster, Brian 

Ginsburg, Olle Larsson, Stuart Martin, Steven Tuecke, and David 

Woodford, “NPSS on NASA’s Information Power Grid: Using 

CORBA and Globus to Coordinate Multidisciplinary Aeroscience 

Applications,” NASA TM-2000-209956 (2000).

John K. Lytle, “The Numerical Propulsion System Simulation: A 

Multidisciplinary Design System for Aerospace Vehicles,” NASA 

TM-1999-209194 (1999).

John K. Lytle, “The Numerical Propulsion System Simulation: Concept 

to Product,” NTRS 199.802.19326 (1997).

John Lytle, Greg Follen, Cynthia Naiman, Austin Evans, Joseph Veres, 

Karl Owen, and Isaac Lopez, “Numerical Propulsion System 

Simulation (NPSS) 1999 Industry Review,” NASA TM-2000-209795 

(2000).

John Lytle, Greg Follen, Cynthia Naiman, Joseph Veres, Karl Owen, 

and Isacc Lopez, “2000 Numerical Propulsion System Simulation 

Review,” NASA CP-2001-210673 (2001).

Lawrence E. Macioce, John W. Schaefer, and Neal T. Saunders, “The 

Energy Efficient Engine Project,” NASA TM-81566 (1980).

Lori Manthey, “NASA Glenn Research Center UEET (Ultra-Efficient 

Engine Technology) Program: Agenda and Abstracts,” NASA RTOP714-01-4A (2001).

J.E. Mcaulay, “Engine Component Improvement Program—Performance 

Improvement,” AIAA Paper 80-0223 (1980).

Edward T. Meleason and K.O. Johnson, “Unducted-Fan Engine,” LEW14429 (1987).
Case 11 | Advancing Propulsive Technology

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11

Glenn A. Mitchell, “Experimental Aerodynamic Performance of Advanced 

40 Degree-Swept, 10-Blade Propeller Model at Mach 0.6 to 0.85,” 

NASA TM-88969 (1988).

Lester D. Nichols and Christos C. Chamis, “Numerical Propulsion System 

Simulation: An Interdisciplinary Approach,” NASA TM-105181 

(1991).

“NPSS User Guide, Software Release: NPSS 1.6.5,” NASA NPSS-User

(2008).

Dean Olson, “Pratt and Whitney Space Propulsion NPSS Usage,” E-14760 

(2004).

John S. Orne, “Performance Seeking Control Program Overview,” NASA 

95N33011 (1995).

Khary I. Parker, James L. Felder, Thomas M. Lavella, Colleen A. Withrow, 

Albert Y. Lehmann, and V.A. William, “Integrated Control Modeling 

for Propulsion Systems Using NPSS,” NASA TM-2004-212945 (2004).

Ronald C. Plybon, Allan VanDeWall, Rajiv Sampath, Mahadevan 

Balasubramaniam, Ramakrishna Mallina, and Rohinton Irani, “High 

Fidelity System Simulation of Multiple Components in Support of 

the UEET Program,” NASA CR-2006-214230 (2006).

“Pratt & Whitney Pure Power PW1000G Engines,” Pratt & Whitney 

S16154.9.08 (2008).

“Propulsion/ACEE,” NASA FACTS-93/8-81 (1981).

T.W. Putnam, “Digital Electronic Engine Control History,” NASA 

86N25344 (1984).

“Quiet Clean Short-Haul Experimental Engine (QSCEE) Preliminary 

Analyses and Design Report,” vol. 1, NASA CR-134838 (1974).
NASA’s Contributions to Aeronautics

778

11

A.A. Saunders, Jr., and A.C. Hoffman, “Application of Digital Controls 

on the Quiet Clean Short Haul Experimental Engines,” AIAA Paper 

79-1203 (1979).

Neal T. Saunders, “Advanced Component Technologies for EnergyEfficient Turbofan Engines,” NASA TM-81507 (1980).

John R. Szuch, Dale J. Arpasi, and Anthony J. Strazisar, “Enhancing 

aeropropulsion Research with High-Speed Interactive Computing,” 

NASA TM-104374 (1991).

H.D. Sowers and W.E. Coward, “Quiet, Clean, Short-Haul, Experimental 

Engine (QCSEE) Under-The-Wing (UTW) Engine Acoustic Design,” 

NASA CR-135267 (1978).

H.D. Sowers and W.E. Coward, “Quiet, Clean, Short-Haul Experimental 

Engine (QCSEE) Over-The-Wing (OTW) Engine Acoustic Design,” 

NASA CR-135268 (1978).

James F. Stewart, Frank W. Burcham, Jr., and Donald H. Gatlin, “FlightDetermined Benefits of Integrated Flight-Propulsion Control 

Systems,” NASA TM-4393 (1992).

C.L. Stotler and A.P. Coppa, “Containment of Composite Fan Blades,” 

NASA CR-159544 (1979).

Kathleen M. Tacina and Changlie Wey, “NASA Glenn High Pressure Low 

NOx Emissions Research,” NASA TM-2008-214974 (2008).

Michael T. Tong and Scott M. Jones, “An Updated Assessment of NASA 

Ultra-Efficient Engine Technologies,” ISABE-2005-1163 (2005).

Mark G. Turner, John A. Reed, Robert Ryder, and Joseph P. Veres, “MultiFidelity Simulation of a Turbofan Engine With Results Zoomed 

Into Mini-Maps for a Zero-D Cycle Simulation,” NASA TM-2004-

213076 (2004).

Don Weir, ed., “Engine Validation of Noise and Emission Reduction 

Technology Phase 1,” NASA CR-2008-215225 (2008).
Case 11 | Advancing Propulsive Technology

779

11

Conference Presentations:

Alain Depitre, “Aircraft Noise Certification History/Development,” 

 presented at ICAO Noise Certification Workshop, Montreal, 2004.

Edmane Envia, “Progress Toward SFW N+1 Noise Goal,” presented at 

the NASA Fundamental Aeronautics Program 2nd Annual Meeting, 

Atlanta, Oct. 7, 2008.

Sanjay Garg, “NASA Glenn Research in Controls and Diagnostics 

for Intelligent Aerospace Propulsion Systems,” presented at the 

Integrated Condition Management 2006 Conference, Anaheim, CA, 

Nov. 14–16, 2006.

Edward M. Kerwin, Jr., “Procedures for Estimating the Near Field Noise 

of Rotating Aircraft Propellers,” presented at the Fifty-First Meeting of 

the Acoustical Society of America, Cambridge, MA, June 17–23, 1956.

Jeff Schweitzer, “An Overview of Recent Collaboration Research with 

NASA in Ultra High Bypass Technology,” presented at the NASA 

Fundamental Aeronautics 2007 Annual Meeting, New Orleans, Oct. 

30–Nov. 1, 2007.

Ann K. Sehra, “The Numerical Propulsion System Simulation: A Vision 

for Virtual Engine Testing,” presented at the American Society of 

Mechanical Engineers TURBO EXPO, Barcelona, Spain, May 8–11, 

2006.

Michael T. Tong and Scott M. Jones, “An Updated Assessment of NASA 

Ultra-Efficient Engine Technologies,” presented at 17th International 

Symposium on Airbreathing Engines, Munich, Germany, Sept. 4–9, 

2006.
NASA’s Contributions to Aeronautics

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11

Books and Monographs:

Alleviation of Jet Aircraft Noise Near Airports (Washington, DC: U.S. Office 

of Science and Technology, 1966).

Arnold & Porter, U.S. Government Support of the U.S. Commercial Aircraft 

Industry, Prepared for the Commission of the European Communities

(Washington, DC: Arnold & Porter, 1991).

Ronald H. Aungier, Turbine Aerodynamics: Axial-Flow and Radial-Flow 

Turbine Design and Analysis (New York: ASME Press, 2006).

David M. Bearden, Noise Abatement and Control: An Overview of Federal 

Standards and Regulations (Washington, DC: Congressional Research 

Service, 2006).

Roger E. Bilstein, Orders of Magnitude: A History of the NACA and NASA, 

1915–1990, NASA SP-4406 (Washington, DC: NASA, 1989).

Charles D. Bright, The Jet Makers: The Aerospace Industry from 1945 to 

1972 (Lawrence, KS: Regents Press of Kansas, 1978).

Mark D. Bowles and Virginia P. Dawson, “The Advanced Turboprop 

Project: Radical Innovation in a Conservative Environment,” in 

From Engineering Science to Big Science, The NACA and NASA Collier 

Trophy Research Project Winners, NASA SP-4219 (Washington, DC: 

NASA, 1998).

Walter J. Boyne and Donald S. Lopez, eds., The Jet Age: Forty Years of 

Jet Aviation (Washington, DC: Smithsonian Institution Press, 1979).

Margaret Conner, Hans von Ohain: Elegance in Flight (Reston, VA: AIAA, 

2001).

Edward W. Constant, The Origins of the Turbojet Revolution (Baltimore: 

The Johns Hopkins University Press, 1980).
Case 11 | Advancing Propulsive Technology

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11

N.A. Cumpsty, Compressor Aerodynamics (Malabar, FL: Krieger 

Publishing Company, 2004 ed.).

“Damage-Tolerant Fan Casings for Jet Engines,” Spinoff 2006 (Washington, 

DC: NASA, 2006).

Virginia P. Dawson, Engines and Innovation: Lewis Laboratory and 

American Propulsion Technology, NASA SP-4306 (Washington, DC: 

NASA, 1991).

James A. DiCarlo, Hee Mann Yun, Gregory N. Morscher, and Ramakrishna 

T. Bhatt, “High-Performance SiC/SiC Ceramic Composite Systems 

Developed for 1315 C (2400 F) Engine Components,” Research and 

Technology Report 2003 (Cleveland: NASA, 2004).

Michael Donne, Leader of the Skies, Rolls-Royce: The First Seventy-Five 

Years (London: Frederick Muller Limited, 1981).

J.H. Doolittle, The Airport and Its Neighbors, The Report of the President’s 

Airport Commission (Washington, DC: U.S. Government Printing 

Office, 1952).

Jeffrey L. Ethell, Fuel Economy in Aviation, NASA SP-462 (Washington, 

DC: NASA, 1983).

General Electric, Seven Decades of Progress: A Heritage of Aircraft Turbine 

Technology (Fallbrook, CA: Aero Publishers, Inc., 1979).

“A Giant Step in Jetliner Propulsion,” Spinoff 1996 (Washington, DC: 

NASA, 1996).

Vicki L. Golich and Thomas E. Pinelli, Knowledge Diffusion in the U.S. 

Aerospace Industry (London: Alex Publishing, 1998).

John Golley, Genesis of the Jet: Frank Whittle and the Invention of the Jet 

Engine (Ramsbury, Marlborough: The Crowood Press, 1997).
NASA’s Contributions to Aeronautics

782

11

Bob Griffiths, “Composite Fan Blade Containment Case: Innovative Use 

of Carbon-Fiber Braid Yields a Ductile Structure that Resists Blade 

Impact,” High Performance Composites (May 1, 2005).

Bill Gunston, The Development of Jet and Turbine Aero Engines

(Minneapolis: Motorbooks International, 2006).

Roy D. Hager and Deborah Vrabel, Advanced Turboprop Project, NASA 

SP-495 (Washington, DC: NASA, 1988).

T.A. Heppenheimer, Turbulent Skies: The History of Commercial Aviation

(New York: John Wiley & Sons, Inc., 1986).

Klaus Hunecke, Jet Engines: Fundamentals of Theory, Design and 

Operation (Minneapolis: Motorbooks International, 2003).

Lori A. Manthey, “Ultra-Efficient Engine Technology (UEET) Program,” 

Research and Technology Report 2001 (NASA, 2002).

Jack D. Mattingly, William H. Heiser, and David T. Pratt, Aircraft Engine 

Design (Reston, VA: AIAA, 2002).

Andrew Nahum, Frank Whittle: Invention of the Jet (Duxford, UK: Icon 

Books, Ltd., in association with the Science Museum, 2004).

Cynthia G. Naiman and Gregory J. Follen, “Numerical Propulsion 

System Simulation—A Common Tool for Aerospace Propulsion 

Being Developed,” Research and Technology Report 2000 (Cleveland: 

NASA, 2001).

NASA Glenn Research Center at Lewis Field: Achieving the Extraordinary

(Cleveland: NASA, 1999).

Guy Norris and Mark Wagner, Boeing 777: The Technological Marvel

(Osceola, WI: MBI Publishing, 2001).

Brian H. Rowe with Martin Ducheny, The Power to Fly: An Engineer’s 

Life (Reston, VA: AIAA, 2005).
Case 11 | Advancing Propulsive Technology

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11

Newell D. Sanders, Aircraft Engine Noise Reduction, NASA SP-311 

(Washington, DC: NSA, 1972).

Herb Saravanamuttoo, Gordon Rogers, and Henry Cohen, Gas Turbine 

Theory, 5th ed. (Upper Saddle River, NJ: Prentice Hall, 2001).

Robert Schlaifer and S.D. Heron, Development of Aircraft Engines and 

Fuels (Boston: Harvard University Graduate School of Business 

Administration, 1950).

Joe Shaw, “Ultra-Efficient Engine Technology Project Continued to 

Contribute to Breakthrough Technologies,” Research and Technology 

Report 2002 (Cleveland: NASA, 2003).

James St. Peter, The History of Aircraft Gas Turbine Engine Development in 

the United States . . . A Tradition of Excellence (Atlanta: International 

Gas Turbine Institute of The American Society of Mechanical 

Engineers, 1999).

Laurel J. Strauber and Cynthia G. Naiman, “Numerical Propulsion 

System Simulation (NPSS): An Award Winning Propulsion System 

Simulation Tool,” Research and Technology Report 2001 (Cleveland: 

NASA, 2002).

James E. Tomayko, Computers Take Flight: A History of NASA’s Pioneering 

Digital Fly-By-Wire Project, NASA SP-4224 (Washington, DC: NASA, 

2000).

Tom Tucker, Touchdown: The Development of Propulsion Controlled 

Aircraft at NASA Dryden (Washington, DC: NASA, 1999).

Frank Whittle, Jet: The Story of a Pioneer (New York: Philosophical 

Library, 1954).

Louis J. Williams, Small Transport Aircraft Technology (Honolulu: 

University Press of the Pacific, 2001).
784

The F-16XL-2 was one of two test aircraft that NASA Dryden used in the 1990s to test the application of laminar flow at supersonic speeds. NASA.
785

Leaner and Greener: 

Fuel Efficiency Takes Flight

Caitlin Harrington

12

CASE

12

Decades of NASA research have led to breakthroughs in understanding the physical processes of pollution and determining how to secure 

unprecedented levels of propulsion and aerodynamic efficiency to 

reduce emissions. Goaded by recurring fuel supply crises, NASA 

has responded with a series of research plans that have dramatically 

improved the efficiency of gas turbine propulsion systems, the lift-todrag ratio of new aircraft designs, and myriad other challenges.

ALTHOUGH NASA’S AERONAUTICS BUDGET has fallen dramatically 

in recent years,1 the Agency has nevertheless managed to spearhead some of America’s biggest breakthroughs in fuel-efficient and 

environmentally friendly aircraft technology. The National Aeronautics and 

Space Administration (NASA) has engaged in major programs to increase 

aircraft fuel efficiency that have laid the groundwork for engines, airframes, 

and new energy sources—such as alternative fuel and fuel cells—that are 

still in use today. NASA’s research on aircraft emissions in the 1970s also 

was groundbreaking, leading to a widely accepted view at the national—and 

later, global—level that pollution can damage the ozone layer and spawning 

a series of efforts inside and outside NASA to reduce aircraft emissions.2

This case study will explore NASA’s efforts to improve the fuel efficiency of aircraft and also reduce emissions, with a heavy emphasis 

on the 1970s, when the energy crisis and environmental concerns created a national demand for “lean and green” airplanes.3 The launch of 

1. The Administration’s fiscal year 2010 budget request for NASA is $507 million, compared with the 

1998 aeronautics budget of $1.5 billion. Opening Remarks of Senator Barbara Mikulski, Chairwoman of the Senate Appropriations Subcommittee on Commerce, Justice, Science, and Related Agencies. 

Hearing of the Fiscal Year 2010 Budget Request for NASA. U.S. Congress. May 21, 2009.

2. Joseph Chambers, Innovation in Flight (SP-2005-4539), p. 46.

3. The author thanks Robert Arrighi, archivist, Wyle Information Systems, History Program at Glenn, for 

providing instrumental documents, and Air Force Capt. Jarrett S. Lee, for providing insights and support.
NASA’s Contributions to Aeronautics

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Sputnik in 1957 and the resulting space race with the Soviet Union 

spurred the National Advisory Committee for Aeronautics (NACA)—

subsequently restructured within the new National Aeronautics and 

Space Administration—to shift its research heavily toward rocketry—

at the expense of aeronautics—until the mid-1960s.4 But as commercial air travel grew in the 1960s, NASA began to embark on a series of 

ambitious programs that connected aeronautics, energy, and the environment. This case study will discuss some of NASA’s most important 

programs in this area.

Key propulsion initiatives to be discussed include the Energy Efficient 

Engine program—perhaps NASA’s greatest contribution to fuel-efficient 

flight—as well as later efforts to increase propulsion efficiency, including the Advanced Subsonic Technology (AST) initiative and the Ultra 

Efficient Engine Technology (UEET) program. Another propulsion effort 

that paved the way for the development of fuel-efficient engine technology was the Advanced Turboprop, which led to current NASA and industry attempts to develop fuel-efficient “open rotor” concepts.

In addition to propulsion research, this case study will also explore 

several NASA programs aimed at improving aircraft structures to promote fuel efficiency, including initiatives to develop supercritical wings 

and winglets and efforts to employ laminar flow concepts. NASA has also 

sought to develop alternative fuels to improve performance, maximize 

efficiency, and minimize emissions; this case study will touch on liquid 

hydrogen research conducted by NASA’s predecessor—the NACA—as 

well as subsequent attempts to develop synthetic fuels to replace hydrocarbon-based jet fuel.

NASA’s Involvement in Energy Efficiency and Emissions Reduction

The goal of improving aircraft fuel efficiency is one shared by aerospace 

engineers everywhere: with increased efficiency come the exciting possibilities of reduced fuel costs and increased performance in terms of 

speed, range, or payload. American engineers recognized the potential 

early on and were quick to create a center of gravity for their efforts to 

improve the fuel efficiency of aircraft engines. The NACA established 

the Aircraft Engine Research Laboratory—later known as NASA Lewis 

and then NASA Glenn—in 1941 in Cleveland, OH, as the Nation’s nerve 

4. Virginia P. Dawson, Engines and Innovation: Lewis Laboratory and American Propulsion Technology, NASA SP-4306 (1991), ch. 9.
Case 12 | Leaner and Greener: Fuel Efficiency Takes Flight

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center for propulsion research.5 The lab first worked on fast fixes for piston engines in production for use in World War II, but it later moved on 

to pursue some of America’s most forward-leaning advances in jet and 

rocket propulsion.6 Improving fuel efficiency was naturally at the center of the laboratory’s propulsion research, and many of NASA’s most 

important fuel-saving engine concepts and technology originated there.7 

While NASA Glenn spearheaded the majority of aircraft fuel efficiency 

research, NASA Langley also played a critical role in the development 

of new fuel-saving aircraft structures.8

NASA’s efforts to develop aircraft technology that both increased fuel 

efficiency and reduced emissions reached their nadir in the 1970s. From 

the time of Sputnik to the late 1960s, space dominated NASA’s focus, particularly the drive to land on the Moon. But in the late 1960s, and particularly after introduction of the wide-body Boeing 747, the Agency turned 

increasing attention toward air transport, consistent with air transport 

itself dramatically increasing as a means of global mobility. Government 

and airline interest in improving jet fuel efficiency was high. However, 

NASA Lewis struggled to reenter the air-breathing propulsion game 

because the laboratory had lost much of its aeronautics expertise during 

the Sputnik crisis and now faced competition for Government support.9 

Aircraft engine companies had developed their own research facilities, 

and the U.S. Air Force (USAF) had completed its propulsion wind tunnel 

facility at Arnold Engineering Development Center in Tullahoma, TN, in 

1961.10 NASA scientists and engineers needed a new aeronautics niche. 

Luckily for them, they found it with the arrival of the oil embargo of 1973 

and the coinciding emergence of a national awareness of environmental concerns. NASA’s “clean and green” research agenda had been born.

The Organization of the Petroleum Exporting Countries (OPEC) oil 

embargo led Americans to realize that the Nation’s economy and military 

5. Neal T. Saunders and Arthur J. Glassman, “Lewis Aeropropulsion Technology: Remembering the 

Past and Challenging the Future,” NASA Lewis Research Center, 1991.

6. Ibid.

7. Ibid.

8. Phone interview of Dennis Huff by Caitlin Harrington, Mar. 31, 2009. See also Valerie J. Lyons 

and Arlene S. Levine, “An Overview of NASA’s Contributions to Energy Technology,” NASA AIAA2008-5641 (Jan. 2009).

9. Ibid.

10. Ibid. See also Arnold Engineering Development Center, Propulsion and Wind Tunnel Facility,

http://www.nimr.org/systems/images/pwt.html, accessed July 29, 2009.
NASA’s Contributions to Aeronautics

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12

were far too dependent on foreign sources of energy. In 1973, 64 percent 

of U.S. oil imports came from OPEC countries.11 The airline industry 

was particularly hard hit; jet fuel prices jumped from 12 cents to over 

$1 per gallon, and annual fuel expenditures increased to $1 billion—

triple the earnings of airlines.12 During the oil crisis, fuel accounted for 

half the airlines’ operating costs,13 and those operating costs were rising faster than the rate of inflation and faster than efficiencies in the airlines’ own operations could reduce them.14 The airline lobby descended 

on Capitol Hill, warning that its struggles to maintain profitability in the 

face of rising fuel costs were a bellwether for the Nation’s entire economy. Lawmakers turned to NASA to for help.

In 1975, the U.S. Senate asked NASA to create the Aircraft 

Energy Efficiency (ACEE) program, with the twin goals of lowering 

the fuel burn of existing U.S. commercial aircraft and building new 

fuel-efficient aircraft to match foreign competition.15 The 10-year, $670 

million ACEE yielded two of NASA’s greatest contributions to aircraft 

fuel-efficiency research. The most significant was the Energy Efficient 

Engine (E Cubed) program, which spawned technology still used in 

gas turbine engines today. The second key element of ACEE was the 

Advanced Turboprop (ATP), a bold plan to build an energy-efficient 

open-rotor engine. The open-rotor concept never made it into the mainstream, but aircraft propulsion research today still draws from ATP concepts, as this case study will later explain. Other technology developed 

under ACEE led to improved aerodynamic structures and laminar flow, 

as well as the design of supercritical wings, winglets, and composites.

Around the same time as ACEE, NASA began to sharpen its focus on 

the reduction of aircraft emissions. Space exploration had opened the 

Nation’s eyes to the fragility of the planet and the potential impact that 

11. Statistics available from Energy Information Administration, U.S. Imports by Country of Origin, 

http://tonto.eia.doe.gov/dnav/pet/pet_move_impcus_a2_nus_epc0_im0_mbblpd_a.htm,

accessed July 29, 2009.

12. Mark D. Bowles and Virginia Dawson, “The Advanced Turboprop Project: Radical Innovation 

in a Conservative Environment,” From Engineering Science to Big Science, Pamela Mack, ed., 

(Washington, DC: NASA, 1998), p. 7.

13. Ibid.

14. General Electric, “Keeping a Bold Promise,” GE Aviation, Evendale, OH, Sept. 12, 1983.

15. “The Energy Conservative Engines Program at Lewis,” Lewis News, Aug. 3, 1979, History 

Office archives, NASA Glenn Research Center.
Case 12 | Leaner and Greener: Fuel Efficiency Takes Flight

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Case 12 | Leaner and Greener: Fuel Efficiency Takes Flight

NASA slide circa 1985 showing Advanced Turboprop (ATP) requirements. NASA.

humans could have on the environment.16 The U.S. Congress pushed 

NASA to become increasingly involved in projects to study the impact 

of stratospheric flight on the ozone layer following the cancellation of 

the Supersonic Transport (SST) in 1971. The Agency provided highaltitude research aircraft, balloons, and sounding rockets for the 

Climactic Impact Assessment Program (CIAP), which was launched by 

the Department of Transportation (DOT) to examine whether the environmental concerns that helped kill the SST were valid.17

DOT and NASA’s CIAP research led to the discovery that aircraft 

emissions could, in fact, damage the ozone layer. CIAP results showed 

that nitrogen oxides would indeed cause ozone depletion if hundreds of 

Concorde and Tu-144 aircraft—the Concorde’s Russian cousin—were to 

fly as planned. Following the release of CIAP, Congress then called on 

NASA to conduct further research into the impacts of stratospheric flight 

on the ozone layer, prompting NASA and DOT to move forward with a 

16. Lyons and Levine, “An Overview of NASA’s Contributions to Energy Technology.”

17. Erick M. Conway, High Speed Dreams (Baltimore: Johns Hopkins University Press, 2005), 

p. 166.
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series of studies that by the 1980s were pointing to the conclusion that 

SSTs were less dangerous to the ozone layer than first thought.18 The 

findings gave NASA reason to believe that improvements in combustor 

technology might be enough to effectively mitigate the ozone problem.

After conducting its breakthrough ozone research, NASA has fairly 

consistently included clean combustor goals in many of its aeronautics projects in an effort to reduce aircraft emissions (examples include 

the Ultra Efficient Engine Technology program and Advanced Subsonic 

Technology program). Today, NASA has broadened its aeronautics research 

to focus not only on NOx (the collective term for water vapor, nitrogen 

oxide, and nitrogen dioxide), but also carbon dioxide and other pollutants.19

NASA’s research in this area is seen as increasingly important as the 

view that aircraft emissions harm air quality and contribute to climate 

change becomes more widely accepted. The United Nations International 

Panel on Climate Change (IPCC) issued a report in 2007 stating that aircraft emissions account for about 2 percent of all human-generated carbon dioxide emissions, which are the most significant greenhouse gas.20

The report also found that aviation accounts for about 3 percent of the 

potential warming effect of global emissions that could impact Earth’s climate.21 The report forecasts that by 2050, the aviation industry (including 

aircraft emissions) will produce about 3 percent of global carbon dioxide 

and 5 percent of the potential warming effect generated by human activity.22

In addition to NASA’s growing interest in climate change, the Agency’s 

research on improving the fuel efficiency of aircraft has also continued 

at a relatively steady pace over the years, although it has seemed to fluctuate to some extent in relation to oil prices. The oil shocks of the 1970s 

spurred a flurry of activity, from the E Cubed to the ATP and alternative fuels research. But interest in ambitious aircraft fuel-efficiency programs seemed to wane during the 1990s, when oil prices were low. Now 

18. Chambers, Innovation in Flight.

19. Phone interview with Dan Bulzan, NASA’s Associate Principle Investigator for the Subsonic 

Fixed Wing and Supersonic Aeronautics Project, by Harrington, Mar. 12, 2009.

20. International Panel on Climate Change, “Fourth Assessment Report,” 2007.

21. Ibid.

22. IPCC’s forecasts are based on projections regarding the future economy and the development 

of new energy-saving aircraft technology, so they are inherently uncertain. See U.S. Government 

Accountability Office, “Aviation and Climate Change: Aircraft Emissions Expected to Grow, but 

Technological and Operational Improvements and Government Policies Can Help Control Emissions,” GA0-09-544 (Jan. 2009).
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that oil prices are high again, however, fuel-efficiency programs seem to 

be back in vogue. (Several alternative fuels research efforts now underway at NASA will be discussed later in this case study.)

One example of the correlation between oil prices and the level of 

NASA’s interest in fuel-efficiency programs is the ATP, NASA’s ambitious 

plan to return to open-rotor engines. The concept never made it into 

mainstream use, partly because of widespread concerns that open-rotor 

engines are too noisy for commercial airline passengers,23 but also partly 

because fuel prices began to fall and there was no longer a demand for 

expensive but highly energy-efficient engines. “We were developing the 

ATP in the late ’70s and early ’80s during the fuel crisis. And while fuel 

prices went up, they didn’t continue to escalate like we originally thought 

they might, so the utility just went down; it just wasn’t cost effective,” 

said John Baughman, Manager of Military Advanced System design 

at General Electric (GE).24 With oil prices once again on the rise today, 

however, there are several new initiatives underway that take off where 

E Cubed and the ATP left off.

Lean and Clean Propulsion Systems

NASA’s efforts to improve engine design stand out as the Agency’s greatest breakthroughs in “lean and green” engine development because of 

their continuing relevance today. Engineers are constantly seeking to 

increase efficiency to make their engines more attractive to commercial airlines: with increased efficiency comes reduced fuel costs and 

increased performance in terms of speed, range, or payload.25 Emissions 

have also remained a concern for commercial aviation. The International 

Civil Aviation Organization (ICAO) has released increasingly strict standards for NOx emissions since 1981.26 The Environmental Protection 

23. Phone interview of David Ed Crow by Harrington, Pratt & Whitney Customer Training Center, 

Hartford, CT, Mar. 27, 2009. For noise concerns related to Advanced Turboprop, see also Bowles 

and Dawson, “The Advanced Turboprop Project,” From Engineering Science to Big Science,

Pamela Mack, ed., p. 27.

24. Phone interview of John Baughman, General Electric’s Manager of Military Advanced Systems 

Design, by Harrington, Apr. 24, 2009.

25. For example, see U.S. Patent No. 4,550,561, “Method for Improving the Fuel Efficiency of a 

Gas Turbine Engine,” Nov. 5, 1985, awarded to NASA Lewis researcher George A. Coffinberry 

and assigned to the U.S. Government.

26. International Civil Aviation Organization, Environment Section: Civil Aircraft Emissions, http://

www.icao.int/icao/en/env/aee.htm, accessed July 29, 2009.
NASA’s Contributions to Aeronautics

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Agency has adopted emissions standards to match those of ICAO and also 

has issued emissions standards for aircraft and aircraft engines under 

the Clean Air Act.27

NASA’s most important contribution to fuel-efficient aircraft technology to date has arguably been E Cubed, a program focused on improving propulsion systems mainly to increase fuel efficiency. The end goal 

was not to produce a production-ready fuel-efficient engine, but rather 

to develop technologies that could—and did—result in propulsion efficiency breakthroughs at major U.S. engine companies. These breakthroughs included advances in thermal and propulsive efficiency, as 

well as improvements in the design of component engine parts. Today, 

General Electric and Pratt & Whitney (P&W) continue to produce engines 

and evaluate propulsion system designs based on research conducted 

under the E Cubed program.

The U.S. Government’s high expectations for E Cubed were reflected 

in the program’s budget, which stood at about $250 million, in 1979 dollars.28 The money was divided between P&W and GE, which each used 

the funding to sweep its most cutting-edge technology into a demonstrator engine that would showcase the latest technology for conserving fuel, 

reducing emissions, and mitigating noise. Lawmakers funded E Cubed 

with the expectation that it would lead to a dramatic 12-percent reduction in specific fuel consumption (SFC), a term to describe the mass of 

fuel needed to provide a certain amount of thrust for a given period.29 

Other E Cubed goals included a 5-percent reduction in direct operating costs, a 50-percent reduction in the rate of performance deterioration, and further reductions in noise and emissions levels compared to 

other turbofan engines at the time.30

The investment paid off in spades. What began as a proposal on 

Capitol Hill in 1975 to improve aircraft engine efficiency ended in 198331

with GE and P&W testing engine demonstrators that improved SFC 

between 14 and 15 percent, exceeding the 12-percent goal. The dem27. U.S. Government Accountability Office, “Aviation’s Effects on the Global Atmosphere Are 

Potentially Significant and Expected to Grow,” GAO/RCED-00-57 (Feb. 2000) available at 

http://www.gao.gov/archive/2000/rc00057.pdf, accessed July 29, 2009.

28. “The Energy Conservative Engines Program at Lewis,” Lewis News, Aug. 3, 1979.

29. Ibid.

30. Ibid.

31. General Electric, “Keeping a Bold Promise,” GE Aviation, Evendale, OH, Sept. 12, 1983.
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onstrators were also able to achieve a reduction in emissions. A NASA 

report from 1984 hailed E Cubed for helping to “keep American engine 

technology at the forefront of the world market.”32 Engineers involved 

in E Cubed at both GE and P&W said the technology advances were 

game changing for the aircraft propulsion industry.

“The E Cubed program is probably the single biggest impact that 

NASA has ever had on aircraft propulsion,” GE’s John Baughman said. 

“The improvements in fuel efficiency and noise and emissions that have 

evolved from the E Cubed program are going to be with us for years to 

come.”33 Ed Crow, former Senior Vice President of Engineering at P&W, 

agreed that E Cubed marked the pinnacle of NASA’s involvement in 

improving aircraft fuel efficiency. “This was a huge program,” he said. “It 

was NASA and the Government’s attempt to make a huge step forward.”34

E Cubed spurred propulsion research that led to improved fuel efficiency in three fundamental ways:

First, E Cubed allowed both GE and P&W to improve the thermal 

efficiency of their engine designs. Company engineers were able to significantly increase the engine-pressure ratio, which means the pressure 

inside the combustor becomes much higher than atmospheric pressure. They were able to achieve the higher pressure ratio by improving the efficiency of the engine’s compressor, which condenses air and 

forces it into the combustor.

In fact, one of the most significant outcomes of the E Cubed program was GE’s development of a new “E Cubed compressor” that dramatically increased the pressure ratio while significantly reducing the 

number of compression stages. If there are too many stages, the engine 

can become big, heavy, and long; what is gained in fuel efficiency may 

be lost in the weight and cost of the engine. GE’s answer to that problem was to develop a compressor that had only 10 stages and produced 

a pressure ratio of about 23 to 1, compared to the company’s previous 

compressors, which had 14 stages and produced a pressure ratio of 

14 to 1.35 That compressor is still in use today in GE’s latest engines, 

including the GE-90.36

32. Jeffrey Ethell, “Fuel Economy in Aviation,” NASA SP-462 (1983), p. 42.

33. Harrington, Baughman interview.

34. Harrington, Crow interview.

35. Harrington, Baughman interview.

36. Ibid.
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P&W’s E Cubed demonstrator had a bigger, 14-stage compressor, 

but the company was able to increase the pressure ratio by modifying the compressor blades to allow for increased loading per stage. 

P&W’s engines prior to E Cubed had pressure ratios around 20 to 1; 

P&W’s E Cubed demonstrator took pressure ratios to about 33 to 1, 

according to Crow.37

The second major improvement enabled by E Cubed research was a 

substantial increase in propulsive efficiency. Air moves most efficiently 

through an engine when its velocity doesn’t change much. The way to 

ensure that the velocity remains relatively constant is to maximize the 

engine’s bypass ratio: in other words, a relatively large mass of air must 

bypass the engine core—where air is mixed with fuel—and go straight 

out the back of the engine at a relatively low exhaust speed. Both GE 

and P&W employed more efficient turbines and improved aerodynamics on the fan blades to increase the bypass ratio to about 7 to 1 (compared with about 4 to 1 on P&W’s older engines).38

Finally, E Cubed enabled major improvements in engine component parts. This was critical, because other efficiencies can’t be 

maximized unless the engine parts are lightweight, durable, and aerodynamic. Increasing the pressure ratio, for example, leads to very high 

temperatures that can stress the engine. Both P&W and GE developed materials and cooling systems to ensure that engine components 

did not become too hot.

In addition to efforts to improve fuel efficiency, E Cubed gave both 

GE and P&W opportunities to build combustors that would reduce 

emissions. E Cubed emissions goals were based on the Environmental 

Protection Agency’s 1981 guidelines and called for reductions in carbon monoxide, hydrocarbons, NOx, and smoke. Both companies developed their emissions-curbing combustor technology under NASA’s 

Experimental Clean Combustor program, which ran from 1972 to 1976. 

Their main efforts were focused on controlling where and in what proportions air and fuel were mixed inside the combustor. Managing the 

fuel/air mix inside the combustor is critical to maximize combustion 

efficiency (and reduce carbon dioxide emissions as a natural byproduct) and to ensure that temperatures do not get so high that NOx is 

generated. GE tackled the mixing issue by developing a dual annular 

37. Harrington, Crow interview.

38. Ethell, “Fuel Economy in Aviation,” p. 30.
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combustor, while P&W went with a two-stage combustor that had two 

in-line combustor zones to control emissions.39

Ultimately, E Cubed provided the financial backing required for 

both GE & P&W to pursue propulsion technology that has fed into 

their biggest engine lines. GE’s E Cubed compressor technology is used 

to power three types of GE engines, including the GE90-115B, which 

powers the Boeing 777-300ER and holds the world record for thrust.40

Other GE engines incorporating the E Cubed compressor include the 

GP-7200, which recently went into service on the Airbus A380, and the 

GE-NX, which is about to enter service on the Boeing 787.41 P&W also 

got some mileage out of the technologies developed under E Cubed. The 

company’s E Cubed demonstrator engine served as the inspiration for 

the PW2037, which fed into other engine designs that today power the 

Boeing 757 commercial airliner (the engine is designated PW2000) and 

the U.S. military’s C-17 cargo aircraft (the engine is designated F117).42

High-Speed Research

When NASA decided to start a High-Speed Research (HSR) program 

in 1990, it quickly decided to draw in the E Cubed combustor research 

to address previous concerns about emissions. The goal of HSR was to 

develop a second generation of High-Speed Civil Transport (HSCT) aircraft with better performance than the Supersonic Transport project of 

the 1970s in several areas, including emissions. The project sought to lay 

the research foundation for industry to pursue a supersonic civil transport aircraft that could fly 300 passengers at more than 1,500 miles per 

hour, or Mach 2, crossing the Atlantic or Pacific Ocean in half the time 

of subsonic jets. The program had an aggressive NOx goal because there 

were still concerns, held over from the days of the SST in the 1970s, that 

a super-fast, high-flying jet could damage the ozone layer.43

NASA’s Atmospheric Effects of Stratospheric Aircraft project was 

used to guide the development of environmental standards for the 

new HSCT exhaust emissions. The study yielded optimistic findings: 

39. Ibid., p. 31–42.

40. Harrington, Baughman interview.

41. Ibid.

42. Harrington, Crow interview.

43. NASA, NASA’s High Speed Research Program, http://oea.larc.nasa.gov/PAIS/HSROverview2.html, accessed July 29, 2009.
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there would be negligible environmental impact from a fleet of 500 

HSCT aircraft using advanced technology engine components.44 The 

HSR set a NOx emission index goal of 5 grams per kilogram of fuel 

burned, or 90 percent better than conventional technology at the time.45

NASA sought to meet the NOx goal primarily through major advancements in combustion technologies. The HSR effort was canceled in 

1999 because of budget constraints, but HSR laid the groundwork for 

future development of clean combustion technologies under the AST 

and UEET programs discussed below.

Advanced Subsonic Technology Program and UEET

NASA started a project in the mid-1990s known as the Advanced Subsonic 

Technology program. Like HSR before it, the AST focused heavily on 

reducing emissions through new combustor technology. The overall 

objective of the AST was to spur technology innovation to ensure U.S. 

leadership in developing civil transport aircraft. That meant lowering 

NOx emissions, which not only raised concern in local airport communities but also by this time had become a global concern because of 

potential damage to the ozone layer. The AST sought to spur the development of new low-emissions combustors that could achieve at least 

a 50-percent reduction in NOx from 1996 International Civil Aviation 

Organization standards. The AST program also sought to develop techniques that would better measure how NOx impacts the environment.46

GE, P&W, Allison Engines, and AlliedSignal engines all participated 

in the project.47 Once again, the challenge for these companies was to control combustion in such a way that it would minimize emissions. This 

required carefully managing the way fuel and air mix inside the combustor 

to avoid extremely hot temperatures at which NOx would be created, or at 

least reducing the length of time that the gases are at their hottest point.

44. Ibid.

45. Committee on High Speed Research, Aeronautics and Space Engineering Board, Commission 

on Engineering and Technical Systems, National Research Council, U.S. Supersonic Aircraft: Assessing NASA’s High Speed Research Program (Washington, DC: National Academies Press, 1997), 

p. 53.

46. “Advanced Subsonic Technology Program Plan,” Office of Aero-Space Technology, NASA 

(May 1999).

47. John Rohde, “Overview of the NASA AST and UEET Emissions Reductions Projects,” Mar. 1, 

2002, http://www.techtransfer.berkeley.edu/aviation02downloads/JohnRohde.pdf, accessed July 

29, 2009.
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Ultimately the AST emissions reduction project achieved its goal of 

reducing NOx emissions by more than 50 percent over the ICAO standard, a feat that was accomplished not with actual engine demonstrators 

but with a “piloted airblast fuel preparation chamber.”48

Despite their relative success, however, NASA’s efforts to improve 

engine efficiency and reduce emissions began to face budget cuts in 2000. 

Funding for NASA’s Atmospheric Effects of Aviation project, which was 

the only Government program to assess the effects of aircraft emissions 

at cruise altitudes on climate change, was canceled in 2000.49 Investments 

in the AST and the HSR also came to an end. However, NASA did manage 

to salvage parts of the AST aimed at reducing emissions by rolling those 

projects into the new Ultra Efficient Engine Technology program in 2000.50

UEET was a 6-year, nearly $300 million program managed by 

NASA Glenn that began in October 1999 and included participation 

from NASA Centers Ames, Goddard, and Langley; engine companies 

GE Aircraft Engines, Pratt & Whitney, Honeywell, Allison/Rolls Royce, 

and Williams International; and airplane manufacturers Boeing and 

Lockheed Martin.51

UEET sought to develop new engine technologies that would dramatically increase turbine performance and efficiency. It sought to reduce 

NOx emissions by 70 percent within 10 years and 80 percent within 25 

years, using the 1996 International Civil Aviation Organization guidelines 

as a baseline.52 The UEET project also sought to reduce carbon dioxide 

emissions by 20 percent and 50 percent in the same timeframes, using 1997 

subsonic aircraft technology as a baseline.53 The dual goals posed a major 

challenge because current aircraft engine technologies typically require a 

tradeoff between NOx and carbon emissions; when engines are designed 

to minimize carbon dioxide emissions, they tend to generate more NOx.

In the case of the UEET project, improving fuel efficiency was 

expected to lead to a reduction in carbon dioxide emissions by at least 

48. Ibid.

49. U.S. Government Accountability Office, “Aviation’s Effects on the Global Atmosphere Are 

Potentially Significant and Expected to Grow.”

50. “Advanced Subsonic Technology Program Plan.”

51. NASA, “Ultra Efficient Engine Technology Program,” June 25, 2001, http://www.grc.nasa.

gov/WWW/RT/RT2000/2000/2100shaw.html, accessed July 29, 2009.

52. Joe Shaw and Catherine Pettie, “NASA Ultra Efficient Engine Technology Project Overview,” 

NASA Glenn Research Center NASA CP-2004-212963, vol. 1.

53. Ibid.
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8 percent: the less fuel burned, the less carbon dioxide released.54 The 

UEET program was expected to maximize fuel efficiency, requiring engine 

operations at pressure ratios as high as 55 to 1 and turbine inlet temperatures of 3,100 degrees Fahrenheit (°F).55 However, highly efficient 

engines tend to run at very hot temperatures, which lead to the generation of more NOx. Therefore, in order to reduce NOx, the UEET program 

also sought to develop new fuel/air mixing processes and separate engine 

component technologies that would reduce NOx emissions 70 percent 

from 1996 ICAO standards for takeoff and landing conditions and also 

minimize NOx impact during cruise to avoid harming Earth’s ozone layer.

Under UEET, NASA worked on ceramic matrix composite (CMC) 

combustor liners and other engine parts that can withstand the high 

temperatures required to maximize energy efficiency and reduce carbon emissions while also lowering NOx emissions. These engine parts, 

particularly combustor liners, would need to endure the high temperatures at which engines operate most efficiently without the benefit of 

cooling air. Cooling air, which is normally used to cool the hottest parts 

of an engine, is unacceptable in an engine designed to minimize NOx, 

because it would create stoichiometric fuel-air mixtures—meaning 

the number of fuel and air molecules would be optimized so the gases 

would be at their hottest point—thereby producing high levels of NOx 

in regions close to the combustor liner.56

NASA’s sponsorship of the AST and the UEET also fed into the development of two game-changing combustor concepts that can lead to a 

significant reduction in NOx emissions. These are the Lean Pre-mixed, 

Pre-vaporized (LPP) and Rich, Quick Mix, Lean (RQL) combustor concepts. P&W and GE have since adopted these concepts to develop combustors for their own engine product lines. Both concepts focus on 

improving the way fuel and air mix inside the engine to ensure that 

core temperatures do not get so high that they produce NOx emissions.

GE has drawn from the LPP combustor concept to develop its Twin 

Annular Pre-mixing Swirler (TAPS) combustor. Under the LPP concept, 

54. NASA, Industry Roundtable: Aerospace Industry Enterprise, NASA Headquarters, http://www.

aeronautics.nasa.gov/events/showcase/environ.htm?goto=index.htm, accessed July 29, 2009.

55. Ibid.

56. Committee on High Speed Research, Aeronautics and Space Engineering Board, Commission 

on Engineering and Technical Systems, National Research Council, U.S. Supersonic Aircraft: Assessing NASA’s High Speed Research Program (1997), p. 55.
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air from the high-pressure compressor comes into the combustor through 

two swirlers adjacent to the fuel nozzles. The swirlers premix the fuel 

and combustion air upstream from the combustion zone, creating 

a lean (more air than fuel) homogenous mixture that can combust 

inside the engine without reaching the hottest temperatures, at which 

NOx is created.57

NASA support also helped lay the groundwork for P&W’s Technology 

for Advanced Low Nitrogen Oxide (TALON) low-emissions combustor, 

which reduces NOx emissions through the RQL process. The front end 

of the combustor burns very rich (more fuel than air), a process that 

suppresses the formation of NOx. The combustor then transitions in 

milliseconds to burning lean. The air must mix very rapidly with the 

combustion products from the rich first stage to prevent NOx formation as the rich gases are diluted.58 The goal is to spend almost no time 

at extremely hot temperatures, at which air and fuel particles are evenly 

matched, because this produces NOx.59

Today, NASA continues to study the difficult problem of increasing fuel efficiency and reducing NOx, carbon dioxide, and other emissions. At NASA Glenn, researchers are using an Advanced Subsonic 

Combustion Rig (ASCR), which simulates gas turbine combustion, to 

engage in ongoing emissions testing. P&W, GE, Rolls Royce, and United 

Technologies Corporation are continuing contracts with NASA to work 

on low-emissions combustor concepts.

“The [ICAO] regulations for NOx keep getting more stringent,” 

said Dan Bulzan, NASA’s associate principle investigator for the subsonic fixed wing and supersonic aeronautics project. “You can’t just sit 

there with your old combustor and expect to meet the NOx emissions 

regulations. The Europeans are quite aggressive and active in this 

area as well. There is a competition on who can produce the lowest 

emissions combustor.”60

Advanced Turboprop Project

Another significant program to emerge from NASA’s ACEE program 

was the Advanced Turboprop project, which lasted from 1976 to 1987. 

57. U.S. Supersonic Aircraft, p. 54.

58. Ibid.

59. Harrington, Crow interview.

60. Harrington, Bulzan interview.
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Like E Cubed, the ATP was largely focused on improving fuel efficiency. 

The project sought to move away from the turbofan and improve on 

the open-rotor (propeller) technology of 1950s. Open rotors have high 

bypass ratios and therefore hold great potential to dramatically increase 

fuel efficiency. NASA believed an advanced turboprop could lead to a 

reduction in fuel consumption of 20 to 30 percent over existing turbofan engines with comparable performance and cabin comfort (acceptable noise and vibration) at a Mach 0.8 and an altitude of 30,000 feet.61

There were two major obstacles to returning to an open-rotor system, however. The most fundamental problem was that propellers typically lose efficiency as they turn more quickly at higher flight speeds. The 

challenge of the ATP was to find a way to ensure that propellers could 

operate efficiently at the same flight speeds as turbojet engines. This 

would require a design that allowed the fan to operate at slow speeds 

to maximize efficiency while the turbine operates fast to achieve adequate thrust. Another major obstacle facing NASA’s ATP was the fact 

that turboprop engines tend to be very noisy, making them less than 

ideal for commercial airline use. NASA’s ATP sought to overcome the 

noise problem and increase fuel efficiency by adopting the concept of 

swept propeller blades.

The ATP generated considerable interest from the aeronautics 

research community, growing from a NASA contract with the Nation’s 

last major propeller manufacturer, Hamilton Standard, to a project 

that involved 40 industrial contracts, 15 university grants, and work at 

4 NASA research Centers—Lewis, Langley, Dryden, and Ames. NASA 

engineers, along with a large industry team, won the Collier Trophy for 

developing a new fuel-efficient turboprop in 1987.62

NASA initially contracted with Allison, P&W, and Hamilton Standard 

to develop a propeller for the ATP that rotated in one direction. This 

was called a “single rotation tractor system” and included a gearbox, 

which enabled the propeller and turbines to operate at different speeds. 

The NASA/industry team first conducted preliminary ground-testing. 

It combined the Hamilton Standard SR-7A prop fan with the Allison 

turbo shaft engine and a gearbox and performed 50 hours of success61. Bowles and Dawson, “The Advanced Turboprop Project,” p. 10.

62. NACA/NASA Research and Development Projects Receiving the Robert J Collier Trophy, 

http://history.nasa.gov/Timeline/collier.html, accessed July 29, 2009.
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Schematic drawing of the NASA propfan testbed, showing modifications and features proposed 

for the basic Grumman Gulfstream airframe. NASA.

ful stationary tests in May and June 1986.63 Next, the engine parts were 

shipped to Savannah, GA, and reassembled on a modified Gulfstream II 

with a single-blade turboprop on its left wing. Flight-testing took place in 

1987, validating NASA’s predictions of a 20 to 30 percent fuel savings.64

Meanwhile, P&W’s main rival, GE, was quietly developing its own 

approach to the ATP known as the “unducted fan.” GE released the 

design to NASA in 1983, and NASA Headquarters instructed NASA 

Lewis to cooperate with GE on development and testing.65 Citing concerns about weight and durability, P&W decided not to use a gearbox 

to allow the propellers and the turbines to turn at different speeds.66 

Instead, the company developed a counter-rotating pusher system. They 

mounted two counter-rotating propellers on the rear of the plane, which 

pushed it into flight. They also put counter-rotating blade rows in the 

turbine. The counter-rotating turbine blades were turning relatively 

63. Bowles and Dawson, “The Advanced Turboprop Project.”

64. Edwin J. Graber, “Overview of NASA PTA Propfan Flight Test Program,” NASA Glenn Research 

Center, Nov. 1, 1987, CASI Doc. 19880006423; and B.H. Little, D.T. Poland, H.W. Bartel, 

C.C. Withers, and P.C. Brown, “Propfan Test Assessment (PTA) Final Project Report,” NASA CR185138 (1989).

65. Bowles and Dawson, “The Advanced Turboprop Project.”

66. Harrington, Baughman interview.
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slowly to accommodate the fan, but because they were turning in opposite 

directions, their relative speed was high and therefore highly efficient.67

GE performed ground tests of the unducted fan in 1985 that showed 

a 20 percent fuel-conservation rate.68 Then, in 1986, a year before the 

NASA/industry team flight test, GE mounted the unducted fan—the propellers and the fan mounted behind an F404 engine—on a Boeing 727 

airplane and conducted a successful flight test.69

Mark Bowles and Virginia Dawson have noted in their analysis of the 

ATP that the competition between the two ATP concepts and industry’s 

willingness to invest in the open-rotor technology fostered public acceptance of the turboprop concept.70 But despite the growing momentum 

and the technical success of the ATP project, the open rotor was never 

adopted for widespread use on commercial aircraft. P&W’s Crow said 

that the main reason was that it was just too noisy.71 “This was clearly 

more fuel-efficient technology, but it was not customer friendly at all,” 

said Crow. Another problem was that the rising fuel prices that had 

spurred NASA to work on energy-efficient technology were now going 

back down. There was no longer a favorable ratio of cost to develop turboprop technology versus savings in fuel burn.72 “In one sense of the 

word it was a failure,” said Crow. “Neither GE nor Pratt nor Boeing nor 

anyone else wanted us to commercialize those things.”

Nevertheless, the ATP yielded important technological breakthroughs 

that fed into later engine technology developments at both GE and P&W. 

Crow said the ATP set the stage for the development of P&W’s latest engine, 

the geared turbofan.73 That engine is not an open-rotor system, but it does 

use a gearbox to allow the fan to turn more slowly than the turbines. The 

fan moves a large amount of air past the engine core without changing 

the velocity of the air very much. This enables a high bypass ratio, thereby 

increasing fuel efficiency; the bypass ratio is 8 to 1 in the 14,000–17,000-

pound thrust class and 12 to 1 in the 17,000–23,000-pound thrust class.74

67. Ibid.

68. James J. Haggerty, “Propfan Update,” Aerospace, fall/winter 1986, pp. 10–11.

69. Graber, “Overview of NASA PTA”; Harrington, Baughman interview.

70. Bowles and Dawson, “The Advanced Turboprop Project.”

71. Harrington, Crow interview.

72. Bowles and Dawson, “The Advanced Turboprop Project.”

73. Harrington, Crow interview.

74. United Technologies, Inc., Leading Change: The Geared Turbofan Engine, Pratt & Whitney 

Customer Training Center, Hartford, CT, 2009.
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GE renewed its ATP research to compete with P&W’s geared turbofan, announcing in 2008 that it would consider both open rotor and 

encased engine concepts for its new engine core development program, 

known as E Core. The company announced an agreement with NASA 

in the fall of 2008 to conduct a joint study on the feasibility of an openrotor engine design. In 2009, GE plans to revisit its original open-rotor 

fan designs to serve as a baseline. GE and NASA will then conduct wind 

tunnel tests using the same rig that was used for the ATP. 75 Snecma, GE’s 

50/50 partner in CFM International—an engine manufacturing partnership—will participate in fan blade design testing. GE says the new E 

Core design—whether it adopts an open rotor or not—aims to increase 

fuel efficiency 16 percent above the baseline (a conventional turbofan 

configuration) in narrow-body and regional aircraft.76

Another major breakthrough resulting from the ATP was the development of computational fluid dynamics (CFD), which allowed engineers to predict the efficiency of new propulsion systems more accurately. 

“What computational fluid dynamics allowed us to do was to design a 

new air foil based on what the flow field needed rather than proscribing a fixed air foil before you even get started with a design process,” 

said Dennis Huff, NASA’s Deputy Chief of the Aeropropulsion Division. 

“It was the difference between two- and three-dimensional analysis; 

you could take into account how the fan interacted with nacelle and 

certain aerodynamic losses that would occur. You could model numerically, whereas the correlations before were more empirically based.”77 

Initially, companies were reluctant to embrace NASA’s new approach 

because they distrusted computational codes and wanted to rely on 

existing design methods, according to Huff. However, NASA continued 

to verify and validate the design methods until the companies began to 

accept them as standard practice. “I would say by the time we came out 

of the Advanced Turboprop project, we had a lot of these aerodynamic 

CFD tools in place that were proven on the turboprop, and we saw the 

companies developing codes for the turbo engine,” Huff said.78

75. Harrington, Baughman interview.

76. General Electric, “GE Launches Engine Core for New Commercial Airplanes,” July 13, 2008, 

available at http://www.businesswire.com/portal/site/home/permalink/?ndmViewId=news_

view&newsId=20080713005020, accessed July 29, 2009.

77. Harrington, Huff interview.

78. Harrington, Huff interview.
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Aircraft Materials and Structures

While refinements in engine design have been the cornerstone of 

NASA’s efforts to improve fuel efficiency, the Agency has also sought to 

improve airframe structures and materials. The ACEE included not only 

propulsion improvement programs but also efforts to develop lightweight composite airframe materials and new aerodynamic structures 

that would increase fuel efficiency. Composite materials, which consist 

of a strong fiber such as glass and a resin that binds the fibers together, 

hold the potential to dramatically reduce the weight—and therefore 

the fuel efficiency—of aircraft.

Initially, Boeing began to investigate composite materials, using 

fiberglass for major parts such as the radome on the 707 and 747 commercial airliners.79 Starting around 1962, composite sandwich parts 

comprised of fiberglass-epoxy materials were applied to aircraft such as 

the Boeing 727 in a highly labor-intensive process.80 The next advance in 

composites was the use of graphite composite secondary aircraft structures, such as wing control surfaces, wing trailing and leading edges, 

vertical fin and stabilizer control surfaces, and landing gear doors.81

NASA research on composite materials began to gain momentum 

in 1972, when NASA and the Air Force undertook a study known as 

Long Range Planning Study for Composites (RECAST) to examine the 

state of existing composites research. The RECAST study found two 

major obstacles to the use of composites: high costs and lack of confidence in the materials.82

However, by 1976, interest in composite materials had picked up 

steam because they are lighter than aluminum and therefore have the 

potential to increase aircraft fuel efficiency. Research on composites 

was formally wrapped into ACEE in the form of the Composite Primary 

Aircraft Structures program. NASA hoped that research on composites 

would yield a fuel savings for large aircraft of 15 percent by the 1990s.

NASA’s efforts under ACEE ultimately led the aircraft manufacturing industry to normalize the use of composites in its manufacturing 

79. Chambers, “Concept to Reality: Contributions of the NASA Langley Research Center to Civil 

Aircraft of the 1990s,” NASA SP 2003-4529 (2003), http://oea.larc.nasa.gov/PAIS/

Concept2Reality/composites.html, accessed July 29, 2009.

80. Ibid.

81. Ibid.

82. Ethell, “Fuel Economy in Aviation,” p. 59.
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processes, driving down costs and making composites far more common in aircraft structures. “Ever since the ACEE program has existed, 

manufacturers have been encouraged by the leap forward they have 

been able to make in composites,” Jeffrey Ethell, the late aviation author 

and analyst, wrote in his 1983 account NASA’s fuel-efficiency programs. 

“They have moved from what were expensive, exotic materials to 

routine manufacture by workers inexperienced in composite structures.”83 

Today, composite materials have widely replaced metallic materials 

on parts of an aircraft’s tail, wings, fuselage, engine cowlings, and 

landing gear doors.84 

NASA research under ACEE also led to the development of improved 

aerodynamic structures and active controls. This aspect of ACEE was 

known as the Energy Efficient Transport (EET) program. Aerodynamic 

structures can improve the way that the aircraft’s geometry affects the 

airflow over its entire surface. Active controls are flight control systems 

that can use computers and sensors to move aircraft surfaces to limit 

unwanted motion or aerodynamic loads on the aircraft structure and 

to increase stability. Active controls lighten the weight of the aircraft, 

because they replace heavy hydraulic lines, rods, and hinges. They also 

allow for reductions in the size and weight of the wing and tail. Both 

aerodynamic structures and active controls can increase fuel efficiency 

because they reduce weight and drag.85

One highly significant aerodynamic structure that was explored 

under ACEE was the supercritical wing. During the 1960s and 1970s, 

Richard Whitcomb, an aeronautical engineer at NASA Langley Research 

Center, led the development of the new airfoil shape, which has a flattened top surface to reduce drag and tends to be rounder on the bottom, with a downward curve at the trailing edge to increase lift. ACEE 

research at NASA Dryden led to the finding that the supercritical wing 

could lead to increased cruising speed and flight range, as well as an 

83. Ibid., p. 71.

84. “NASA Technology,” Dryden Flight Research Center, Oct. 21, 2008, available at http://

www.nasa.gov/centers/dryden/news/X-Press/stories/2008/10_08_technology.html, accessed 

July 29, 2009.

85. Ethell, “Fuel Economy in Aviation,” p. 77; see also N.A. Radovich, D. Dreim, D.A. O’Keefe, 

L. Linner, S.K. Pathak, J.S. Reaser, D. Richardson, J. Sweers, and F. Conner, “Study for the Optimization of a Transport Aircraft Wing for Maximum Fuel Efficiency,” vol. 1: “Methodology, Criteria, 

Aeroelastic Model Definition, and Results,” NASA CR-172551 (1985).
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increase in fuel efficiency of about 15 percent over conventional-wing 

aircraft. Supercritical wings are now in widespread use on modern 

subsonic commercial transport aircraft.86

Whitcomb also conducted research on winglets, which are vertical extensions of wingtips that can improve an aircraft’s fuel efficiency 

and range. He predicted that adding winglets to transport-size aircraft 

would lead to improved cruising efficiencies between 6 and 9 percent. 

In 1979 and 1980, flight tests involving a U.S. Air Force KC-135 aerial 

refueling tanker demonstrated an increased mileage rate of 6.5 percent.87 

The first big commercial aircraft to feature winglets was the MD-11, 

built by McDonnell-Douglas, which is now a part of Boeing. Today, 

winglets can be are commonly found on many U.S.- and foreign-made 

commercial airliners.88

Laminar flow is another important fuel-saving aircraft concept spearheaded by NASA. Aircraft designed to maximize laminar flow offer the 

potential for as much as a 30-percent decrease in fuel usage, a benefit 

that can be traded for increases in range and endurance. The idea behind 

laminar flow is to minimize turbulence in the boundary layer—a layer 

of air that skims over the aircraft’s surface. The amount of turbulence in 

the boundary layer increases along with the speed of the aircraft’s surface and the distance air travels along that surface. The more turbulence, 

the more frictional drag the aircraft will experience. In a subsonic transport aircraft, about half the fuel required to maintain level flight in cruise 

results from the necessity to overcome frictional drag in the boundary layer.89

There are two types of methods used to achieve laminar flow: active 

and passive. Active Laminar Flow Control (LFC) seeks to reduce turbulence in the boundary layer by removing a small amount of fluid (air) 

from the boundary layer. Active LFC test sections on an aircraft wing 

contain tiny holes or slots that siphon off the most turbulent air by using 

an internal suction system. Passive laminar flow does not involve a suc86. Dryden Flight Research Center, “NASA Technology.” Whitcomb’s work is thoroughly discussed 

by Jeremy Kinney in a companion case study.

87. Dryden Flight Research Center, “NASA Dryden Technology Facts: Winglets,” http://www.

nasa.gov/centers/dryden/about/Organizations/Technology/Facts/TF-2004-15-DFRC. html,

accessed July 29, 2009.

88. Ibid.

89. Albert L. Braslow, A History of Suction-Type Laminar Flow Control with Emphasis on Flight 

Research, Monographs in Aerospace History, No. 13 (Washington, DC: NASA History Division, 

Office of Policy and Plans, NASA Headquarters, 1999), p. 1.
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An F-16XL flow visualization test. This F-16 Scamp model was tested in the NASA Langley 

Research Center Basic Aerodynamics Research Tunnel. This was a basic flow visualization test 

using a laser light sheet to illuminate the smoke. NASA.

tion system to remove turbulent air; instead, it relies on careful contouring of the wing’s surface to reduce turbulence.90

In 1990, NASA and Boeing sponsored flight tests of a Boeing 757 that 

used a hybrid of both active and passive LFC. The holes or slots used in 

active LFC can get clogged with bugs. As a result, NASA and Boeing used 

90. Dryden Flight Research Center, “Fact Sheet: F-16XL Laminar Flow Research Aircraft,” updated 

Mar. 1, 2008, http://www.nasa.gov/centers/dryden/news/FactSheets/FS-023-DFRC.html, 

accessed July 29, 2009.
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a hybrid LFC system on the 757 that limited the air extraction system 

to the leading edge of the wing, followed by a run of the natural laminar flow.91 Based on the flight tests, engineers calculated that the application of hybrid LFC on a 300-passenger, long-range subsonic transport 

could provide a 15-percent reduction in fuel burned, compared with 

a conventional equivalent.92

NASA laminar flow research continued to evolve, with NASA Dryden 

conducting flight tests on two F-16 test aircraft known as the F-16XL-1 

and F-16XL-2 in the early and mid-1990s. The purpose was to test the 

application of active and passive laminar flow at supersonic speeds. 

Technical data from the tests are available to inform the development 

of future high-speed aircraft, including commercial transports.93

Today, laminar flow research continues, although active LFC, required 

for large transport aircraft, has not yet made its way into widespread use 

on commercial aircraft. However, NASA is continuing work in this area. 

NASA’s subsonic fixed wing project, the largest of its four aeronautics 

programs, is working on projects to reduce noise, emissions, and fuel 

burn on commercial-transport-size aircraft by employing several technology concepts, including laminar flow control. The Agency is hoping 

to develop technology to reduce fuel burn for both a next generation of 

narrow-body aircraft (N+1) and a next generation of hybrid wing/body 

aircraft (N+2).94 NASA is expected to conduct wind tunnel tests of two 

hybrid wing body (also known as blended wing body) aircraft known as 

N2A and N2B in 2011. Those aircraft, which will incorporate hybrid LFC, 

are expected to reduce fuel burn by as much as 40 percent.95

Together with this research on emissions and fuel burn has come a 

heightened awareness on reducing aircraft noise. One example of a very 

91. Mary L. Sandy and H. Keith Henry, “Flight Tests Prove Concept for Jet Liner Fuel Economy,” 

NASA Headquarters Press Release, Aug. 23, 1990, http://www.nasa.gov/home/

hqnews/1990/90-115.txt, accessed July 29, 2009.

92. Braslow, A History of Suction-Type Laminar Flow Control with Emphasis on Flight Research, p. 32.

93. Dryden Flight Research Center, “Fact Sheet: F-16XL Laminar Flow Research Aircraft.”

94. Graham Warwick, “NASA Steps Up Research into Fuel Efficient Aircraft,” Flight International,

June 14, 2007, available at http://www.flightglobal.com/articles/2007/06/14/214635/

nasa-steps-up-research-into-efficient-aircraft.html, accessed July 29, 2009.

95. Warwick, “NASA Pushes Blended Wing/Body,” Aviation Week, Jan. 13, 2009, available at http://www.aviationweek.com/aw/generic/story.jsp?id=news/Body011309.

xml&headline=NASA%20Pushes%20Blended%20Wing/Body%20&channel=space, accessed July 

29, 2009.
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beneficial technical “fix” to the noise problem is the chevron exhaust 

nozzle, so called because it has a serrated edge resembling a circular 

saw blade, or a series of interlinked chevrons. The exhaust nozzle chevron has become a feature of recent aircraft design, though how to best 

configure chevron shapes to achieve maximum noise-reduction benefit without losing important propulsive efficiencies is not yet a refined 

science. The takeoff noise reduction benefits, when “traded off” against 

potential losses in cruise efficiency, clearly required continued study, 

in much the same fashion that, in the piston-engine era, earlier NACA 

engineers grappled with assessing the benefits of the controllable-pitch 

propeller and the best way to configure early radial engine cowlings. 

As that resulted in the emergence of the NACA cowling as a staple and 

indeed, design standard, for future aircraft design, so too, presumably, 

will NASA’s work lead to better understanding of the benefits and design 

tradeoffs that must be made for chevron design.96

Hydrogen Research Leads to Rockets and Fuel Cells

While most of NASA’s aircraft fuel-efficiency research grew out of the 

reality jolt of the 1970s oil crisis and environmental concerns, there is 

at least one notable exception. Researchers first began to investigate liquid hydrogen as an alternative to hydrocarbon-based fuel in the mid1950s because they suspected major performance efficiencies could 

be gained.97 The Lewis Flight Propulsion Laboratory issued a seminal 

report in April 1955—although it was not declassified until September 

1962—suggesting that liquid hydrogen might have a positive impact on 

the performance of high-altitude military aircraft (subsonic and supersonic bombers, fighters, and reconnaissance aircraft flying at 75,000 to 

85,000 feet).98 While the report raised the aviation community’s awareness of the potential for hydrogen as a fuel source, it did not lead to 

widespread use in aircraft because of technical problems with using 

96. Travis L. Turner, Randolph H. Cabell, Roberto J. Cano, and Richard J. Silcox, “Testing of 

SMA-Enabled Active Chevron Prototypes Under Representative Flow Conditions,” NASA Langley 

Research Center, Report LAR-17332, Paper 6928-36, CASI Doc. 20080014174 (2008); Vance 

Dippold, III, “CFD Analyses and Jet-Noise Predictions of Chevron Nozzles with Vortex Stabilization,” 

NASA TM-2008-215150 (2008).

97. Abe Silverstein and Eldon Hall, “Research Memorandum: Liquid Hydrogen as Jet Fuel for High 

Altitude Aircraft,” Lewis Flight Propulsion Laboratory, Cleveland, OH, N63-12541, (Apr. 15, 1955).

98. Ibid.
NASA’s Contributions to Aeronautics

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The Martin B-57B light bomber that Lewis modified to test hydrogen propulsion, shown on the 

snowy ramp of Lewis Laboratory (now Glenn Research Center) in 1958. NASA.

hydrogen inside an aircraft. Again, this early interest in hydrogen did 

not reflect environmental or conservation concerns, but rather, a desire 

to achieve much higher flight vehicle performance.

In the report, two NACA researchers—Abe Silverstein and Eldon 

Hall—argued the case for using liquid hydrogen, noting it has a special advantage as an aviation fuel: a high heating value. This means that 

it takes less hydrogen fuel than hydrocarbon fuel to achieve the same 

thrust. That advantage could prove particularly important at high altitudes, the researchers noted, where maximizing fuel efficiency is critical to make up for other penalties associated with flying high.

Indeed, one of the downsides of a high-altitude flight regime—in 

which atmospheric pressure is low—is that it generally requires heavy, 

high-pressure-ratio engines to ensure combustion is sustained and thrust 

levels are adequate. The NACA report speculated that it might be possible to use lighter engines—albeit less-efficient ones, with lower pressure 

ratios—if liquid hydrogen were used for fuel. Liquid hydrogen requires 

less combustion volume than hydrocarbons do, making shorter and 

lighter engines feasible. And, with its high heating value, liquid hydrogen fuel generates more thrust per pound than hydrocarbons do, even 

if it’s being used in a light engine running at a lower pressure ratio. The 

report posited that if every pound of weight saved by using a lighter 
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A 1978 NASA conceptual design study for a liquid-hydrogen-powered jetliner. Note the large volume required by liquid hydrogen, compared with conventional hydrocarbon-fueled airliners. NASA.

engine could be replaced by a pound of liquid hydrogen fuel, an aircraft 

could be over twice as effective in extending its range at high altitudes.

After Silverstein and Hall issued their report, the NACA conducted 

experiments showing that hydrogen had a high combustion efficiency 

in a turbojet combustor even in low-pressure conditions. In 1956, NACA 

researchers at Lewis made “three completely successful flights” using 

liquid hydrogen in one engine of a modified Martin B-57B jet bomber, 

thereby effectively demonstrating that liquid hydrogen could be handled 

and jettisoned safely and was feasible for use in aircraft.99 Meanwhile, 

from 1956 to 1958, the U.S. Air Force began work on a secret project, 

known as Suntan, to develop a high-altitude, hydrogen-fueled aircraft 

with performance superior to the secret U-2 spy plane of the Counter 

Intelligence Agency (CIA).

The use of liquid hydrogen in aircraft would have marked a major 

breakthrough in terms of high altitude flight because engine weight is 

99. Donald R. Mulholland, Loren W. Acker, Harold H. Christenson, and William V. Gough, “Flight 

Investigation of a Liquid-Hydrogen Fuel System,” NACA RM-E57F19a (1957).
NASA’s Contributions to Aeronautics

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“the single most important variable determining the height at which 

an airplane can fly.”100 Liquid hydrogen offered the potential to fly at 

high altitudes at an extended range. Despite its potential, however, neither the NACA nor the Air Force was able to convince enough stakeholders in Government and industry that liquid hydrogen was a viable 

candidate for aviation fuel.

Hydrogen’s excellent combustion qualities raised questions about 

whether it could be safely transported or carried inside aircraft. To be 

sure, NACA flight tests demonstrated that safe handling of hydrogen fuel 

on the ground and in the air was possible. Also, the Air Force conducted 

tests in which liquid hydrogen tanks under pressure were ruptured, and 

it found that in many cases the hydrogen quickly escaped without ignition. Yet concerns about safety persisted, and, in a tight budget climate, 

hydrogen-fuelled aircraft lost out to other priorities. After receiving a full 

briefing on Suntan, Gen. Curtis E. Lemay, the former head of Strategic 

Air Command who had moved up to Vice Chief of Staff in July 1957, 

raised concerns about safety. “What,” he said, “put my pilots up there 

with a bomb?”101

Others questioned whether using liquid hydrogen would truly yield 

big gains in aircraft range at high altitudes. Hydrogen has a high volume—10 times that of hydrocarbons—which means that the aircraft 

fuselage has to be bigger and weigh more to accommodate the fuel. 

Silverstein and Hall argued that there would be more room for hydrogen fuel tanks in high-altitude aircraft, which would need larger wings 

and therefore a bigger fuselage to provide lift in the thin air of the upper 

atmosphere (a bigger fuselage would mean more room for hydrogen fuel 

tanks). But while it might have been possible to extend the range of aircraft because of the increased efficiency of liquid hydrogen, others questioned whether hydrogen-fuelled aircraft would still be fairly limited by 

the tremendous amount of fuel storage capacity that hydrogen requires. 

Kelly Johnson, the Lockheed Martin engineer who designed the U-2 and 

the hydrogen-fueled CL-400 for the Air Force’s Suntan project, said he 

could see a range growth of only 3 percent from adding more hydrogen 

fuel storage capacity to his CL-400 design. “We have crammed the max100. Silverstein and Hall, “Research Memorandum: Liquid Hydrogen as Jet Fuel for High Altitude 

Aircraft.”

101. John L. Sloop, “Liquid Hydrogen as a Propulsion Fuel,” NASA SP-4404 (Washington, DC: 

NASA, 1978).
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imum amount of hydrogen in the fuselage that it can hold. You do not 

carry hydrogen in the flat surfaces of the wing,” he said.102

While liquid hydrogen is highly energetic, it has far less energy 

density than hydrocarbon fuels. Thus, to get an equivalent amount of 

energy from hydrogen requires a much greater volume. Accordingly, a 

hydrogen airplane would have extremely large fuel tanks, which, having to be supercold as well, pose significant technical challenges to aircraft designers. Researchers have not yet found a way to overcome the 

challenges associated with hydrogen’s large volume, which forces aircraft design compromises and requires complex ground transportation, 

storage distribution, and vent capture system. Moreover, hydrogen is 

not a viable source of energy in itself; producing it requires the use of 

other sources of energy—such as electric power produced by nuclear 

fusion as well as a large source of clean water. However, in one respect, 

hydrogen could “pay back” this “debt,” for it could be used to enrich 

the production process of synthetic fuel, achieving similar production 

efficiencies while reducing the amount of water and coal traditionally 

required for enrichment.103

Despite these technical challenges, NASA’s research on the use of 

hydrogen to power aircraft did lead to some important findings: namely, 

that hydrogen is a potentially promising turbojet fuel in a high-altitude, 

low-speed flight regime. These conditions favor a fuel that can operate 

efficiently in low-pressure conditions. High altitudes also favor a largevolume aircraft, helping to offset the disadvantage of hydrogen’s low 

density. Given these attractive characteristics, the prospect of using 

hydrogen as an aircraft propellant has continued to resurface in the past 

decade, especially when the cost of hydrocarbon-based fuel rises. For 

example, NASA’s Zero CO2 research project sought to eliminate carbon 

dioxide and lower NOx emissions by converting propulsion systems to 

hydrogen fuel.104 One new propulsion technology that NASA engineers 

considered as part of Zero CO2 was the use of fuel cells, which are discussed below. A NASA Glenn Web page updated as recently as 2008 says 

that the Combustion Branch of NASA’s Propulsion Division is still study102. Ibid.

103. D. Daggett, O. Hadaller, R. Hendricks, and R. Walther, “Alternative Fuels and Their Potential 

Impact on Aviation,” NASA TM-2006-214365 (2006); see also C.G. Jay, “Hydrogen Enrichment 

of Synthetic Fuel,” NASA Tech Briefs, vol. 3, no. 1 (June 1, 1978), NTTC Doc. 19780000039.

104. Lyons and Levine, “An Overview of NASA’s Contributions to Energy Technology.”
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ing hydrogen combustion to demonstrate that hydrogen can be used as 

an aviation fuel to minimize emissions.105

The NACA’s early research on hydrogen-fuelled aircraft also created an awareness of hydrogen as a potential fuel source that did not 

exist before Silverstein and Hall embarked on their study. This awareness helped lead to important breakthroughs in rocketry and fuel 

cell research. In particular, research on the use of hydrogen in airbreathing aircraft laid the groundwork for the successful development 

of hydrogen-fueled rockets in the mid-1950s. In fact, Silverstein and 

Hall’s research helped to inform NASA’s decision in 1959 to use liquid 

hydrogen as a propellant in the upper stage of the Saturn launch 

vehicle. That decision was one of the keys to the success of the Apollo 

Moon landing missions of the 1960s and 1970s.106

The NACA’s early efforts to draw attention to hydrogen as a power 

source also led to the development of fuel cells for the Apollo and Gemini 

capsules. Apollo employed the world’s first fuel cells, which used hydrogen and oxygen to generate onboard power for Apollo command and 

service modules. Fuel cells are essentially plastic membranes treated 

with a special catalyst; hydrogen seeps into the membrane and meets 

up with the oxygen inside to generate electricity and water. The fuel 

cells used on the Apollo proved so successful that they were once again 

employed on the Space Shuttle orbiter.

Fuel Cells for Aircraft

Observing the success of fuel cells in space flight, NASA researchers in 

the late 1990s began to explore the potential for fuel cells to power aircraft. The cells were attractive to NASA’s aeronautics directorate because 

they have near-zero emissions, are quiet and efficient, and can work in 

any environment where the temperature is lower than the cell’s operating temperature. Valerie Lyons, NASA’s Division Chief of Power and 

In-Space Propulsion, said she began pushing for NASA aeronautics to 

pursue fuel cell research about 10 years ago. “I would venture to say 

they hadn’t really looked at it before that,” she said. “When I looked at 

the fuel cell area, I said, ‘This is pretty interesting, can we use this?’”107 

105. NASA Glenn Research Center Propulsion Systems Division Combustion Branch, updated July 

24, 2008, available at http://www.grc.nasa.gov/WWW/combustion, accessed July 29, 2009.

106. Sloop, “Liquid Hydrogen as a Propulsion Fuel.”

107. Harrington, Lyons phone interview, Apr. 27, 2009.
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One of NASA’s main fuel cell initiatives was the Next Generation 

Clean Aircraft Power (NEXCAP) program, which sought to use the cells 

as auxiliary power units (APUs) for aircraft.108 The APUs could be used 

in onboard electrical systems to power a grounded aircraft, providing 

an alternative to wasting fuel and producing emissions by drawing on 

power from an idling engine. NASA researchers hoped that the APU 

research would eventually lead to the design and test flight in 2030 of 

an electric airplane, which would rely on fuel cells for propulsion.109 

While NASA’s electric airplane never came to fruition, Boeing flew the 

first piloted airplane powered by hydrogen fuel cells in 2008.110 NASA 

researchers maintain a strong interest in the potentiality of fuel cells to 

meet the future energy needs of the Nation.

The 1970s and the Rise of Synthetic Fuels

NASA’s interest in alternative fuels did not end with liquid hydrogen; synthetic fuel research, joined with research on new, more aerodynamically efficient aircraft configurations, took off in the 1970s and 1980s, as rising oil 

prices and a growing concern about mankind’s (and aviation’s) impact on 

the environment pushed researchers to seek alternatives to oil-based fuel.111 

In 1979, NASA Langley released an aircraft fuel study that compared 

liquid hydrogen, liquid methane, and synthetic aviation kerosene derived 

from coal or oil shale.112 The study took into account factors including 

cost, capital requirements, and energy resources required to make the 

fuel. These factors were considered in light of the practicality of using 

the fuel in terms of the fuel production processes, transportation, storage, and its suitability for use on aircraft. Environmental emissions and 

safety aspects of the fuel also were considered. The study concluded that 

all three fuels met the criteria, but that synthetic aviation kerosene was 

the most attractive because it was the least expensive.113 

108. Anita Liang, Chief, Aeropropulsion Projects Office, “Emerging Fuel Cell Developments at 

NASA for Aircraft Applications,” Glenn Research Center, 2003, available at http://www.netl.doe.

gov/publications/proceedings/03/seca/Anita%20Liang.pdf, accessed July 29, 2009.

109. Ibid.

110. Boeing, “Boeing Successfully Flies First Fuel-Cell Powered Plane,” Apr. 3, 2008, http://

www.boeing.com/news/releases/2008/q2/080403a_nr.html, accessed July 29, 2009.

111. See, for example, A.L. Nagel, W.J. Alford, and J.F. Dugan, “Future Long-Range Transports: 

Prospects for Improved Fuel Efficiency,” NASA TM-X-72659 (1975).

112. Robert D. Wicofski, “Comparison of Alternate Fuel for Aircraft,” NASA TM-80155 (Sept. 1979).

113. Ibid.
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Despite the promising findings of NASA’s study, however, synthetic 

fuel never made it into mainstream production. The fuel’s capital costs are 

still relatively high when compared with oil-based jet fuel, because new 

synthetic fuel production plants have to be built to produce the fuel.114 

Private industry has been hesitant to get into this business, fearing it would 

not make a return on its investment. If oil prices were to drop—as they 

did in the mid-1980s— companies that invested in synthetic aircraft fuel 

production would find it difficult to compete with cheap oil-based jet fuel.

Regardless of industry’s hesitation, Government efforts to develop 

and test alternative fuels are springing to life again as a result of a return 

to high oil prices and a growing concern about the impact of emissions 

on air quality and climate change. The U.S. Air Force has engaged in a 

systematic process to certify all of its aircraft to fly on a 50/50 blend of 

oil-based jet fuel and synthetic fuel. Air Force officials hoped that testing 

and flying their own aircraft on synthetic fuels would encourage commercial airlines to do the same, believing that if the service and airline 

industry could create a buyer’s market for synthetic fuel, then the energy 

industry might be more amenable to investing the money required to 

build synthetic fuel plants for mass production.115 

NASA has also begun testing the performance and emissions of two 

synthetic fuels derived from coal and natural gas. While the Air Force’s 

interest in alternative fuels is largely related to concerns about oil price 

volatility and the national security risks of relying on foreign oil suppliers, NASA has embarked on alternative fuels research largely to study 

the potential for reducing emissions. NASA’s research effort, which is 

being conducted at NASA Dryden, seeks to closely measure particulate 

levels. “Even though there are no current regulations for particulates, 

we see particulates as being very important,” said Bulzan, who is leading the alternative fuels effort. “They are very important to local air quality when the aircraft is taking off and landing at the airport, and they 

can also generate cloud formation that can affect global warming.”116 

114. A plant that produces Fischer-Tropsch synthetic fuel, which is being used by the U.S. Air Force to 

power jet engines, requires five times the capital costs as an oil refinery. See Gerrard Cowan and Harrington, “Adding Fuel to the Fire: Energy and Defence,” Jane’s Defence Weekly, Nov. 21, 2008.

115. Harrington, “USAF Promotes Fuel Alternative,” Jane’s Defence Weekly, Mar. 16, 2007.

116. NASA has had a long-standing interest in such research. See, for example, R. Bradford, W.T. 

Atkins, R.M. Bass, R. Dascher, J. Dunkin, N. Luce, W. Seward, and D. Warren, “Coal Conversion 

and Synthetic-Fuel Production,” NASA Tech Briefs, vol. 5, no. 1 (Aug. 1, 1980), p. 56.
Case 12 | Leaner and Greener: Fuel Efficiency Takes Flight

817

12

Both the USAF and NASA are using synthetic fuel derived from a 

process developed by the Germans in World War II known as FischerTropsch. In this process, a mixture of carbon monoxide and hydrogen is 

used to create liquid hydrocarbons for fuel. NASA Dryden’s latest alternative fuels testing, which took place in early 2009, involved fueling a 

grounded DC-8 with both 100-percent synthetic fuel and a 50/50 blend. 

The test results are being compared with baseline tests of hydrocarbon 

fuel emissions tests performed in the DC-8 in 2004. Air Force researchers were on hand to help measure the emissions.117

NASA and the Air Force are also working with Boeing to explore 

the possibility of using biofuel, which may prove to be cleaner than fuel 

derived from the Fischer-Tropsch process. The main obstacle to biofuel 

use at this time is the fact that it is difficult to procure in large quantities. 

For example, algae are attractive feedstock for biofuel, but the problem 

lies in being able to grow enough. NASA has begun to take on the feedstock problem by setting up a Greenlab Research Facility at NASA Glenn, 

where NASA researchers are seeking to optimize the growing conditions 

for algae and halophytes, which are plants tolerant of salt water.118 

In conclusion, the oil crisis and growing environmental awareness 

of the 1970s presented a critical opportunity for NASA to reclaim its 

mantle at the forefront of aeronautics research. NASA-led programs in 

fuel-efficient engines, aircraft structures, and composites—as well as 

the Agency’s contribution to computational fluid dynamics—planted the 

seeds that gave private industry the confidence and technological knowhow to pursue bold aircraft fuel-efficiency initiatives on its own. Without 

NASA’s E Cubed program, U.S. engine companies may not have had 

the financial resources to develop their fuel-saving, emissions-reducing 

TAPS and TALON combustors. E Cubed also spawned the open-rotor 

engine concept, which is still informing engine fuel-efficiency efforts 

today. The turbulent 1970s also created the opening for NASA Langley’s 

Richard Whitcomb to proceed full throttle with efforts to develop 

supercritical wings and winglets that have revolutionized fuel-efficient 

airframe design. And NASA’s research on alternative fuels during the 

1970s, if stillborn, nevertheless set the stage for the Agency to play a 

significant role in the Government’s revitalized alternative fuels research 

that came with the dawning of the 21st century.

117. Harrington, Bulzan interview.

118. Ibid.
NASA’s Contributions to Aeronautics

818

12

Addressing the Nation’s scientific leadership in 2009, President 

Barack Obama compared the energy challenge facing America to the 

shock of Sputnik in 1957, declaring it the Nation’s new “great project.”119 

Reflecting the increasing emphasis and rising priorities of Federal environmental research, NASA had received funding to support global climate 

studies, while NASA’s aeronautics research received additional funding to 

“improve aircraft performance while reducing noise, emissions, and fuel 

consumption.”120 Clearly, NASA’s experience in energy and aeronautics 

positioned the Agency well to continue playing a major role in these areas.

As the Agency enters the second decade of the 21st century, much 

remains to be done to increase aircraft fuel efficiency, but much, likewise, has already been accomplished. To NASA’s aeronautics researchers, inheritors of a legacy of accomplishment in flight, the energy and 

environmental challenges of the new century constitute an exciting stimulus, one as profoundly intriguing as any of the other challenges—supersonic flight and landing on the Moon among them—that the NACA and 

NASA have faced before. Those challenges, too, had appeared daunting. 

But just as creative NACA–NASA research overcame them, those in the 

Agency charged with responsibility for pursuing the energy and environmental challenges of the new century were confident that they, and 

the Agency, would once again see their efforts crowned with success.

119. President Barack Obama, Remarks by the President at the National Academy of Sciences 

Annual Meeting, (Washington, DC: Office of the Press Secretary, Apr. 27, 2009). Available at 

http://www.whitehouse.gov/the_press_office/Remarks-by-the-President-at-the-National-Academy-ofSciences-Annual-Meeting, accessed July 29, 2009.

120. The White House, “NASA Budget Highlights: The National Aeronautics and Space Administration 2010 Budget,” updated Apr. 28, 2009, available at http://www.nasa.gov/news/

budget/index.html, accessed July 29, 2009.
Case 12 | Leaner and Greener: Fuel Efficiency Takes Flight

819

12

Recommended Additional Readings

Reports, Papers, Articles, and Presentations:

Theodore G. Ayers and James B. Hallissy, “Historical Background and 

Design Evolution of the Transonic Aircraft Technology Supercritical 

Wing,” NASA TM-81356 (1981).

Darrell H. Baldwin and Bradford S. Linscott, “The Federal Wind Program 

at NASA Lewis Research Center,” NASA TM-83480 (1983).

R. Bradford, W.T. Atkins, R.M. Bass, R. Dascher, J. Dunkin, N. Luce, 

W. Seward, and D. Warren, “Coal Conversion and Synthetic-Fuel 

Production,” NASA Tech Briefs, vol. 5, no. 1 (Aug. 1, 1980), p. 56.

B.P. Collins, A.L. Haines, and C.J. Wales, “A Concept for a Fuel Efficient 

Flight Planning Aid for General Aviation,” NASA CR-3533 (1982).

Gerrard Cowan and Caitlin Harrington, “Adding Fuel to the Fire: Energy 

and Defence,” Jane’s Defence Weekly, Nov. 21, 2008.

D. Daggett, O. Hadaller, R Hendricks, and R. Walther, “Alternative Fuels 

and Their Potential Impact on Aviation,” NASA TM-2006-214365 (2006).

John H. Del Frate and Gary B. Cosentino, “Recent Flight Test Experience 

with Uninhabited Aerial Vehicles at the NASA Dryden Flight Research 

Center,” NASA TM-1998-206546 (1998).

Vance Dippold, III, “CFD Analyses and Jet-Noise Predictions of Chevron 

Nozzles with Vortex Stabilization,” NASA TM-2008-215150 (2008).

Edwin J. Graber, “Overview of NASA PTA Propfan Flight Test Program,” 

NASA Lewis [Glenn] Research Center, Nov. 1, 1987, CASI Doc 

198.800.06423.

Caitlin Harrington, “USAF Promotes Fuel Alternative,” Jane’s Defence 

Weekly, Mar. 16, 2007.

C.G. Jay, “Hydrogen Enrichment of Synthetic Fuel,” NASA Tech Briefs, 

vol. 3, no. 1 (June 1, 1978), NTTC Doc. 197.800.00039.
NASA’s Contributions to Aeronautics

820

12

B.H. Little, D.T. Poland, H.W. Bartel, C.C. Withers, and P.C. Brown, 

“Propfan Test Assessment (PTA) Final Project Report,” NASA 

CR-185138 (1989).

Valerie J. Lyons and Arlene S. Levine, “An Overview of NASA’s Contributions 

to Energy Technology,” NASA AIAA-2008-5641 (Jan. 2009).

Donald R. Mulholland, Loren W. Acker, Harold H. Christenson, and 

William V. Gough, “Flight Investigation of a Liquid-Hydrogen Fuel 

System,” RM E57F19a (1957).

A.L. Nagel, W.J. Alford, and J.F. Dugan, “Future Long-Range Transports: 

Prospects for Improved Fuel Efficiency,” NASA TM-X-72659 (1975).

Palmer C. Putnam, “Wind Power: Yesterday, Today, and Tomorrow,” 

NASA CP-2230 (1982).

N.A. Radovich, D. Dreim, D.A. O’Keefe, L. Linner, S.K. Pathak, J.S. Reaser, 

D. Richardson, J. Sweers, and F. Conner, “Study for the Optimization 

of a Transport Aircraft Wing for Maximum Fuel Efficiency,” vol. 1: 

“Methodology, Criteria, Aeroelastic Model Definition, and Results,” 

NASA CR-172551 (1985).

Travis L. Turner, Randolph H. Cabell, Roberto J. Cano, and Richard J. 

Silcox, “Testing of SMA-Enabled Active Chevron Prototypes Under 

Representative Flow Conditions,” NASA Langley Research Center, 

Report LAR-17332, Paper 6928-36, CASI Doc. 200.800.14174.

U.S. Federal Energy Administration, “Project Independence Blueprint. 

Task Force. Report. Solar Energy,” U.S. Department of Commerce, 

National Technical Information Service (1974).

U.S. Government Accountability Office, “Aviation’s Effects on the Global 

Atmosphere Are Potentially Significant and Expected to Grow,” GAO/

RCED-00-57 (Feb. 2000).

U.S. National Research Council [Committee on High Speed Research, 

Aeronautics and Space Engineering Board of the Commission on 
Case 12 | Leaner and Greener: Fuel Efficiency Takes Flight

821

12

Engineering and Technical Systems], U.S, Supersonic Aircraft: Assessing 

NASA’s High Speed Research Program (Washington: NRC, 1997).

Robert D. Wicofski, “Comparison of Alternate Fuel for Aircraft,” NASA 

TM-80155 (1979).

Books and Monographs:

Joseph R. Chambers, Partners in Freedom: Contributions of the Langley 

Research Center to U.S. Military Aircraft of the 1990s, SP-2000-4519 

(Washington: NASA, 2000).

Erick M. Conway, High Speed Dreams (Baltimore: Johns Hopkins 

University Press, 2005).

Virginia P. Dawson, Engines and Innovation: Lewis Laboratory and American 

Propulsion Technology, SP-4306 (Washington, DC: NASA, 1991).

Jeffrey Ethell, Fuel Economy in Aviation, NASA SP-462 (1983).

Alexander Laufer, et al., Shared Voyage: Learning and Unlearning from 

Remarkable Projects, NASA History Series, NASA SP-2005-4111, 

(Washington, DC: NASA, 2005).

Anita Liang, Emerging Fuel Cell Developments at NASA for Aircraft 

Applications (Cleveland: NASA Glenn Research Center, 2003).

Pamela Mack, ed., From Engineering Science to Big Science, NASA 

SP-4219 (Washington, DC: NASA, 1998).

Palmer C. Putnam, Power from the Wind (New York: Van Nostrand 

Reinhold Company, 1948).

Robert W. Righter, Wind Energy in America: A History (Norman, OK: 

University of Oklahoma Press, 1996).

John L. Sloop, Liquid Hydrogen as a Propulsion Fuel, NASA SP-4404 

(Washington. DC: NASA, 1978).
822

A full-pitch, fiberglass Vestas wind turbine in Rome in 2009. The Danish company remains the 

largest in the wind turbine industry. R.P. Hallion.
823

Good Stewards: NASA’s 

Role in Alternative Energy

Bruce I. Larrimer

CASE

13

Consistent with its responsibilities to exploit aeronautics technology for 

the benefit of the American people, NASA has pioneered the development and application of alternative energy sources. Its work is arguably most evident in wind energy and solar power for high-altitude 

remotely piloted vehicles. Here, NASA’s work in aerodynamics, solar 

power, lightweight structural design, and electronic flight controls has 

proven crucial to the evolution of novel aerospace craft.

T

HIS CASE STUDY REVIEWS two separate National Aeronautics and 

Space Administration (NASA) programs that each involved research 

and development (R&D) in the use of alternative energy. The first 

part of the case study covers NASA’s participation in the Federal Wind 

Energy Program from 1974 through 1988. NASA’s work in the wind 

energy area included design and fabrication of large horizontal-axis wind 

turbine (HAWT) generators, and the conduct of supporting research 

and technology projects. The second part of the case study reviews 

NASA’s development and testing of high-altitude, long-endurance solarpowered unmanned aerial vehicles (UAVs). This program, which ran from 

1994 through 2003, was part of the Agency’s Environmental Research 

and Aircraft Sensor Technology Program.

Wind Energy Program and Large Horizontal-Axis Wind Turbines 

(1974–1988)

The energy crisis of the 1970s brought about renewed interest in the 

development of alternative energy sources, including harnessing wind 

power for the generation of electricity. This renewed interest led to the 

establishment of the Federal Wind Energy Program in 1974 as part of the 

Nation’s solar energy program. The initial program overview, technical 

analysis, and objectives were formalized by the Project Independence 

Interagency Solar Task Force that was formed in April 1974 and 

chaired by the National Science Foundation (NSF). Approximately 100 

13
NASA’s Contributions to Aeronautics

824

13

individuals—representing various Government agencies, universities, 

research laboratories, private industries, and consulting firms—participated in the task force project. Thirteen of the participants were 

from NASA, including six from NASA Lewis (now NASA John H. 

Glenn Research Center). The task force’s final findings were outlined in 

the November 1974 “Project Independence Blueprint” report. The task 

force identified the six following “most promising” technologies for 

converting solar energy to a variety of useful energy forms: (1) solar 

heating and cooling of buildings, (2) solar thermal energy conversion, 

(3) wind energy conversion, (4) bioconversion to fuels, (5) ocean thermal energy conversion, and (6) photovoltaic electric power systems. 

The task force noted that the objective of the wind energy conversion 

part of the program was to improve the efficiency of wind turbine systems in a variety of applications and to reduce their costs. In regard to 

site selection, the task force concluded that the first attainment of economic viability in the United States would occur in areas such as the 

Great Plains, Alaska, the Great Lakes, the Atlantic and Pacific coasts, New 

England, and Hawaii. It concluded that the key to large-scale application 

of wind energy conversion systems was the reduction of costs through 

advanced technology, new materials, mass production, and the use of field 

fabrication techniques. Finally, the task force noted that a closely monitored program of proof-of-concept experiments was expected to reduce 

cost and constraint uncertainties.1

As a prelude to the formation of the wind energy program, NASA 

Lewis made significant contributions to a wind energy workshop that 

reviewed both the current status of wind energy and assessed the potential of wind power. This workshop was held as part of the Research 

Applied to National Needs (RANN) project that led to the National 

Science Foundation’s role in the initial planning of a 5-year sustained 

wind energy program. In January 1975, the wind energy program was 

transferred to the newly formed Energy Research and Development 

Administration (ERDA), which was incorporated into the newly formed 

U.S. Department of Energy (DOE) in 1977.

Pursuant to the initial agreement between NASA and the NSF, which 

had no research centers of its own, NASA’s Lewis Research Center at 

1. Federal Energy Administration, under direction of National Science Foundation, “Project Independence Blueprint,” Solar Energy PB 248 507 (reproduced by National Technical Information 

Service, U.S. Department of Commerce, Springfield, VA, Nov. 1974), pp. I-3, I-20-22.
Case 13 | Good Stewards: NASA’s Role in Alternative Energy

825

Lewis Field in Cleveland, OH, was given overall project management for 

the portion of the Wind Energy Program that involved the development 

and fabrication of large experimental horizontal-axis wind turbines. 

NASA Lewis’s responsibilities also included the conduct of supporting 

research and technology for the wind turbine conversion systems. This 

sponsorship continued under the Department of Energy once DOE took 

over the Federal Wind Energy Program. Louis Divone, who initially 

selected NASA Lewis to participate in the program, was the wind energy 

program manager for the NSF and later for ERDA and DOE. The program goal was the development of the technology for safe, reliable, and 

environmentally acceptable large wind turbine systems that could generate significant amounts of electricity at costs competitive with conventional electricity-generating systems.

NASA Lewis engineers were very interested in getting involved in 

the Wind Energy Program and realized early on that they could make 

significant contributions because of the Research Center’s long experience and expertise in propeller and power systems, aerodynamics, 

materials, and structures testing. The selection of NASA Lewis also represented an interesting historical context. Over 85 years earlier, in 1887–

1888, in Cleveland, OH, an engineer by the name of Charles F. Brush 

constructed a 60-foot, 80,000-pound wind-electric dynamo that is generally credited as being the first automatically operating wind turbine 

for electricity generation. Brush’s wind turbine, which supplied power 

for his home for up to 10 years, could produce a maximum 12,000 

watts of direct current that charged 12 batteries that in turn ran 350 

incandescent lights, 2 arc lights, and a number of electric motors. His 

dynamo made 50 revolutions to 1 revolution of the wind wheel, which 

consisted of 144 wooden blades and was 56 feet in diameter, accounting 

for 1,800 square feet of total blade surface swept area. The wind dynamo 

had an automatic regulator that prevented the power from running 

above 90 volts at any speed. Brush later dismantled his wind dynamo, 

apparently without attempting to develop a unit that could feed into a 

central power network.2

The use of wind power to generate electricity achieved a degree 

of success in rural and remote areas of the United States in the 1920s 

and 1930s. These generators, however, were small, stand-alone wind2. See Robert W. Righter, Wind Energy in America: A History (Norman, OK: University of Oklahoma Press, 1996), pp. 44–49 and 106.

13
826

13

NASA’s Contributions to Aeronautics

electric systems such as those designed and marketed by Marcellus and 

Joseph Jacobs, who built three-bladed systems, and the Windcharger 

Corporation, which built two-bladed generators. Most of these efforts 

were abandoned in the 1940s and 1950s because of the expansion of 

electrical utility networks, especially in response to passage of the Rural 

Electrification Act of 1937 and the availability of low-cost fossil fuels.

The first American effort to build a large wind turbine to feed into a 

power network was undertaken by Palmer Cosslett Putnam. This effort 

was funded by the S. Morgan Smith Company, which constructed and 

installed a 1.25-megawatt wind turbine at Grandpa’s Knob, VT. Prior to 

fabrication of his turbine, Putnam considered a number of questions 

that were still being debated years later, including whether to build a 

vertical- or horizontal-axis wind turbine; if horizontal, how many blades 

should there be; whether the generator should be aloft or on the ground; 

whether the drive should be mechanical or hydraulic; whether the tower 

should rotate or be stationary; and what size generator should be used. 

He noted that examples of all of these configurations existed in writings 

on wind power. Putnam, with the concurrence of both Beauchamp and 

Burwell Smith, decided on using the horizontal-axis, two-bladed stainless steel configuration, with a mechanically driven synchronous generator mounted aloft. He then concluded that the optimum size of a wind 

turbine generator (WTG) was close to 2 megawatts and noted that studies indicated that increased efficiency appeared to be flat between 2 and 

3 megawatts.3 The Smith-Putnam wind turbine, which supplied power to 

the Central Vermont Public Service Corporation’s power network, started 

operations on October 19, 1941, and operated intermittently for a total 

electric generation period of approximately 16 months. A bearing failure caused a blade separation accident, and the project was terminated 

in March 1945. While the turbine was not rebuilt, the system’s operation 

demonstrated that wind could be harnessed on a large scale to produce 

electricity. The power company, as well as others, envisioned that wind 

turbines would operate in conjunction with hydroelectric power systems.4

In the late 1950s, a German engineer by the name of Ulrich Hütter 

also built a smaller, 100-kilowatt wind turbine generator (the Hütter3. Palmer C. Putnam, “Wind Power: Yesterday, Today, and Tomorrow,” NASA CP-2230 (Proceedings of a workshop sponsored by U.S. Department of Energy, Wind Energy Technology Division, 

and NASA Lewis Research Center, Cleveland, OH, July 28–30, 1981).

4. Righter, Wind Energy in America, pp. 126–137.
Case 13 | Good Stewards: NASA’s Role in Alternative Energy

827

Allgaier wind turbine) that was tied into a power utility grid. Hütter’s 

machine used a 112-foot-diameter, two-bladed downwind rotor with full 

span pitch control. The blades were mounted on a teetered hub. In preparation to commence work on its own wind turbines, NASA purchased 

the plans from Hütter and considered or incorporated a number of design 

criteria and features of both the Smith-Putnam and Hütter-Allgaier wind 

turbine generators.5 NASA also participated in a joint NASA–Danish financing of the restoration of the wind turbine, which was completed in 1977. 

NASA Lewis later did aerodynamics testing and modeling of the Gedser 

wind turbine information using the Mod-0 testbed turbine.

NASA Lewis’s involvement in wind energy leading up to its selection to 

oversee the wind turbine development portion of the Federal Wind Energy 

Program included designing, at the request of Puerto Rico, a wind turbine to generate electricity for the Island of Culebra. This project grew out 

of an unrelated NASA Lewis 1972 project to take wind measurements in 

Puerto Rico. Later on, under the Wind Energy Program, NASA returned 

to Puerto Rico to build one of the Agency’s first-generation (Mod-0A) 

wind turbine machines. NASA Lewis’s involvement in the Wind Energy 

Program also was enhanced by its research of past wind energy projects 

and its projection of the future feasibility of using wind power to generate electricity for U.S. power utility networks. NASA’s overview and findings were presented as a paper at a symposium held in Washington, DC, 

that brought together past developers of wind turbines, including Palmer 

Putnam, Beauchamp Smith, Marcellus Jacobs, and Ulrich Hütter, as well 

as a new group of interested wind energy advocates.

1973 RANN Symposium Sponsored by the National Science Foundation

In reviewing the current status and potential of wind energy, Ronald 

Thomas and Joseph M. Savino, both from NASA’s Lewis Research Center, 

in November 1973 presented a paper at the Research Applied to National 

Needs Symposium in Washington, DC, sponsored by the National Science 

Foundation. The paper reviewed past experience with wind generators, 

problems to be overcome, the feasibility of wind power to help meet 

energy needs, and the planned Wind Energy Program. Thomas and 

Savino pointed out that the Dutch had used windmills for years to provide power for pumping water and grinding grain; that the Russians built 

5. “MOD-0A 200 Kilowatt Wind Turbine Generator Design and Analysis Report,” executive summary, DOE/NASA/0163-1; NASA CR-165127; AESD-TME-3051 (Aug. 1980), p. 2.

13
NASA’s Contributions to Aeronautics

828

13

a 100-kilowatt generator at Balaclava in 1931 that feed into a power network; that the Danes used wind as a major source of power for many 

years, including the building of the 200-kilowatt Gedser mill system that 

operated from 1957 through 1968; that the British built several large 

wind generators in the early 1950s; that the Smith-Putnam wind turbine built in Vermont in 1941 supplied power into a hydroelectric power 

grid; and that Germans did fine work in the 1950s and 1960s building 

and testing machines of 10 and 100 kilowatts. The two NASA engineers 

noted, however, that in 1973, no large wind turbines were in operation.

Thomas and Savino concluded that preliminary estimates indicated 

that wind could supply a significant amount of the Nation’s electricity 

needs and that utilizing energy from the wind was technically feasible, 

as evidenced by the past development of wind generators. They added, 

however, that a sustained development effort was needed to obtain economical systems. They noted that the effects of wind variability could 

be reduced by storage systems or connecting wind generators to fossil fuel or hydroelectric systems, or dispersing the generated electricity 

throughout a large grid system. Thomas and Savino6 recommended a 

number of steps that the NASA and National Science Foundation program should take, including: (1) designing, building, and testing modern 

machines for actual applications in order to provide baseline information 

for assessing the potential of wind energy as an electric power source, 

(2) operating wind generators in selected applications for determining 

actual power costs, and (3) identifying subsystems and components that 

might be further reduced in costs.7

NASA–Industry Wind Energy Program Large Horizontal-Axis Wind 

Turbines

The primary objective of the Federal Wind Energy Program and the 

specific objectives of NASA’s portion of the program were outlined in 

a followup technical paper presented in 1975 by Thomas, Savino, and 

Richard L. Puthoff. The paper noted that the overall objective of the 

6. In 2007, Purdue University recognized Joseph Savino’s work on wind energy and other accomplishments during his 41 years with NASA by awarding him its Outstanding Mechanical Engineer 

Award.

7. Ronald L. Thomas and Joseph M. Savino, “Status of Wind Energy Conversion,” NASA TMX-71523 (technical paper presented at RANN Symposium Sponsored by the National Science 

Foundation, Washington, DC, Nov. 18–20, 1973), pp. 1–5.
Case 13 | Good Stewards: NASA’s Role in Alternative Energy

829

program was “to develop the technology for practical cost-competitive 

wind-generator conversion systems that can be used for supplying significant amounts of energy to help meet the nation’s energy needs.”8

The specific objectives of NASA Lewis’s portion of the program were to: 

(1) identify cost-effective configurations and sizes of wind-conversion 

systems; (2) develop the technology needed to produce cost-effective, 

reliable systems; (3) design wind turbine generators that are compatible with user applications, especially with electric utility networks; (4) 

build up industry capability in the design and fabrication of wind turbine generators; and (5) transfer the technology from the program to 

industry for commercial application. To satisfy these objectives, NASA 

Lewis divided the development function into the three following areas: 

(1) design, fabrication, and testing of a 100-kilowatt experimental wind 

turbine generator; (2) optimizing the wind turbines for selected user 

operation; and (3) supporting research and technology for the systems.

The planned workload was divided further by assignment of different tasks to different NASA Research Centers and industry participants. NASA Lewis would provide project management and support in 

aerodynamics, instrumentation, structural dynamics, data reduction, 

machine design, facilities, and test operations. Other NASA Research 

Centers would provide consulting services within their areas of expertise. 

For example, Langley worked on aeroelasticity matters, Ames consulted 

on rotor dynamics, and Marshall provided meteorology support. Initial 

industry participants included Westinghouse, Lockheed Corporation, 

General Electric, Boeing, and Kaman Aerospace.

In order to undertake its project management role, NASA Lewis 

established the Center’s Wind Power Office, which consisted initially 

of three operational units—one covering the development of an experimental 100-kilowatt wind turbine, one handling the industry-built utility-operated wind turbines, and one providing supporting research and 

technology. The engineers in these offices basically worked together in a 

less formal structure, crossing over between various operational areas. 

Also, the internal organization apparently underwent several changes 

during the program’s existence. For example, in 1976, the program was 

8. R. Thomas, R. Puthoff, and J. Savino, “Plans and Status of the NASA-Lewis Research Center 

Wind Energy Project,” (technical paper to be presented at the Joint Power Conference cosponsored 

by the Institute of Electronic and Electrical Engineers and American Society of Mechanical Engineers, Portland, OR, Sept. 28–Oct. 1, 1975), p. 1.

13
NASA’s Contributions to Aeronautics

830

13

NASA Mod-0 testbed wind turbine, Plum Brook Station, Sandusky, OH. NASA.

directed by the Wind Power Office as part of the Solar Energy Branch. 

The first two office managers were Ronald Thomas and William Robbins. 

By 1982, the organization consisted of a Wind Energy Project Office, 

which was once again under the supervision of Thomas and was part of 

the Wind and Stationary Power Division. The office consisted of a project development and support section under the supervision of James P. 

Couch (who managed the Mod-2 project), a research and technology section headed by Patrick M. Finnegan, and a wind turbine analysis section 

under the direction of David A. Spera. By 1984, the program organization had changed again with the Wind Energy Project Office, which was 

under the supervision of Darrell H. Baldwin, becoming part of the Energy 

Technology Division. The office consisted of a technology section under 

Richard L. Puthoff and an analysis section headed by David A. Spera. The 

last NASA Lewis wind energy program manager was Arthur Birchenough.
Case 13 | Good Stewards: NASA’s Role in Alternative Energy

831

NASA’s Experimental (Mod-0) 100-Kilowatt Wind Turbine Generator 

(1975–1987)

Between 1974 and 1988, NASA Lewis led the U.S. program for large 

wind turbine development, which included the design and installation 

of 13 power-utility-size turbines. The 13 wind turbines included an initial testbed turbine designated the Mod-0 and 3 generations of followup 

wind turbines designated Mod-0A/Mod-1, Mod-2, and Mod-5. As noted in 

the Project Independence task force report, the initial 100-kilowatt wind 

turbine project and related supporting research was to be performed inhouse by NASA Lewis, while the remaining 100-kilowatt systems, megawatt systems, and large-scale multiunit systems subprograms were to 

be performed by contractors under NASA Lewis direction. Each successive generation of technology increased reliability and efficiency while 

reducing the cost of electricity. These advances were made by gaining a 

better understanding of the system-design drivers, improving the analytical design tools, verifying design methods with operating field data, and 

incorporating new technology and innovative designs. However, before 

these systems could be fabricated and installed, NASA Lewis needed to 

design and construct an experimental testbed wind turbine generator.

NASA’s first experimental wind turbine (the Mod-0) was constructed 

at Plum Brook Station in Sandusky, OH, and first achieved rated speed 

and power in December 1975. The initial design of the Mod-0 drew 

upon some of the previous information from the Smith-Putnam and 

Hütter-Allgaier turbines. The primary objectives of the Mod-0 wind turbine generator were to provide engineering data for future use as a base 

for the entire Federal Wind Energy Program and to serve as a testbed 

for the various components and subsystems, including the testing of 

different design concepts for blades, hubs, pitch-change mechanisms, 

system controls, and generators. Also, a very important function of 

the Mod-0 was to validate a number of computer models, codes, tools, 

and control algorithms.

The Mod-0 was an experimental 100-kilowatt wind turbine generator that at a wind speed of 18 miles per hour (mph) was expected to 

generate 180,000 kilowatthours of electricity per year in the form of 

440-volt, 3-phase, 60-cycle alternating current output. The initial testbed system, which included two metal blades that were each 62-feet long 

from hub to blade tip located downwind of the tower, was mounted on a 

100-foot, four-legged steel lattice (pinned truss design) tower. The drive 

train and rotor were in a nacelle with a centerline 100 feet above ground. 

13
NASA’s Contributions to Aeronautics

832

13

The blades, which were built by Lockheed and were based on NASA’s 

and Lockheed’s experience with airplane wing designs, were capable of 

pitch change (up and down movement) and full feather (angle of the 

blade change so that wind resistance is minimized). The hub was of the 

rigid type, meaning that the blades were bolted to the main shaft. A yaw 

(deviation from a straight path) control aligned the wind turbine with 

the wind direction, and pitch control was used for startup, shutdown, 

and power control functions. When the wind turbine was in a shutdown 

mode, the blades were feathered and free to slowly rotate. The system 

was linked to a public utility power network through an automatic synchronizer that converted direct current to alternating current.9

A number of lessons were learned from the Mod-0 testbed. One of 

the first problems involved the detection of larger than expected blade 

bending incidents that would have eventually caused early fatigue failure 

of the blades. The blade bending occurred for both the flatwise (out-ofplane) and edgewise (in-plane) blade positions. Followup study of this 

problem determined that high blade loads that resulted in the bending 

of the blades were caused by impulses applied to the blade each time it 

passed through the wake of the tower. Basically, the pinned truss design 

of the tower was blocking the airflow to a much greater degree than 

anticipated. The cause of this problem, which related to the flatwise 

load factors, was confirmed by site wind measurements and wind tunnel tower model tests. The initial measure taken to reduce the blocking 

effect was to remove the stairway from the tower. Eventually, however, 

NASA developed the soft tube style tower that later became the standard construction method for most wind turbine towers. Followup study 

of the edgewise blade loads determined that the problem was caused 

by excessive nacelle yawing (side-to-side) motion. This problem was 

addressed by replacing a single yaw drive, which aligns the rotor with 

the wind direction, with a dual yaw drive, and by adding three brakes 

to the yaw system to provide additional stiffness.10

Both of the above measures reduced the bending problems below 

the predicted level. Detection of these problems on the testbed Mod-0 

9. Richard L. Puthoff, “Fabrication and Assembly of the ERDA/NASA 100-Kilowatt Experimental 

Wind Turbine,” NASA TM-X-3390 (1976), pp. 1-9; and Thomas, Puthoff, and Savino, “Plans and 

Status of the NASA-Lewis Research Center Wind Energy Project,” pp. 3–4.

10. Ronald L. Thomas and Richard M. Donovon, “Large Wind Turbine Generators.” NASA TM73767 (1978), pp. 3–5.
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resulted in reevaluation of the analytical tools and the subsequent redesign of the wind turbine that proved extremely important in the design of 

NASA’s subsequent horizontal-axis large wind turbines. In regard to other 

operational testing of the Mod-0 system, NASA engineers determined that 

the wind turbine controls for speed, power, and yaw worked satisfactorily 

and that synchronization to the power utility network was successfully 

demonstrated. Also, the startup, utility operation, shutdown, and standby 

subsystems worked in a satisfactory manner. Finally, the Mod-0 was used 

to check out remote operation that was planned for future power utility 

systems. In summary, the Mod-0 project satisfied its primary objective 

of providing the entire Federal Wind Energy Program with early operations and performance data and through continued experience with 

testing new concepts and components. While NASA Lewis was ready to 

move forward with fabrication of its next level Mod-0A and Mod-1 wind 

turbines, the Mod-0 testbed continued to provide valuable testing of new 

configurations, components, and concepts for over 11 more years.

First Generation DOE–NASA Wind Turbine Systems (Mod-0A and Mod-1) 

(1977–1982)

The Mod-0 testbed wind turbine system was upgraded from 100 kilowatts to a 200-kilowatt system that became the Mod-0A. Installation of 

the first Mod-0A system was completed in November 1977, with one 

additional machine installed each year through 1980 at four locations: 

Clayton, NM; Culebra, PR; Block Island, RI; and Oahu, HI. This first 

generation of wind turbines completed its planned experimental operations in 1982 and was removed from service.

The basic components and systems of the Mod-0A consisted of the 

rotor- and pitch-change mechanism, drive train, nacelle equipment, 

yaw drive mechanism and brake, tower and foundation, electrical system and components, and control systems. The rotor consisted of the 

blades, hub, pitch-change mechanism, and hydraulic system. The drive 

train included the low-speed shaft, speed increaser, high-speed shaft, 

belt drive, fluid coupling, and rotor blades. The electrical system and 

components were the generator, switchgear, transformer, utility connection, and slip rings. The control systems were the blade pitch, yaw, 

generator control, and safety system.11

11. “MOD-0A 200 Kilowatt Wind Turbine Generator Design and Analysis Report,” executive summary, DOE/NASA/0163-1; NASA CR-165127; AESD-TME-3051 (Aug. 1980), p. 1.

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13

Similar to the Mod-0 testbed, the Mod-0A horizontal-axis machines 

had a 125-foot-diameter downwind rotor mounted on a 100-foot 

rigid pinned truss tower. However, this more powerful first generation of turbines had a rated power of 200 kilowatts at a wind speed of 

18 miles per hour and made 40 revolutions per minute. The turbine 

had two aluminum blades that were each 59.9 feet long. The 

Westinghouse Electric Corporation was selected, by competitive bidding, 

as the contractor for building the Mod-0A, and Lockheed was selected 

to design and build the blades. NASA and Westinghouse personnel 

were involved in the installation, site tests, and checkout of the wind 

turbine systems.

The primary goal of the Mod-0A wind turbine was to gain experience and obtain early operation performance data with horizontalaxis wind turbines in power utility environments, including resolving 

issues relating to power generation quality, and safety, and 

procedures for system startup, synchronization, and shutdown. This 

goal included demonstrating automatic operation of the turbine and 

assessing machine compatibility with utility power systems, as well as 

determining reliability and maintenance requirements. To accomplish 

this primary goal, small power utility companies or remote location 

sites were selected in order to study problems that might result from a 

significant percentage of power input into a power grid. NASA engineers 

also wanted to determine the reaction of the public and power utility 

companies to the operation of the turbines. The Mod-0A systems 

were online collectively for over 38,000 hours, generating over 3,600 

megawatthours of electricity into power utility networks. NASA determined that while some early reliability and rotor-blade life problems 

needed to be corrected, overall the Mod-0A wind turbine systems 

accomplished the engineering and research objectives of this phase of 

the program and made significant contributions to second- and 

third-generation machines that were to follow the Mod-0A and 

Mod-1 projects. Interface of the Mod-0A with the power utilities demonstrated satisfactory operating results during their initial tests from November 1977 to March 1978. The wind turbine was 

successfully synchronized to the utility network in an unattended 

mode. Also, dynamic blade loads during the initial operating period 

were in good agreement with the calculation using the MOSTAB 

computer code. Finally, successful testing on the Mod-0 provided 

the database that led the way for private development of a wide 
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range of small wind turbines that were placed in use during the 

late 1980s.12

Closely related to the Mod-0A turbine was the Mod-1 project, for 

which planning started in 1976, with installation of the machine taking 

place in May 1979. In addition to noise level and television interference 

testing (see below), the primary objective of the Mod-1 program was to 

demonstrate the feasibility of remote utility wind turbine control. Three 

technical assessments were planned to evaluate machine performance, 

interface with the power utility, and examine the effects on the environment. This system was a one-of-a-kind prototype that was much larger 

than the Mod-0A, with a rated power of 2,000 kilowatts (later reduced 

to 1,350) and a blade swept diameter of 200 feet. The Mod-1 was the 

largest wind turbine constructed up to that time. Considerable testing 

was done on the Mod-1 because the last experience with megawatt-size 

wind turbines was nearly 40 years earlier with the Smith-Putnam 1.25- 

megawatt machine, a very different design. Full-span blade pitch was 

used to control the rotor speed at a constant 35 revolutions per minute 

(later reduced to 23 rpm). The machine was mounted on a steel tubular 

truss tower that was 12 feet square at the top and 48 feet square at the 

bottom. General Electric was the prime contractor for designing, fabricating, and installing the Mod-1. The two steel blades were manufactured 

by the Boeing Engineering and Construction Company. There was also 

a set of composite rotor blades manufactured by the Kaman Aerospace 

Corporation that was fully compatible for testing on the Mod-1 system. 

The wind turbine, which was in Boone, NC, was tested with the Blue 

Ridge Electrical Membership Corporation from July 1979 to January 

1981. The machine, operating in fully automatic synchronized mode, 

fed into the power network within utility standards.13

One of the testing objectives of this first-generation prototype was 

to determine noise levels and any potential electromagnetic interference with microwave relay, radio, and television associated with 

12. Louis V. Divone, “Evolution of the Modern Wind Turbine,” ch. 3, in David A. Spera, ed., Wind 

Turbine Technology: Fundamental Concepts of Wind Turbine Engineering (New York: ASME Press, 

1994) pp. 116–117.

13. D.A. Spera, L.A. Viterna, T.R. Richards, and H.E. Neustadter, “Preliminary Analysis of Performance and Loads Data from the 2-Megawatt MOD-1 Wind Turbine Generator,” NASA TM-81408 

(1979), pp. 1–3; L.A. Viterna, “The NASA-LeRC Wind Turbine Sound Prediction Code,” presented 

at Second DOE/NASA Wind Turbine Dynamics Workshop, Cleveland, Feb. 24–26, 1981.

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13

mountainous terrain. These potential problems were among those 

identified by an initial study undertaken by NASA Lewis, General 

Electric, and the Solar Energy Research Institute. An analytical 

model developed at NASA Lewis of acoustic emissions from the rotor 

recommended that the rotor speed be reduced from 35 to 23 revolutions per minute, and the 2,000-kilowatt generator was replaced with 

a 1,350-kilowatt, 1,200-rpm generator. This change to the power train 

made a significant reduction in measured rotor noise. During the noise 

testing, however, the Mod-1, like the Mod-0A, experienced a failure in 

the low-speed shaft of the drive train and, because NASA engineers 

determined that both machines had accomplished their purposes, they 

were removed from the utility sites. Lessons learned from the engineering studies and testing of the first-generation wind turbine systems indicated the need for technological improvements to make the machines 

more acceptable for large utility applications. These lessons proved valuable in the design, construction, and operation of the next generation 

of DOE–NASA wind turbines. Other contributions from the Mod-1 program included low-cost wind turbine design concepts and metal and 

composite blade design and fabrication. Also, computer codes were 

verified for dynamic and loads analysis.

Although the Mod-1 was a one-of-kind prototype, there was a conceptual design that was designated as the Mod-1A. The conceptual design 

incorporated improvements identified during the Mod-1 project but, 

because of schedule and budget constraints, were not able to be used in 

fabrication of the Mod-1 machine. One of the improvements involved 

ideas to lessen the weight of the wind turbine. Also, one of the proposed 

configurations made use of a teetered hub and upwind blades with partial span control. Although the Mod-1A was not built, many of the ideas 

were incorporated into the second- and third-generation DOE–NASA 

wind turbines.

Second-Generation DOE–NASA Wind Turbine Systems (Mod-2)

While the primary objectives of the Mod-0, Mod-0A, and Mod-1 programs were research and development, the primary goal of the second-generation Mod-2 project was for direct and efficient commercial 

application. The Mod-2 program was designed to determine the potential cost-effectiveness of megawatt-sized remote site operation wind turbines when located in areas of moderate (14 mph) winds. Significant 

changes from the Mod-0 and Mod-1 included use of a soft-shell-type 
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DOE–NASA Mod-2 megawatt wind turbine cluster, Goldendale, WA. NASA.

tower, an epicyclic gearbox, a quill shaft to attenuate torque and power 

oscillations, and a rotor designed primarily to commercial steel fabrication standards. Other significant changes were the switch from fixed 

to a teetered (pivot connection) hub rotor, which reduced rotor fatigue, 

weight, and cost; use of tip control rather than full span control; and 

orienting the rotor upwind rather than downwind, which reduced rotor 

fatigue and resulted in a 2.5-percent increase in power produced by the 

system. Each of these changes resulted in a favorable decrease in the cost 

of electricity. One of the more important changes, as noted in a Boeing 

conference presentation, was the switch from the stiff truss type tower 

to a soft shell tower that weighed less, was much cheaper to fabricate, 

and enabled the use of heavy but economical and reliable rotor designs.14

Four primary Mod-2 wind turbine units were designed, built, and 

operated under the second-generation phase of the DOE–NASA program. The first three machines were built as a cluster at Goldendale, 

WA, where the Department of Energy selected the Bonneville Power 

14. Richard R. Douglas, “Large Wind Turbine Design Characteristics and R&D Requirements,” 

NASA CP-2106; DOE Publication CONF-7904111 (workshop held at Cleveland, OH, Apr. 

24–26, 1979), pp. 61–65.

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13

DOE–NASA Mod-2 wind turbine nacelle and rotor. NASA.

Administration as the participating utility. The operation of several wind 

turbines at one site afforded NASA the opportunity to study the effects 

of single and multiple wind turbines operating together while feeding 

into a power network. The Goldendale project demonstrated the successful operation of a cluster of large NASA Mod-2 horizontal-axis wind 

turbines operating in an unattended mode within a power grid. For construction of these machines, DOE–NASA awarded a competitively bid 

contract in 1977 to Boeing. The first of the three wind turbines started 

operation in November 1980, and the two additional machines went into 

service between March and May 1981. As of January 1985, the threeturbine cluster had generated over 5,100 megawatthours of electricity 

while synchronized to the power grid for over 4,100 hours. The Mod-2 

machines had a rated power of 2.5 megawatts, a rotor-blade diameter of 

300 feet, and a hub height (distance of the center of blade rotation to the 

ground) of 300 feet. Boeing evaluated a number of design options and 

tradeoffs, including upwind or downwind rotors, two- or three-bladed 
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839

rotors, teetered or rigid hubs, soft or rigid towers, and a number of 

different drive train and power generation configurations. A fourth 

2.5-megawatt Mod-2 wind turbine was purchased by the Department of 

the Interior, Bureau of Reclamation, for installation near Medicine Bow, 

WY, and a fifth turbine unit was purchased by Pacific Gas and Electric for 

operation in Solano County, CA.15

System Verification Units

In addition to the DOE–NASA units, NASA Lewis participated with 

the Bureau of Reclamation in the experimentation with two other turbines near Medicine Bow, WY. Both of these machines were designated 

as system verification units (SVU) because of their purpose of verifying the concept of integrating wind turbine generators with hydroelectric power networks. This was viewed as an important step in the 

Bureau of Reclamation’s long-range program of supplementing hydroelectric power generation with wind turbine power generation. One of 

the two turbines was a new design developed by the Hamilton Standard 

Division of United Technologies Corp., a 4-megawatt WTS-4 system, in 

the Medicine Bow area. A Swedish company, Karlskronavarvet (KKRV), 

was selected as a major subcontractor responsible for the design and 

fabrication of the nacelle hardware. The WTS-4 had a two-blade fiberglass downwind rotor that was 256.4 feet in diameter. For over 20 years, 

this 4-megawatt machine remained the largest power rated wind turbine 

generator ever built. In a reverse role, an additional 3-megawatt version 

of the same machine was built for the Swedish government, with KKRV 

as the prime contractor and Hamilton Standard as the subcontractor.16

The other SVU turbine was a Mod-2 design. While NASA engineers 

determined that the initial Mod-2 wind turbine generator performance 

was acceptable, they noted areas where improvement was needed. 

The problems encountered were primarily hardware-oriented and 

were attributed to fabrication or design deficiencies. Identification of 

these problems led to a number of modifications, including changes in 

the hydraulic, electric, and control systems; rework of the rotor hub 

flange; addition of a forced-lubrication system; and design of a new 

low-speed shaft.

15. Darrell H. Baldwin and Jerry Kennard, “Development of Large, Horizontal-Axis Wind Turbines,” 

NASA TM-86950 (1985), pp. 3–5 and Table 1.

16. Ibid., p. 4; and telephone interview of Larry Viterna by author, Apr. 27, 2009.

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13

DOE–NASA Mod-5B 3.2-megawatt wind turbine, Oahu, HI. NASA.

Third-Generation Advanced Multimegawatt Wind Turbines—The Mod-5 

Program (1980–1988)

The third-generation (Mod-5) program, which started in 1980, was 

intended to incorporate the experiences from the earlier DOE–NASA 

wind turbines, especially the Mod-2 experiences, into a final proof-ofconcept system for commercial use by an electric utility company. Two 

construction contracts were awarded to build the Mod-5 turbines—one 

unit to General Electric, which was designated the Mod-5A, and one 

unit to Boeing, which was designated the Mod-5B. As intermediate steps 

between the Mod-2 and Mod-5, two conceptual studies were undertaken 

for fabrication of both an advanced large wind turbine designated the 

Mod-3 and a medium turbine designated the Mod-4. Likewise, both a 

large-scale Mod-5 and medium-scale Mod-6 were planned as the final 

Wind Energy Program turbines. The Mod-3 and Mod-4 studies, however, 

were not carried through to construction of the turbines, and the Mod-6 

program was canceled because of budget constraints and changing priorities resulting from a decline in oil prices following the end of the oil 
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841

crisis of the 1970s. Also, General Electric chose not to proceed beyond 

the design phase with its Mod-5A. As a result, only the Boeing Mod-5B 

was constructed and placed into power utility service.17

Although its design was never built, General Electric did complete 

the detailed design work and all of the significant development tests and 

documented the entire Mod-5A program. The planned Mod-5A system 

contained many interesting features that NASA Lewis chose to preserve 

for future reference. The Mod-5A wind turbine was expected to generate 

electricity at a cost competitive with conventional forms of power generation once the turbines were in volume production. The program was 

divided into three phases: conceptual design, which was completed in 

March 1981; preliminary design, which was completed in May 1982; and 

final design, which was started in June 1982. The Mod-5A was planned 

to have a 7.3-megawatt generator, a 400-foot-diameter two-bladed teetered rotor, and hydraulically actuated ailerons over the outboard 40 

percent of the blade span to regulate the power and control shutdown. 

The blades were to be made of epoxy-bonded wood laminates. The yaw 

drive was to include a hydraulically actuated disk brake system, and the 

tower was to be a soft-designed welded steel plate cylindrical shell with 

a conical base. The Mod-5A was designed to operate in wind speeds of 

between 12 and 60 mph at hub height. The system was designed for automatic unattended operation and for a design life of 30 years.18

The Mod-5B, which was the only Mod-5 unit built, was physically 

the world’s largest wind turbine generator. The Mod-5B represented very 

advanced technology, including an upwind teetered rotor, compact planetary gearbox, pitchable tip blade control, soft-shell-type tower, and a 

variable-speed electrical induction generator/control system. Variable 

speed control enabled the turbine speed to vary with the wind speed, 

resulting in an increase energy capture and a decrease in fatigue loads 

on the drive train. The system underwent a number of design changes 

before the final fabricated version was built. For example, the turbine 

originally was planned to have a blade swept diameter of 304 feet. This 

was increased to 420 feet and finally reduced to 320 feet because of 

the use of blade steel tips and control improvements. Also, the turbine generator was planned initially to be rated at 4.4 megawatts. This 

17. Divone, “Evolution of Modern Wind Turbines,” ch. 3, in Spera, Wind Turbine Technology, p. 130.

18. General Electric Company, Advanced Energy Programs Department, “MOD-5A Wind Turbine 

Generator Program Design Report,” vol. 1, executive summary, NASA CR-174734 (1984), pp. 1–6.

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13

was increased to 7.2 megawatts and then decreased to the final version 

3.2 megawatts because of development of better tip control and load 

management. The rotor weighed 319,000 pounds and was mounted 

on a 200-foot tower. Extensive testing of the Mod-5B system was conducted, including 580 hours of operational testing and 660 hours of performance and structural testing. Performance testing alone generated 

over 72 reports reviewing test results and problems resolved.19

The Mod-5B was the first large-scale wind turbine to operate successfully at variable rotational speeds, which varied from 13 to 17.3 

revolutions per minute depending on the wind speed. In addition, the 

Mod-5B was the first large wind turbine with an apparent possibility of lasting 30 years. The turbine, with a total system weight of 1.3 

million pounds, was installed at Kahuku on the north shore of Oahu, 

HI, in 1987 and was operated first by Hawaiian Electric Incorporated 

and later by the Makani Uwila Power Corporation. The turbine started 

rated power rotation July 1, 1987. In January 1988, the Mod-5B was 

sold to the power utility, which continued to operate the unit as part 

of its power generation network until the small power utility ceased 

operations in 1996. In 1991, the Mod-5B produced a single wind turbine record of 1,256 megawatthours of electricity. The Mod-5B was operated in conjunction with 15 Westinghouse 600-kilowatt wind turbines. 

While the Westinghouse turbines were not part of the NASA program, 

the design of the turbines combined successful technology from NASA’s 

Mod-0A and Mod-2 programs.20

The Mod-5B, which represented a significant decrease over the Mod-2 

turbines in the cost of production of electricity, was designed for the 

sole purpose of providing electrical power for a major utility network. 

To achieve this goal, a number of changes were made over the Mod-2 

systems, including changes in concepts, size, and design refinements. 

These changes were reflected in more than 20 engineering studies, which 

addressed issues such as variable pitch versus fixed pitch, optimum 

machine size, steel shell versus truss tower, blade aerodynamics, material selection, rotor control, tower height, cluster optimization, and 

19. Boeing Aerospace Company, “MOD-5B Wind Turbine System Final Report,” vol. 1, executive 

summary, NASA CR-180896 (1988), p. 2–2, and vol. 3, “Acceptance Testing,” NASA CR180898 (1988), pp. v–vi, 1–1, 8–1, and B–1.

20. Divone, “Evolution of Modern Wind Turbines,” in Spera, Wind Turbine Technology, pp. 

130–131.
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gearbox configuration. For example, the studies indicated that loads 

problem was the decisive factor with regard to the use of a partial span 

variable pitch system rather than a fixed pitch rotor system, dynamic 

simulation led to selection of the variable speed generator, analysis 

of operational data enabled a significant reduction in the weight and 

size of the gearbox, and the development of weight and cost trend data 

for use in size optimization studies resulted in the formulation of 

machine sizing programs.21

A number of design elements resulted in significant contributions 

to the success of the Mod-5B wind turbine. Aerodynamic improvement 

over the Mod-2, including improvements in vortex generators, trailing 

edge tabs, and better shape control, resulted in an 18-percent energy 

capture increase. Improved variable speed design resulted in an increase 

of greater than 7 percent (up to as high as 11 percent) over an equivalent synchronous generator system. Both cycloconverter efficiency and 

control optimization of rotor speed versus wind speed proved to be better than anticipated. Use of the variable speed generator system to control power output directly, as opposed to the pitch power control on the 

Mod-2, substantially reduced blade activity, especially at below rated 

power levels. The variable speed design also resulted in a substantial 

reduction in structural loads. Adequate structural integrity was demonstrated for all stress measurement locations. Lessons learned during 

the earlier operation of the Mod-2 systems resulted in improved yaw 

and pitch systems. Extensive laboratory simulation of control hardware 

and software likewise reduced control problems compared with Mod-2 

systems.22 In summary, the Mod-5B machine represented a reliable 

proof-of-concept large horizontal-axis wind turbine conversion system 

capable of long-life production of electricity into a power grid system, 

thus fulfilling the DOE–NASA program objectives.

The Mod-5B was the last DOE–NASA wind turbine generator built 

under the Federal Wind Energy Program. In his paper on the Mod-5B 

wind turbine system, Boeing engineer R.R. Douglass noted the following size versus cost problem relating to the purchase of large wind turbines faced by power utility companies:

21. R.R. Douglas, Boeing Engineering and Construction Company, “Conceptual Design of the 7 

Megawatt Mod-5B Wind Turbine Generator Final Report,” N83-19272 (1982), pp. 1–7; and Boeing 

Aerospace Company, “Mod-5B Wind Turbine System Final Report,” vol. 3, pp. v, vi, 1–1, 8–1, and B–1.

22. “Mod-5B Wind Turbine Generator Final Report,” vol. 1, pp. 7–1 and 7–2.

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13

. . . large scale commercialization of large wind turbines suffers from the chicken and egg syndrome. That is, costs of units 

are so high when produced one or two at a time on prototype 

tooling that the utilities can scarcely afford to buy them. On 

the other hand, industry cannot possibly afford to invest the 

huge capital required for an automated high rate production 

capability without an established order base. To break this log 

jam will require a great deal of cooperation between government, industry, and the utilities.23

Boeing noted, however, in its final Mod-5B report that: “In summary 

the Mod-5B demonstrated the potential to generate at least 11 percent 

more revenue at a given site than the original design goal. It also demonstrated that multi-megawatt class wind turbines can be developed 

with high dependability which ultimately should show up in reduced 

operation and maintenance costs.”24

NASA’s Wind Turbine Supporting Research and Technology Contributions

A very significant NASA Lewis contribution to wind turbine development 

involved the Center’s Supporting Research and Technology (SR&T) program. The primary objectives of this component of NASA’s overall wind 

energy program were to gather and report new experimental data on various aspects of wind turbine operation and to provide more accurate analytical methods for predicting wind turbine operation and performance. 

The research and technology activity covered the four following areas: 

(1) aerodynamics, (2) structural dynamics and aeroelasticity, (3) composite materials, and (4) multiple wind turbine system interaction. In the 

area of aerodynamics, NASA testing indicated that rounded blade tips 

improved rotor performance as compared with square rotor tips, resulting in an increase in peak rotor efficiency by approximately 10 percent. 

Also in the aerodynamics area, significant improvements were made in 

the design and fabrication of the rotor blades. Early NASA rotor blades 

used standard airfoil shapes from the aircraft industry, but wind turbine 

rotors operated over a significantly wider range of angles of attack (angles 

between the centerline of the blade and incoming airstream). The rotor 

23. Douglass, “Conceptual Design of the 7 Megawatt Mod-5 Wind Turbine Generator,” p. 6.

24. Boeing Aerospace Company, “Mod-5B Wind Turbine System Final Report,” vol. 1, executive 

summary, NASA CR-180896 (Mar. 1988), p. 8–1.
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blades also needed to be designed to last up to 20 or 30 years, which 

represented a challenging problem because of the extremely high number of cyclic loads involved in operating wind turbines. To help solve 

these problems, NASA awarded development grants to the Ohio State 

University to design and wind tunnel test various blade models, and to 

the University of Wichita to wind tunnel test a rotor airfoil with ailerons.25

In the structural dynamics area, NASA was presented with problems related to wind loading conditions, including wind shear (variation of wind velocity with altitude), nonuniform wind gusts over the 

swept rotor area, and directional changes in the wind velocity vector field. NASA overcame this problem by developing a variable speed 

generator system that permitted the rotor speed to vary with the wind 

condition, thus producing constant power.

Development work on the blade component of the wind turbine 

systems, including selecting the material for fabrication of the blades, 

represents another example of supporting technology. As noted above, 

NASA Lewis brought considerable structural design expertise in this 

area to the wind energy program as a result of previous work on helicopter rotor blades. Early in the program, NASA tested blades made 

of steel, aluminum, and wood. For the 2-megawatt Mod-1 phase of the 

program, however, NASA Lewis decided to contract with the Kaman 

Aerospace Corporation for the design, manufacture, and groundtesting of two 100-foot fiberglass composite blades. NASA provided the 

general design parameters, as well as the static and fatigue load information, required for Kaman to complete the structural design of the 

blades. As noted in Kaman’s report on the project, the use of fiberglass, 

which later became the preferred material for most wind turbine blades, 

had a number of advantages, including nearly unlimited design flexibility in adopting optimum planform tapers, wall thickness taper, twist, 

and natural frequency control; resistance to corrosion and other environmental effects; low notch sensitivity with slow failure propagation 

rate; low television interference; and low cost potential because of adaptability to highly automated production methods.26

25. Darrell H. Baldwin and Bradford S. Linscott, “The Federal Wind Program at NASA Lewis 

Research Center,” NASA TM-83480 (1983), pp. 5–9.

26. W.R. Batesole, Kaman Aerospace Corporation, “Fiberglass Composite Blades for the 2 MW 

Mod-1 Wind Turbine Generator,” NASA CR-2230, Cleveland, OH, July 28–30, 1981, pp. 

215–226.

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13

The above efforts resulted in a significant number of technical 

reports, analytical tests and studies, and computer models based upon 

contributions of a number of NASA, university, and industry engineers 

and technicians. Many of the findings grew out of tests conducted on the 

Mod-0 testbed wind turbine at Plum Brook Station. One is work done 

by Larry A. Viterna, a senior NASA Lewis engineer working on the wind 

energy project, in aerodynamics. In studying wind turbine performance 

at high angles of attack, he developed a method (often referred to as the 

Viterna method or model) that is widely used throughout the wind turbine industry and is integrated into design codes that are available from 

the Department of Energy. The codes have been approved for worldwide 

certification of wind turbines. Tests with the Mod-0 and Gedser wind 

turbines formed the basis for his work on this analytical model, which, 

while not widely accepted at the time, later gained wide acceptance. 

Twenty-five years later, in 2006, NASA recognized Larry Viterna and Bob 

Corrigan, who assisted Viterna on data testing, with the Agency’s Space 

Act Award from the Inventions and Contributions Board.27

Benefits of NASA’s “Good Stewardship” Regarding the Agency’s 

Participation in the Federal Wind Energy Program

NASA Lewis’s involvement in the Federal Wind Energy Program 

from 1974 through 1988 brought a high degree of engineering experience and expertise to the project that had a lasting impact on the 

development of use of wind energy both in the United States and internationally. During this program, NASA developed the world’s first 

multimegawatt horizontal-axis wind turbines, the dominant wind 

turbine design in use throughout the world today.

NASA Lewis was able to make a quick start and contribution to 

the program because of the Research Center’s longstanding experience and expertise in aerodynamics, power systems, materials, and 

structures. The first task that NASA Lewis accomplished was to bring 

forward and document past efforts in wind turbine development, including work undertaken by Palmer Putnam (Smith-Putnam wind turbine), Ulrich Hütter (Hütter-Allgaier wind turbine), and the Danish 

27. Viterna and David C. Janetzke, “Theoretical and Experimental Power From Large HorizontalAxis Wind Turbines,” NASA TM-82944 (Sept. 1982); and “Wind Energy Reaps Rewards,” NASA 

release dated May 31, 2006, http://www.nasa.gov/vision/earth/technologies/wind_turbines_

prt.htm, accessed Aug. 20, 2009.
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Gedser mill. This information and database served both to get NASA 

Lewis involved in the Wind Energy Program and to form an initial data 

and experience foundation to build upon. Throughout the program, 

NASA Lewis continued to develop new concepts and testing and 

modeling techniques that gained wide use within the wind energy field. 

It documented the research and development efforts and made this 

information available for industry and others working on wind 

turbine development.

Lasting accomplishments from NASA’s program involvement 

included development of the soft shell tubular tower, variable speed 

asynchronous generators, structural dynamics, engineering modeling, design methods, and composite materials technology. NASA 

Lewis’s experience with aircraft propellers and helicopter rotors had 

quickly enabled the Research Center to develop and experiment with 

different blade designs, control systems, and materials. A significant 

blade development program advanced the use of steel, aluminum, 

wood epoxy composites, and later fiberglass composite blades that 

generally became the standard blade material. Finally, as presented 

in detail above, NASA was involved in the development, building, 

and testing of 13 large horizontal-axis wind turbines, with both the 

Mod-2 and Mod-5B turbines demonstrating the feasibility of operating large wind turbines in a power network environment. With the 

end of the energy crisis of the 1970s and the resulting end of most 

U.S. Government funding, the electric power market was unable to 

support the investment in the new large wind turbine technology. 

Development interest moved toward the construction and operation 

of smaller wind turbine generators for niche markets that could be 

supported where energy costs remained high.

NASA Lewis’s involvement in the wind energy program started 

winding down in the early 1980s, and, by 1988, the program was 

basically turned over to the Department of Energy. With the decline 

in energy prices, U.S. turbine builders generally left the business, 

leaving Denmark and other European nations to develop the 

commercial wind turbine market.

While NASA Lewis had developed a 4-megawatt wind turbine in 

1982, Denmark had developed systems with power levels only 10 percent of that at that time. However, with steady public policy and product development, Denmark had captured much of the $15 billion world 

market by 2004.

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13

TABLE 1

COMPARATIVE WIND TURBINE TECHNOLOGICAL DEVELOPMENT, 1981–2007

TURBINE TYPE Nibe A NASA WTS-4 Vestas

YEAR 1981 1982 2007

COUNTRY OF ORIGIN Denmark United States Denmark

POWER (IN KW) 630 4,000 1,800

TIP HEIGHT (FEET) 230 425 355

POWER REGULATION Partial pitch Full pitch Full pitch

BLADE NUMBER 3 2 3

BLADE MATERIAL Steel/fiberglass Fiberglass Fiberglass

TOWER STRUCTURE Concrete Steel tubular Steel tubular

Source: Larry A. Viterna, NASA.

Most of the technology developed by NASA, however, continued to 

represent a significant contribution to wind power generation applicable both to large and small wind turbine systems. In recent years, interest has been renewed in building larger-size wind turbines, and General 

Electric, which was involved in the DOE–NASA wind energy program, 

has now become the largest U.S. manufacture of wind power generators 

and, in 2007, was among world’s top three manufacturers of wind turbine systems. The Danish company Vestas remained the largest company 

in the wind turbine field. GE products currently include 1.5-, 2.5-, and, 

for offshore use, 3.6-megawatt systems. New companies, such as Clipper 

Wind Power, with its manufacturing plant in Cedar Rapids, IA, and Nordic 

Windpower also have entered the large turbine fabrication business in the 

United States. Clipper, which is a U.S.–U.K. company, installed its first 

system at Medicine Bow, WY, which was the location of a DOE–NASA 

Mod-2 unit. In the first quarter of 2007, the company installed eight commercial 2.5-megawatt Clipper Liberty machines. Nordic Windpower, 

which represents a merger of Swedish, U.S., and U.K. teams, markets 

its 1-megawatt unit that encompasses a two-bladed teetered rotor that 

evolved from the WTS-4 wind turbine under the NASA Lewis program.

In summary, NASA developed and made available to industry significant technology and turbine hardware designs through its “good 

stewardship” of wind energy development from 1974 through 1988. 

NASA thus played a leading role in the international development and 

utilization of wind power to help address the Nation’s energy needs 

today. In doing so, NASA Lewis fulfilled its primary wind program goal 
Case 13 | Good Stewards: NASA’s Role in Alternative Energy

849

of developing and transferring to industry the technology for safe, 

reliable, and environmentally acceptable large wind turbine systems 

capable of generating significant amounts electricity at cost competitive 

prices. In 2008, the United States achieved the No. 1 world ranking 

for total installed capacity of wind turbine systems for the generation 

of electricity.

Solar Propulsion for High-Altitude Long-Endurance Unmanned Aerial Vehicles

Another area of NASA involvement in the development and use of alternative energy was work on solar propulsion for High-Altitude LongEndurance (HALE) unmanned aerial vehicles (remotely piloted vehicles). 

Work in this area evolved out of the Agency’s Environmental Research 

Aircraft and Sensor Technology (ERAST) program that started in 1994. 

This program, which was a joint NASA/industry effort through a Joint 

Sponsored Research Agreement (JSRA), was under the direction of 

NASA’s Dryden Flight Research Center. The primary objectives of the 

ERAST program were to develop and transfer advanced technology to 

an emerging American unmanned aerial vehicle industry, and to conduct flight demonstrations of the new technologies in controlled environments to validate the capability of UAVs to undertake operational 

science missions. A related and important aspect of this mission was the 

development, miniaturization, and integration of special purpose sensors and imaging equipment for the solar-powered aircraft. These goals 

were in line with both the revolutionary vehicles development aspect of 

NASA’s Office of Aerospace Technology aeronautics blueprint and with 

NASA’s Earth Science Enterprise efforts to expand scientific knowledge 

of the Earth system using NASA’s unique capabilities from the standpoint of space, aircraft, and onsite platforms.28

Specific program objectives were to develop UAV capabilities for 

flying at extremely high altitudes and for long periods of time; demonstrate payload capabilities and sensors for atmospheric research; address 

and resolve UAV certification and operational issues; demonstrate the 

UAV’s usefulness to scientific, Government, and civil customers; and 

foster the emergence of a robust UAV industry in the United States.29

28. “ERAST: Environmental Research Aircraft and Sensor Technology,” NASA FS-2002-08-020 

DFRC, pp. 1–3.

29. Mishap Investigation Board, “Investigation of the Helios Prototype Aircraft Mishap,” vol. 1 

(Mishap Report), NASA (Jan. 2004), p. 16.

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13

 The ERAST program envisioned missions that included remote 

sensing for Earth science studies, hyperspectral imaging for agriculture 

monitoring, tracking of severe storms, and serving as telecommunications relay platforms. Related missions called for the development and 

testing of lightweight microminiaturized sensors, lightweight materials, avionics, aerodynamics, and other forms of propulsion suitable for 

extreme altitudes and flight duration.30

The ERAST program involved the development and testing of 

four generations of solar-powered UAVs, including the Pathfinder, the 

Pathfinder Plus, the Centurion, and the Helios Prototype. Because of 

budget limitations, the Helios Prototype was reconfigured in what could 

be considered a fifth-generation test vehicle for long-endurance flying 

(see below). Earlier UAVs, such as the Perseus, Theseus, and Proteus, 

relied on gasoline-powered engines. The first solar-powered UAV was the 

RAPTOR/Pathfinder, also known as the High-Altitude Solar (HALSOL) 

aircraft, which that was originally developed by the U.S. Ballistic Missile 

Defense Organization (BMDO—now the Missile Defense Agency) as part 

of a classified Government project and subsequently turned over to NASA 

for the ERAST program. In addition to BMDO’s interest in having NASA 

take over solar vehicle development, a workshop held in Truckee, CA, 

in 1989 played an important role in the origin of the ERAST program.

The Truckee Workshop and Conference Report

In July 1989, NASA Ames sponsored a workshop on requirements for 

the development and use of very high-altitude aircraft for atmospheric 

research. The primary objectives of the workshop were to assess the scientific justification for development of new aircraft that would support 

stratospheric research beyond the altitudes attainable by NASA’s ER-2 

aircraft and to determine the aircraft characteristics (ceiling, altitude, 

payload capabilities, range, flight duration, and operational capabilities) 

required to perform the stratospheric research missions. Approximately 

35 stratospheric scientists and aircraft design and operations experts 

attended the conference, either as participants or as observers. Nineteen 

of these attendees were from NASA (1 for NASA Langley, 16 from NASA 

Ames, and 2 representing both NASA Dryden and Ames); 4 were from universities and institutes, including Harvard University and Pennsylvania 

30. “Pathfinder Solar-Powered Aircraft” NASA, Dryden Flight Research Center Fact Sheet, FS-034.
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State University; and 6 represented aviation companies, including Boeing 

Aerospace, Aurora Flight Sciences and Lockheed. Crofton Farmer, representing the Jet Propulsion Laboratory, served as workshop chair, and 

Philip Russell, from NASA Ames, was the workshop organizer and report 

editor. The attendees represented a broad range of expertise, including 

9 aircraft and development experts, 3 aircraft operations representatives, 2 aeronautical science experts, 2 Earth science specialists, 1 instrument management expert (Steven Wegener from NASA Ames, who later 

directed the science and payload projects for the solar UAV program), 

1 general management observer, and 17 stratospheric scientists.31

The workshop considered pressing scientific questions that required 

advanced aircraft capabilities in order to accomplished a number of 

proposed science related missions, including: (1) answering important 

polar vortex questions, including determining what causes ozone loss 

above the dehydration region in Antarctica and to what extent the losses 

are transmitted to the middle latitudes; (2) determining high-altitude 

photochemistry in tropical and middle latitudes; (3) determining the 

impact and degree of airborne transport of certain chemicals; and (4) 

studying volcanic, stratospheric cloud/aerosol, greenhouse, and radiation balance. The workshop concluded that carrying out the above 

missions would require flights at a cruise altitude of 100,000 feet, the 

ability to make a round trip of between 5,000 and 6,000 nautical miles, 

the capability to fly into the polar night and over water more than 200 

nautical miles from land, and carry a payload equal to or greater than 

the ER-2. The workshop report noted that experience with satellites 

pointed out the need for increased emphasis on correlative measurements for current and future remote sensing systems. Previously, balloons had provided most of this information, but balloons presented a 

number of problems, including a low frequency of successful launches, 

the small number of available launch sites worldwide, the inability to follow selected paths, and the difficulty in recovering payloads. The workshop concluded with the following finding:

We recommend development of an aircraft with the capacity 

to carry integrated payloads similar to the ER-2 to significantly 

higher altitude preferably with greater range. It is important 

31. “Global Stratospheric Change: Requirements for a Very-High Altitude Aircraft for Atmospheric 

Research,” (workshop held in Truckee, CA, July 15–16, 1989), NASA CP-10041 (1989), p. xi.

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13

that the aircraft be able to operate over the ocean and in the 

polar night. This may dictate development of an autonomous 

or remotely piloted plane. There is a complementary need to 

explore strategies that would allow payloads of reduced weight 

to reach even higher altitude, enhancing the current capability of balloons.32

High-altitude, long-duration vehicle development, along with 

development of reduced weight instrumentation, both became goals 

of the ERAST program.

Joint Sponsored Research Agreement: “The ERAST Alliance”

The ERAST program was organized pursuant to a unique arrangement 

known as a Joint Sponsored Research Agreement (JSRA).33 This type 

of agreement was authorized by the National Aeronautics and Space 

Act of 1958, and the specific ERAST agreement was authorized under 

the NASA Administrator’s delegations of March 31, 1992, and February 

25, 1994. The purpose of the agreement was to: (1) develop and demonstrate UAV flight capability at altitudes up to 100,000 feet and up 

to 4 days’ duration; (2) further develop payload integration capabilities responsive to the data collection and measurement requirements 

of the atmospheric community; (3) research activity toward further 

resolution of UAV certification and civil operational issues; (4) 

further demonstrate UAV viability to scientific, Government, and civil 

users, leading to increased applications for UAVs; and (5) effect technology transfers to the parties to develop a robust United States UAV 

industry capable of asserting the lead as the premier provider of 

UAVs for government and civil users worldwide. The agreement, 

which became effective in September 1994, provided for the terms 

and conditions of the arrangement, various participation categories, 

32. “Global Stratospheric Change,” workshop held in Truckee, CA, p. 2.

33. The ERAST program was the first large NASA program to use a Joint Sponsored Research 

Agreement. The JSRA type arrangement was first researched as a new form of contracting option in 

the 1980s at NASA Ames. Ames had previously used this type of joint arrangement on three smaller 

projects. Jack Glazer, a NASA Ames attorney, started research in 1988 on the possibility of using the 

“Other Transaction Authority” clause in the Space Act of 1958 for application to collaborative projects. 

Glazer recommended creating AmTech as a nonprofit company to document JSRA legal and policy 

issues and to establish prototype projects. (See “The Promise of ERAST,” NASA Dryden, Oct. 2004.)
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preliminary budgets for the first 5 years, and operational and 

reporting requirements.34

The agreement established an Alliance Council, which was to meet 

at least twice a year, to coordinate with NASA Dryden’s ERAST Project 

Office on planning the research and development and flight-testing 

to be performed. The joint agreement provided for ERAST management through a NASA program manager and a NASA ERAST project 

manager. Required reports included an annual report, R&D/technical 

reports, monthly progress reports, intellectual property reports, commercialization reports, and management and financial reports. While 

actual program expenditures could and did vary, NASA’s projected financial commitments from 1994 through 2000 were $2.8 million for 1994, 

$5.75 million for 1995, $6.05 million for 1996, $6.35 million for 1997, 

$6.70 million for 1998, $7.25 million for 1999, and $7.25 million for 

2000. Finally, the terms and conditions of the agreement provided for 

extensions through December 31, 2000.35

The above arrangement, however, actually remained in effect after 

2000. In 2002, NASA entered into a followup joint agreement with 

AeroVironment, Inc., of Simi Valley, CA, including its SkyTower subsidiary. The new agreement was intended to streamline existing efforts 

to merge solar-powered UAV development into a single solar-electric platform program with the goal of developing multiple aircraft. This collaborative effort included continued development of the Helios Prototype.36

Industry partners that participated in the JSRA program included 

four primary companies—AeroVironment, Inc. (builder of the four solar 

prototype UAVs), Aurora Flight Sciences (manufacture of the Perseus 

B), General Atomics Aeronautical Systems (builder of the Altus 2), and 

Scaled Composites (developer of the Proteus). American Technology 

Alliances (AmTech) served as facilitator for the alliance, and Karen Risa 

Robbins, a founder of AmTech, played a primary role in development and 

acceptance of the ERAST JSRA. Of the above companies, AeroVironment 

was the primary one involved in the solar-powered part of the ERAST pro34. “Joint Sponsored Research Agreement: Environmental Research Aircraft & Sensor Technology: 

‘ERAST Alliance’ for High Altitude, Long Endurance Unmanned Aerial Vehicles,” (Office of Aeronautics, NASA, Aug. 1994), pp. 1 and 39.

35. Ibid. p. 44.

36. “NASA Dryden, AeroVironment Sign Solar/Electric Aircraft Pact,” NASA Dryden News 

Release 02-48 (Aug. 19, 2002).

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13

gram. There were up to 28 participants in the alliance, including small 

businesses, universities, and nonprofit organizations. NASA also worked 

closely with the Federal Aviation Administration to address a program 

goal of resolving issues related to operation of UAVs in the National 

Airspace System, including development of “see and avoid” sensors and 

“over-the-horizon” communications equipment. Under the joint agreement, NASA was able to provide program management and oversight, 

flight-test facilities, operational support, and project funding. The funding aspect of the joint NASA-industry effort was facilitated because the 

program was permitted to use Federal Acquisition Regulations as guidelines rather than as rules. Furthermore, NASA safety regulations were 

not required to be specifically followed. 37 As ERAST project manager, 

NASA Dryden was responsible for the setting of priorities, determination of technical approaches toward meeting project objectives, project funding and oversight, coordination of facilities for UAV operations, 

development and coordination of payloads for test flights, and foresight 

to ensure that actions taken by ERAST alliance partners satisfied NASA’s 

future needs for UAVs. Each company in the alliance made contributions 

to the project through combinations of money and services. The ERAST 

program, however, required only nominal funding by the companies, and, 

in order to further commercial development of HALE UAVs, NASA offered 

the companies ownership of all hardware developed by the program.38

Jenny Baer-Riedhart, NASA Dryden ERAST Program Manager for 

the first 4 years, described the NASA-industry working relationship 

under the joint agreement as follows:

NASA and the companies agreed on business plans at the 

annual alliance meeting. Each year at this meeting, I laid out 

the requirements for the program, based on input from all of 

the parties. Together, we evaluated our working business plan 

against these requirements. We set programmatic milestones, 

as well as milestones for each of the companies.39

37. Mishap Investigation Board, “Investigation of Helios Prototype Aircraft Mishap,” p. 16.

38. Alexander Laufer, Todd Post, Edward J. Hoffman, and Ronald A. Heifotz, Shared Voyage: 

Learning and Unlearning from Remarkable Projects (Washington, DC: NASA History Division, 

2005), ch. 4, “Flying High on Spirit: The Pathfinder Solar-Powered Airplane,” p. 125.

39. Ibid., p. 130.
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Baer-Riedhart added that NASA and the companies put funding into 

a shared bank account from which AmTech, acting as the go-between 

for the companies and NASA, distributed the funds to the parties. She 

noted that, at first, the companies wanted to get their own money, build 

their own aircraft, and have a flyoff, but that NASA’s vision from the 

start of the alliance was for the companies to get together to build one 

aircraft. Jeffrey Bauer, who was the Chief Engineer at Dryden and later 

served as the last ERAST Program Manager, credits the success of the 

program to the structure and partnerships that formed the alliance, noting: “One of the major attributes of the program is the alliance of government and industry. ERAST is not a contract. We work collectively 

to develop what’s best for the group and community.”40 John Del Frate, 

Dryden solar-power aircraft manager, commenting on the alliance, stated: 

“The technology early on was immature. We knew there would be problems, but the foundation of the program was built on the premise that 

we were allowed to take risks, and that made it very successful.”41 In 

addition to Baer-Riedhart, Del Frate, and Bauer, other NASA Dryden 

senior ERAST program/project managers included James Stewart, John 

Sharkey, and John Hicks.

Adding an industry perspective to the working relationship between 

the ERAST alliance partners, Ray Morgan, then vice president of 

AeroVironment, a company that had over 13 years of experience developing UAVs, noted: “Like most new relationships, the alliance went 

through an initial courtship phase, followed by a few spats, before it settled into an ongoing relationship that worked, more or less, for the good 

of all.” Morgan added that NASA brought considerable expertise to the 

program, including vast experience in developing and testing unique air 

vehicles at high altitudes.42

One area of NASA expertise that Morgan specifically noted was 

the advice that NASA provided AeroVironment regarding how best to 

implement redundant systems for critical components, especially where 

the systems must automatically determine which sensors are working 

properly and which ones are not. AeroVironment had used “single thread” systems across major components for the first Pathfinder 

prototype, meaning that failure of one component would likely cause 

40. Jay Levine, ed., “The X-Press,” vol. 44, issue 2, Dryden Flight Center (May 8, 2002), p. 3.

41. Levine, ed., “The X-Press,” vol. 45, issue 4, Dryden Flight Center (June 2003).

42. Laufer, et al., Shared Voyage, section 2, ch. 4, p. 129.

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13

failure of the UAV.43 The utilization of redundant systems also extended 

to other components of vehicle operation. For example, the control systems for the solar-powered UAVs were remotely piloted through a dual 

radio frequency data link with the vehicle’s automatic control system, 

which likewise achieved redundancy through the use of two identical 

flight computers, uplink receivers, and downlink transmitters. In addition, there was a triple set of airspeed sensors and dual Global Positioning 

System (GPS) receivers.44 Even the fuel cells were originally to be completely redundant, but this plan was abandoned because of budget limitations and fuel cell development problems. This need for redundant 

systems in UAVs was reinforced by NASA Dryden’s experience with testing UAVs, including a number of program mishaps. The NASA team realized that the chance of mission success was greatly improved through the 

use of redundant systems for the UAVs.

Pathfinder—First-Generation ERAST Solar Program Test Vehicle 

(1994–1997)

The first-generation ERAST solar-test program HALE vehicle was 

the Pathfinder, which was designed and built by AeroVironment. 

AeroVironment’s earlier solar aircraft projects included the building of 

the piloted Gossamer Albatross and a scaled-down version known as the 

Gossamer Penguin. This experience assisted the company in building 

the Pathfinder UAV. In addition to Ray Morgan, who was vice president 

of AeroVironment, the company team included a number of experienced 

engineers and technicians, including William Parks, who was the company’s chief engineer for the Centurion and Helios Prototype UAVs, and 

Robert Curtin and Kirk Flittie, who both served as project and later as 

program managers. The program brought honors for Bob Curtin and Ray 

Morgan, who both received the Aviation Week Laurel Award in 1996.45

Pathfinder, which initially was battery-powered, was a remotely 

piloted flying wing that demonstrated a number of technologies, including lightweight composite structures, low wing loading flying-wing 

configuration, redundant and fault tolerant flight control systems, 

43. Ibid., p.131.

44. Derek L. Lisoski and Mark B. Tischler, “Solar Powered Stratosphere Research Aircraft—Flight 

Test and Systems Identification,” presented at the RTO SCI Symposium held in Madrid, Spain, May 

5–7, 1998, and published in RTO MP-11, p. 27-4.

45. Laufer, et al., Shared Voyage, section 4, ch. 4, p. 224, and information from John Del Frate.
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Pathfinder aircraft in flight. NASA.

high-efficiency electric motors, thermal control systems for highaltitude flight, and a high specific power solar array. The remotely piloted 

Pathfinder had 6 electric motors that each weighed 13 pounds and consisted of a fixed-pitch 79-inch propeller and a solid-state motor with internal power electronics, nacelle, and cooling fins. Differential power to two 

wingtip motors on either side was used for lateral control. Wing dihedral 

(upsweep) provided roll stability, and 26 elevator control surfaces were 

attached to the wing’s trailing edge for pitch control. Pathfinder’s solar 

array generated approximately 8,000 watts near solar noon. The solar UAV 

could obtain an airspeed of between 15 and 25 mph and a cruising speed 

of between 17 to 20 mph. The vehicle had a length of 12 feet, a wingspan 

of 98.4 feet, a wing chord (front to rear distance) of 8 feet, a gross weight 

of approximately 560 pounds, a payload capacity of up to 100 pounds, a 

wing aspect ratio (the ratio between the wingspan and the wing chord) 

of 12 to 1, and a power-off glide ratio of 18 to 1. Pathfinder had a maximum bank rate of 5 degrees and a maximum turn rate of 3 degrees per 

second at sea level and 1.7 degrees at 60,000 feet.46

46. Lisoski and Tischler, “Solar Powered Stratospheric Research Aircraft—Flight Test and Systems 

Identification,” pp. 27-1 through 27-5; and “Pathfinder Solar-Powered Aircraft,” NASA Dryden 

Flight Research Center, FS-034.

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13

To gain some introductory understanding and experience with the 

challenges and nuances of solar cell operation, prior to the official start 

of the ERAST program and the transfer of Pathfinder from BMDO, the 

project team had arranged for some solar cell flight tests on local sorties 

over the Edwards dry lake. Pathfinder itself was not equipped with solar 

arrays on early test flights because the ERAST alliance did not want to 

risk any damage to the expensive solar cell arrays until Pathfinder’s flying capabilities could be tested. Pathfinder’s gross weight of 560 pounds 

produced a very low wing-loading load distribution of less than 0.64 

pounds per square foot that significantly increased sensitivity to winds 

during takeoff and landing. This necessitated special training for the 

ground controllers, especially during takeoff and landing of the UAV. 

Pathfinder’s first foray to high altitude took it to 50,500 feet and proved 

immensely productive. “We learned tons from that flight,” John Del Frate 

recalled afterward, noting, “There were a lot of naysayers that were quieted after that flight.”47 Unfortunately, afterward the vehicle dramatically demonstrated its sensitivity to wind, being seriously damaged in 

its hangar when ground crews opened both hangar doors, thus creating a draft that blew Pathfinder into a jet that was in the same hangar.

After a number of developmental flights and further modifications at 

Dryden, Pathfinder was transported to the Island of Kauai, which offered 

a more favorable wind environment and a greater operational area with 

less competing air traffic. From testing at Edwards Air Force Base, the 

NASA-industry alliance learned that weather factors—including wind, 

turbulence, cloud cover, humidity, temperature, and pressure—were critical in attempting to fly the wing-loaded Pathfinder at high altitudes. In 

addition, the team noted that the UAVs would probably not be flying in 

the same conditions as found in standard atmosphere reference tables 

because testing indicated a surprising variance in actual temperature 

in comparison with the tables. The team also noted that the higher a 

solar UAV flies, the greater the downwind drift distance if activation of 

a flight termination system (FTS) is required.48 These factors required 

careful study of historical weather patterns to determine the optimum 

47. Review comments of Del Frate to author, Aug. 4, 2009.

48. As Del Frate recalls: “Flight termination can vary from aircraft to aircraft. We chose to not 

cripple the aircraft if we had a problem, but rather we would cut power to the motors and use a 

small chute off-center to provide asymmetrical drag to put PF into a descending spiral in order to not 

destroy the aircraft. As a result, the wind aloft could carry us quite a ways downwind.”
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site to attempt to set a world record UAV altitude flight. Accordingly, 

NASA selected the Navy’s Pacific Missile Range Facility in Hawaii as the 

location to test the high-altitude capabilities of the solar UAVs.49

In Hawaii, Pathfinder was flown for seven additional flights, one of 

which in 1997 set a world record of over 71,500 feet for a high-altitude 

flight by a propeller-driven aircraft. This broke a 1995 Pathfinder test altitude flight record of 50,500 feet, which had earned NASA recognition as 

1 of the 10 most memorable record flights of 1995. Pathfinder test flights 

also demonstrated solar-powered HALE vehicles’ potential as platforms 

for environmental monitoring and technical demonstration missions by 

gaining additional information relating to the Island of Kauai’s terrestrial 

and coastal ecosystems. These science missions, which employed specially 

build lightweight sensor systems (see below), included detection of forest nutrient status, forest regrowth after damage from Hurricane Iniki in 

1992, sediment concentrations in coastal waters, and assessment of coral 

reef conditions. Experience from Pathfinder test flights, in combination 

with other UAV testing, also yielded a number of lessons regarding hardware reliability, including the following recommended procedures: (1) 

testing the airframe structure as much as possible before flight, particularly the composite airframe joint bondings; (2) testing the vehicle’s systems in an altitude chamber because of the extreme cold and low-pressure 

conditions encountered by high-altitude science aircraft; (3) recognizing 

that UAVs, like aircraft, have a tendency to gain weight; (4) maintaining 

strict configuration control; and (5) ensuring that a redundant system is 

functional before switching from the primary system.50

Pathfinder Plus—Second-Generation ERAST Program Test Vehicle 

(1997–1998)

NASA and AeroVironment upgraded the Pathfinder UAV to an improved 

configuration known as the Pathfinder Plus, which had a longer wingspan, two additional electric motors for a total of eight motors, and 

improved solar cells. The two additional motors were more efficient than 

the six carried forward from Pathfinder. The Pathfinder Plus UAV, which 

was intended to serve as a “transitional” aircraft between Pathfinder 

and the next-generation Centurion UAV, had a wingspan of 121 feet that 

49. Del Frate and Gary B. Cosentino, “Recent Flight Test Experience with Uninhabited Aerial 

Vehicles at NASA Dryden Flight Research Center,” NASA TM-1998-206546 (Apr. 1998), p. 5.

50. Ibid.

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Pathfinder Plus flight in Hawaii. NASA.

included use of four of the five wing sections from the original Pathfinder. 

The new center wing section contained more efficient solar cells that 

converted approximately 19 percent of the solar energy into electrical 

power for the vehicle’s motors, avionics, and communications system. 

This compared with an efficiency rating of only 14 percent for the solar 

cells on the remaining four wing sections from the original Pathfinder. 

The addition of the fifth wing section enabled Pathfinder Plus to generate 12,500 watts, as compared with the original Pathfinder’s 7,500 

watts. Pathfinder Plus had a gross weight of 700 pounds, up to a 150-

pound payload capacity, an aspect ratio of 15 to 1, a wing chord of 8 feet, 

and a power-off glide ratio of 21 to 1. Pathfinder Plus enabled higheraltitude flights and was used to qualify the next-generation Centurion 

wing panel structural design, airfoil, and SunPower Corporation’s solar 

array. Several flight tests were conducted in Hawaii. On its last flight, on 

August 6, 1998, Pathfinder Plus set a world altitude record of 80,201 feet 

for solar-power- and propeller-driven aircraft. These flight tests demonstrated the power, aerodynamic, and systems technologies needed for 

the Centurion and confirmed the practical utility of using high-altitude 

remotely controlled solar powered aircraft for commercial purposes.51

51. “NASA Dryden Fact Sheet—Pathfinder Solar–Powered Aircraft,” Dryden Flight Research Center, 

FS-034, pp. 2–3.
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ERAST Pathfinder Sensor Technology Development

As noted in the last two letters of the ERAST acronym, the development 

of sensor technology for the program constituted a major program goal. 

Science activities of the Pathfinder missions were coordinated by NASA’s 

Ames Research Center. Ames developed and tested a number of scientific 

instruments, including two imaging sensors—a high spectral resolution 

Digital Array Scanned Interferometer (DASI) and an Airborne Real-Time 

Imaging System (ARTIS). Steven Wegener of NASA Ames served as project manager for the science and sensor program. Dougal Maclise was payload project manager, Steven Dunagan was the team leader of the DASI 

project, and Stan Ault was team leader for the ARTIS project. DASI, which 

weighs less than 25 pounds and was mounted under Pathfinder’s wing, 

is a remote sensing instrument that looks at reflected spectral intensities from the Earth. The ARTIS payload was built around a color infrared six-megapixel digital camera. Both sensors were designed to be 

small, lightweight, and interactive in accordance with ERAST program 

goals of miniaturizing flight payloads. Both sensor systems also were 

designed to complement high-altitude studies of atmospheric ozone, landcover changes, and natural hazard studies conducted by NASA’s Earth 

Resources Survey ER-2 aircraft. The Pathfinder’s imaging systems featured a remote interactive operation and near-real-time transmission of 

images to ground stations and the Internet. This capability improved the 

speed, quality, and efficiency of data collection, analysis, and interpretation. The NASA team noted that the rapid availability of information from 

these systems could aid in fast decision making during natural disasters.52

The science and sensor aspects of the ERAST program promoted new 

solar UAV payloads and missions, including disaster management with 

the Global Disaster Information Network, over-the-horizon and real-time 

technologies, support of Earth science enterprises, high-resolution mapping, and promotion of the Commercial Remote Sensing program partnership.53 The advantages of UAVs over satellites and piloted aircraft include: 

(1) long-range capability, including the ability to fly to remote locations 

52. See, for example, “Pathfinder Solar-Powered Aircraft Begins Science Missions,” NASA Release: 

97-45, http://www.nasa.gov/centers/dryden/news/NewsReleases/1997/97-45.html,

accessed Aug. 20, 2009.

53. Stephen S. Wegener and James Brass, “Environmental Research Aircraft and Sensor Technology—New Technologies for Earth Science,” Earth Science Enterprise, Atmospheric Chemistry and 

Dynamics, p. 177.

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13

and cover large areas; (2) long-endurance capability, including the ability 

to fly longer and revisit areas on a frequent basis; (3) high-altitude capability, including the ability to fly above weather or danger; (4) slow-speed 

flight, including the capability to stay near one location; and (5) elimination of pilot exposure, thus enabling long duration or dangerous flights.54

Solar Cells and Fuel Cells for Solar-Powered ERAST Vehicles

NASA had first acquired solar cells from Spectralab but chose cells from 

SunPower Corporation of Sunnyvale, CA, for the ERAST UAVs. These 

photovoltaic cells converted sunlight directly into electricity and were 

lighter and more efficient than other commercially available solar cells at 

that time. Indeed, after NASA flew Helios, SunPower was selected to furnish high-efficiency solar concentrator cells for a NASA Dryden ground 

solar cell test installation, spring-boarding, as John Del Frate recalled 

subsequently, “from the technology developed on the PF+ and Helios 

solar cells.”55 The Dryden solar cell configuration consisted of two fixedangle solar arrays and one sun-tracking array that together generated up 

to 5 kilowatts of direct current. Field-testing at the Dryden site helped 

SunPower lower production costs of its solar cells and identify uses and 

performance of its cells that enabled the company to develop large-scale 

commercial applications, resulting in the mass-produced SunPower A-300 

series solar cells.56 SunPower’s solar cells were selected for use on the 

Pathfinders, Centurion, and Helios Prototype UAVs because of their highefficiency power recovery (more than 50-percent higher than other commercially available cells) and because of their light weight. The solar cells 

designed for the last generation of ERAST UAVs could convert about 19 

percent of the solar energy received into 35 kilowatts of electrical current 

at high noon on a summer day. The solar cells on the ERAST vehicles were 

bifacial, meaning that they could absorb sunlight on both sides of the cells, 

thus enabling the UAVs to catch sunrays reflected upward when flying 

above cloud covers, and were specially developed for use on the aircraft.

54. S.M. Schoenung and S.S. Wegener, “Meteorological and Remote Sensing Applications of 

High Altitude Unmaned Aerial Vehicles,” presented at the Fourth International Airborne Remote Sensing Conference and Exposition/21st Canadian Symposium on Remote Sensing, Ottawa, Ontario, 

Canada, June 21–24, 1999, sections 1.0 (“Introduction”) and 2.0 (“Applications”).

55. Del Frate, e-mail to author, Aug. 4, 2009.

56. See “Originating Technology/NASA Contribution,” pp. 66–67, Environment and Resources 

Management.
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While solar cell technology satisfied the propulsion problem during 

daylight hours, a critical problem relating to long-endurance backup systems remained to be solved for flying during periods of darkness. Without 

solving this problem, solar UAV flight would be limited to approximately 

14 hours in the summer (much less, of course, in the dark of winter), plus 

whatever additional time could be provided by the limited (up to 5 hours 

for the Pathfinder) backup batteries. Although significant improvements 

had been made, batteries failed to satisfy both the weight limitation and 

long duration power generation requirements for the solar-powered UAVs.

As an alternative to batteries, the ERAST alliance tested a number 

of different fuel cells and fuel cell power systems. An initial problem to 

overcome was how to develop lightweight fuel cells because only 440 

pounds of Helios’s takeoff weight of 1,600 pounds were originally planned 

to be allocated to a backup fuel cell power system. Helios required 

approximately 120 kilowatthours of energy to power the craft for up 

to 12 hours of flight during darkness, and, fortunately, the state of fuel 

cell technology had advanced far enough to permit attaining this; earlier efforts back to the early 1980s had been frustrated because fuel 

cell technology was not sufficiently developed at that time. The NASAindustry team later determined, as part of the ERAST program, that a 

hydrogen-oxygen regenerative fuel cell system (RFCS or regen system) 

was the hoped for solution to the problem, and substantial resources 

were committed to the project.

RFCSs are closed systems whereby some of the electrical power produced by the UAV’s solar array during daylight hours is sent to an electrolyzer that takes onboard water and disassociates the water into hydrogen 

gas and oxygen gas, both of which are stored in tanks aboard the vehicle. 

During periods of darkness, the stored gases are recombined in the fuel 

cell, which results in the production of electrical power and water. The 

power is used to maintain systems and altitude. The water is then stored 

for reuse the following day. This cycle theoretically would repeat on a 

24-hour basis for an indefinite time period. NASA and AeroVironment 

also considered, but did not use, a reversible regen system that instead 

of having an electrolyzer and a fuel cell used only a reversible fuel cell 

to do the work of both components.57

57. Michael A. Dornheim, “The Quest for Perpetual Flight: Special Fuel Cells Key to Months-long 

Flight,” Aviation Week & Space Technology (Feb. 28, 2000), pp. 58–61; telephone interview of 

Del Frate by author (Apr. 27, 2009), along with followup e-mail.

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As originally planned, Helios was to carry two separate regen fuel 

cell systems contained in two of four landing gear pods. This not only 

disbursed the weight over the flying wing, but also was in keeping with 

the plan for redundant systems. If one of the two fuel cells failed, Helios 

could still stay aloft for several days, albeit at a lower altitude. Contracts 

to make the fuel cell and electrolyzer were given to two companies—Giner 

of Waltham, MA, and Lynntech, Inc., of College Station, TX. Each of the 

two systems was planned to weigh 200 pounds, including 27 pounds for 

the fuel cell, 18 pounds for the electrolyzer, 40 pounds for oxygen and 

hydrogen tanks, and 45 pounds for water. The remaining 70 pounds consisted of plumbing, controls, and ancillary equipment.58

While the NASA–AeroVironment team made a substantial investment in the RFCS and successfully demonstrated a nearly closed system 

in ground tests, it decided that the system was not yet ready to satisfy the 

planned flight schedule. Because of these technical difficulties and time 

and budget deadlines, NASA and AeroVironment agreed in 2001 to switch 

to a consumable hydrogen-air primary fuel cell system for the Helios 

Prototype’s long-endurance ERAST mission. The fuel cells were already 

in development for the automotive industry. The hydrogen-air fuel cell 

system required Helios to carry its own supply of hydrogen. In periods of 

darkness, power for the UAV would be produced by combining gaseous 

hydrogen and air from the atmosphere in a fuel cell. Because of the low 

air density at high altitudes, a compressor needed to be added to the system. This system, however, would operate only until the hydrogen fuel 

was consumed, but the team thought that the system could still provide 

multiple days of operation and that an advanced version might be able to 

stay aloft for up to 14 days. The installation plan was likewise changed. 

The fuel cell was now placed in one pod with the hydrogen tanks attached 

to the lower surface of the wing near each wingtip. This modification, of 

course, dramatically changed Helios’s structural loadings, transforming 

it from a span-loaded flying wing to a point-loaded vehicle.59

Centurion—Third-Generation ERAST Program Test Vehicle (1996–1998)

The Centurion, which was built in 1998 by AeroVironment, represented 

third-generation advancement on the technology developed in the 

Pathfinder and Pathfinder Plus UAVs. The Centurion, however, was still 

58. Dornheim, “The Quest for Perpetual Flight,” Aviation Week & Space Technology, pp. 58–61.

59. Telephone interview of Del Frate by author (Apr. 2, 2009), and followup e-mail.
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Centurion in flight. NASA.

considered to be a prototype demonstrator for future solar-powered vehicles that could stay airborne for long periods of time. Prior to construction 

of the full-scale Centurion UAV, a quarter-scale model was constructed to 

verify design predictions. The Centurion had a five-section wingspan of 

206 feet, which was more than twice the wingspan of Pathfinder; incorporated a redesigned airfoil; and had a length of 12 feet and an aspect ratio 

of 26 to 1. The aircraft, which had 14 direct-current electric motors, was 

to be powered by bifacial solar cells that covered 80 percent of the upper 

wing surface and had a maximum output of 31 kilowatts. The Centurion 

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13

had a cruising speed of between 17 and 21 mph and could carry a payload of approximately 100 pounds to an altitude of 100,000 feet, or 600 

pounds to 80,000 feet. The primary building material consisted of carbon 

fiber and graphite epoxy composite structure, Kevlar, Styrofoam leading 

edge, and plastic film covering. Centurion flew three low-altitude developmental test flights on battery power at NASA Dryden, verifying handling 

qualities, performance, and structural integrity. The primary mission of 

the Centurion was to verify the handling and performance characteristics 

of an ultra-lightweight all-wing UAV with a wingspan of over 200 feet.60

Helios Prototype checkout flight in Hawaii. NASA.

Helios Prototype—Fourth-Generation and Last ERAST Unmanned Aerial 

Vehicle

The Helios Prototype, which resulted from modifications of the Centurion 

and renaming of the aircraft to Helios, was a proof-of-concept flying 

wing that because of budget limitations had two different configurations: 

a high-altitude configuration designated the HP-01 (1998–2002) and a 

long-endurance configuration designated the HP-03 (2003).61 The two 

primary objectives for the Helios Prototypes were to demonstrate sustained flight at an altitude of 100,000 feet and the ability to fly nonstop 

60. “Centurion,” NASA Dryden Flight Research Center Fact Sheet FS-056.

61. NASA never tested the Helios configuration (both HP-01 and HP-03) at Edwards, only testing 

Centurion. (Information from Del Frate, Aug. 4, 2009.)
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Helios Prototype taking off for record altitude flight. NASA. 

for at least 24 hours, including at least 14 hours above 50,000 feet. Initial 

low-altitude test flights were conducted under battery power at Edwards 

Air Force Base in 1999. Afterward, Helios was equipped with highefficiency photovoltaic solar cells and underwent high-altitude flighttesting in the summer of 2001 at the U.S. Navy Pacific Missile Range 

Facility in Hawaii. On August 13, 2001, following further upgrading of 

systems, the high-altitude configuration (HP-01) reached an altitude of 

96,863 feet, setting a world record for sustained horizontal flight by a 

winged aircraft and effectively satisfying the first Helios objective of highaltitude performance. The previous record was held by the Air Force’s 

SR-71A. The plan for the long-endurance configuration was to use solar 

cells to power the electric motors and subsystems during daylight hours 

and a modified hydrogen-air fuel cell system during the night. The vehicle also was equipped with backup lithium batteries and was batterypowered for its first six low-altitude test flights at Dryden.

Helios used wing dihedral (wing upsweep), engine power, elevator 

control surfaces, and a stability augmentation and control system to 

provide aerodynamic stability and control. Helios added a sixth wing 

panel to the Centurion, giving the remotely powered aircraft a wingspan 

of 247 feet, which was longer than the wingspans of either the U.S. Air 

Force C-5 transport or the Boeing 747 commercial airliner. The aircraft 

had a length of 12 feet, a wing chord of 8 feet, an empty weight of 1,322 

pounds, an aspect ratio of 31 to 1, and wing loading of 0.835 pounds 

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13

per square foot. The craft, which had 14 (only 10 on long-endurance 

HPO3 version) brushless direct-current electric motors each rated at 2 

horsepower (1.5 kW) that each drove a lightweight, 79-inch-diameter 

propeller, was powered by 62,120 bifacial solar cells covering the upper 

wing surface. Bifacial solar cells enabled Helios to convert solar energy 

into electricity when illuminated from either above or below, which 

enabled the vehicle to absorb reflected energy when flying above cloud 

cover. The Helios had a cruising speed of between 19 and 27 mph and 

could carry a payload of up to 726 pounds, including ballast, instruments, experiments, and a supplemental energy system. Helios was 

designed to operate at up to 100,000 feet, which is above 99 percent of 

the Earth’s atmosphere, with a typical endurance mission at 50,000 to 

70,000 feet. Like the Centurion, the Helios Prototype was constructed 

mostly of composite material, including carbon fiber, graphite epoxy, 

Kevlar, Styrofoam, and a thin transparent plastic film. The main tubular wing spar, which was made of carbon fiber and also wrapped with 

Nomex and Kevlar, was thicker on the top and bottom to absorb the 

constant bending motions that occurred during flight. Under-wing pods 

were attached to carry the landing gear, battery backup system, fuel cells, 

flight control computers, and data instruments. The Helios flight control surfaces consisted of 72 trailing-edge elevators that provided pitch 

control. The fixed landing gear consisted of two wheels on each pod.62

At an October 13, 1999, flight demonstration at the Dryden Flight 

Center, Ray Morgan—vice president of AeroVironment, which developed the Helios—noted that the ultimate intention is for Helios-type 

UAVs to “stay in the stratosphere for months at a time and act as an 

11-mile-high tower.” Morgan added that the Helios had a “very unique, 

slow, and stable flight characteristic . . . that means we can fly in fairly 

tight circles and be, essentially, a geo-stationary platform in the sky—

and that has lots of potential application.”63 Some of commercial tasks 

envisioned for future solar-powered HALE aircraft to follow the Helios 

Prototype were storm tracking studies, atmospheric sampling, spectral 

imaging for agricultural and natural resource purposes, pipeline mon62. “Helios Prototype,” Dryden Flight Research Center Fact Sheet FS-068.

63. John D. Hunley and Yvonne Kellogg, compilers, Proceedings of addresses, sessions and workshops of the NASA ERAST Exclusive Preview sponsored by NASA Dryden Flight Research Center, 

Edwards, CA, Oct. 13, 1999, ERAST: Scientific Applications and Technology Commercialization 

(Sept. 2000), p. 4.
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869

itoring, and telecommunications platforms. Morgan further noted that 

reliability of the Helios was obtained in two ways: simplicity, because 

the UAV is simply a flying wing, and redundancy. Each motor pylon 

turned 2 horsepower into 10 pounds of thrust. There was only one 

moving part in each pylon. There were no brushes, gearboxes, or mechanism for variation in the pitch required to operate the UAV.64 Furthermore, 

while the Helios had 72 elevators, the UAV could use differential thrust 

to turn the aircraft without use of the elevators.

The Helios HP03 long-endurance configuration, which included a 

hydrogen-air fuel cell pod, made its first high-altitude flight June 7, 2003. 

The test objectives of this flight were: (1) to demonstrate the readiness of 

the aircraft systems, flight support equipment and instrumentation, and 

flight procedures; (2) to validate the handling and aeroelastic stability 

of the aircraft; (3) to demonstrate the operation of the fuel cell system 

and gaseous hydrogen storage tanks; and (4) to provide training for personnel to staff future multiday flights. The flight was planned to be a 

30-hour endurance flight, but because of leakage in the coolant and compressed air line, the flight had to be aborted after 15 hours. Flight data, 

however, verified that the vehicle was aeroelastically stable under the flight 

conditions expected for the long-endurance flight demonstration.

On its second test flight, on June 26, 2003, following further modifications based on the first flight, the Helios prototype crashed on the 

Navy’s Pacific Missile Range Facility in Hawaii. This flight was intended 

to evaluate the vehicle’s hydrogen fuel cell system designed to power the 

UAV during periods of darkness, which would be necessary for longduration flights. Approximately 30 minutes into its flight and at an altitude of only 2,800 feet, the aircraft encountered turbulence and morphed 

into a high dihedral (wings upswept) configuration that, over time, took 

on an increasingly alarming U-like appearance. This caused Helios to 

become unstable in pitch, and the big UAV began to nose up and down. 

During one descending swoop, the big UAV exceeded its maximum permissible airspeed, causing the wing’s leading edge secondary structure on the outer panels to fail and the solar cells and skin on the upper 

surface to rip off. Shedding bits and pieces of structure, the Helios progressively disintegrated as it plunged downward, being destroyed when 

it impacted the ocean.65

64. Ibid., p. 6.

65. Mishap Investigation Board, “Investigation of the Helios Prototype Aircraft Mishap,” passim.

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13

The Mishap Investigation Board (MIB) concluded that: “In summary, there was no evidence that suggests that there was prior structural failure(s) that would have contributed to the vehicle developing 

the higher-than-normal dihedral deflections [wing upsweep] that caused 

the unstable pitch oscillation and subsequent failure of the vehicle.”66 

The report, in noting the lessons learned, identified both the proximate 

and root causes of the loss of Helios. The proximate cause of the crash 

was determined to be the “high dynamic pressure reached by the aircraft during the last cycle of the unstable pitch oscillation leading to failure of the vehicle’s secondary structure.” Two root causes, however, that 

contributed to or created the proximate cause of the accident were: (1) 

the lack of adequate analysis methods, which led to an inaccurate risk 

assessment of the effects of the configuration change from the HPO1 

(high altitude) to the HPO3 (long endurance) version of the Helios and 

(2) configuration changes to the aircraft, which altered the Helios from a 

spanloader to a highly point-loaded mass distribution on the same structure, significantly reducing design robustness and margins of safety.67

The report noted that the Helios UAV represented “a nonlinear stability and control problem involving complex interactions among the 

flexible structure, unsteady aerodynamics, flight control system, propulsion system, the environmental conditions, and vehicle flight dynamics.” 

As a result of the investigation, the MIB made a number of key recommendations for future UAV development. These included:

1. Develop more advanced, multidisciplinary “timedomain” analysis methods appropriate to highly flexible, “morphing” vehicles;

2. Develop ground-test procedures and techniques appropriate to UAV class vehicles to validate new analysis 

methods and predictions;

3. Improve the technical insight for highly complex 

projects, using the expertise available from all NASA 

Research Centers;

4. Develop multidisciplinary models, which can describe 

the nonlinear dynamic behavior of aircraft modifications 

or perform incremental flight-testing; and

66. Ibid., p. 45.

67. Ibid., pp. 10, 54, 87–89, and Table 10.1.
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5. Provide adequate resources to future programs for more 

incremental flight-testing when large configuration changes 

significantly deviate from the initial design concept.

As already noted, because of budget constraints, the NASA-industry 

team basically had to use one prototype vehicle for two tests: high altitude 

and long endurance. Adding to this factor, as noted in the MIB report, the 

switch from a regenerative fuel cell to a hydrogen-air fuel cell, which was 

necessitated by technical problems with the RFCS, combined with project 

time and budget deadlines, caused a critical change in the load factor on 

the wing structure. The hydrogen-air fuel cell system was more point 

loaded because of having to use a third pod at the centerline of the vehicle to hold the 520-pound primary hydrogen-air fuel cell. The originally 

planned for regenerative fuel cell system would have required only two pods, 

each one-third of the distance from the vehicle’s centerline to the wingtip.68

Even with the loss of the Helios Prototype, the MIB noted both the 

success and the challenge of the solar UAV program, as reflected in the 

following statement from the Mishap report:

During the course of this investigation the MIB discovered 

that the AV [AeroVironment]/NASA technical team had created 

most of the world’s knowledge in the area of High AltitudeLong Endurance (HALE) aircraft design, development, and 

test. This has placed the United States in a position of world 

leadership in this class of vehicle, which has significant strategic implications for the nation. The capability afforded by 

such vehicles is real and unique, and can enable the use of the 

stratosphere for many government and commercial applications. The MIB also found that this class of vehicle is orders 

of magnitude more complex than it appears but that the AV/

NASA technical team had identified and solved the toughest 

technical problems. Although more knowledge can and should 

be pursued as recommended in this report, an adequate knowledge base now exists to design, develop, and deploy operational HALE systems.69

68. Ibid., pp. 10, 89, and Table 10-1. Both HP-01 and HP-03 had a total of five under-wing 

pods; the HP-03 configuration replaced the original center pod with the fuel cell pod.

69. Ibid., p. 10.

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ERAST Program Overview and Accomplishments

The ERAST program, which started in 1994 and ended in 2003, following the loss of the Helios Prototype, accomplished almost all of its 

primary objectives, including development of a solar-powered UAV 

capable of flying at very high altitudes; demonstrating payload capabilities and sensors for atmospheric research; resolving solar-powered 

UAV operational issues; and demonstrating UAV’s usefulness to scientific, Government, and civil customers. The ERAST program also demonstrated that the unique joint NASA-industry ERAST alliance under a 

Joint Sponsored Research Agreement worked well and that good cooperation under this agreement led to efficient use of resources and expedited reaching the project milestones required to satisfy the program’s 

budget and time constraints. In regard to the alliance partner manufacturing the solar cells for the four generations of solar-powered UAVs, 

SunPower Corporation was able to make significant increases in efficiency and lower the cost of its commercial solar cells, which led to its 

successful mass-produced series A-300 solar cells. The one primary objective that remained unfulfilled at the close of the ERAST NASA-industry 

program was obtaining the long-endurance capability to fly for multiple days or longer. As with the earlier solar UAV program in the 1980s, 

a backup power system that would enable solar UAVs to fly in periods of darkness remained the critical problem to be solved. Significant 

progress in this area, however, was made and work continued at NASA 

Glenn under the Low Emissions Alternative Power (LEAP) program on 

the goal of perfecting a lightweight regenerative fuel cell system. NASA 

Glenn made its first closed-loop (system completely sealed) regenerative fuel cell demonstration in September 2003 and demonstrated five 

contiguous back-to-back charge-discharge cycles at full power under 

semiautonomous control in July 2005.70

70. D.J. Bents, “Solar Airplanes and Regenerative Fuel Cells,” presentation to the 43rd annual 

I.R.I.S. show in Mayfield Heights, OH, Oct. 9, 2007.
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Recommended Additional Readings

Reports, Papers, Articles, and Presentations:

T.S. Anderson, et al., “MOD-0A 200-Kilowatt Wind Turbine Generator 

Design and Analysis Report,” executive summary, NASA CR-165127 

(1980).

Darrell H. Baldwin and Jerry Kennard, “Development of Large, 

Horizontal-Axis Wind Turbines,” NASA TM-86950 (1985).

Darrell H. Baldwin and Bradford S. Linscott, “The Federal Wind Program 

at NASA Lewis Research Center,” NASA TM-83480 (1983).

A.G. Birchenough, et al., “Operating Experience with the 200 KW 

MOD-0A Wind Turbine Generators,” NASA Lewis Wind Energy 

Project Office, NASA 83N-15916 (1983).

Boeing Aerospace Company, “MOD-5B Wind Turbine System Final 

Report,” vol. 1 (executive summary), NASA CR-180896 (1988), and 

vol. 3 (“Acceptance Testing”) NASA CR-180898 (1988).

John H. Del Frate and Gary B. Cosentino, “Recent Flight Test Experience 

with Uninhabited Aerial Vehicles at the NASA Dryden Flight Research 

Center,” NASA TM-1998-206546 (1998).

Michael A. Dornheim, “AeroVironment Pushes Limits of Solar Flight,” 

Aviation Week & Space Technology (May 4, 1998).

Michael A. Dornheim, “Giant Centurion Starts Flight Test,” Aviation 

Week & Space Technology (Nov. 16, 1998).

Richard R. Douglas, “The Boeing MOD-2 at 2.5 MW,” workshop held 

at Lewis Research Center, Cleveland, OH, Apr. 24–26, 1979, NASA 

CP-2106 (1979).

R.R. Douglas, “Conceptual Design of the 7 Megawatt MOD-5B Wind 

Turbine Generator,” NASA Accession ID: 83N19272 (1982).

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13

Dryden Flight Research Center Fact Sheets, “Centurion,” NASA FS-056.

Dryden Flight Research Center Fact Sheets, “Helios Prototype,” NASA 

FS-068.

Dryden Flight Research Center Fact Sheets, “Pathfinder Solar-Powered 

Aircraft,” NASA FS-034 and FS-068.

Dryden Flight Research Center Fact Sheets, “ERAST,” NASA FS-020.

Federal Energy Administration, “Project Independence Blueprint. 

Task Force. Report. Solar Energy,” U.S. Department of Commerce, 

National Technical Information Service (1974).

General Electric Company, Space Division, “Mod-1 Wind Turbine 

Generator Analysis and Design Report,” NASA CR-159497 (1979).

General Electric Company, “MOD-5A Wind Turbine Generator Program 

Design Report,” vol. 1 (executive summary), NASA CR-174734 (1984).

John D. Hunley and Yvonne Kellogg, compilers, “ERAST: Scientific 

Applications and Technology Commercialization,” Proceedings of the 

NASA ERAST Exclusive Preview sponsored by NASA Dryden Flight 

Research Center, Edwards, CA, Oct. 13, 1999, NASA CP-2000-209031 

(2000).

Jay Levine, ed., “Special Helios Prototype Edition,” The X-Press, vol. 44, 

issue 2, Dryden Flight Center, Edwards, CA (May 8, 2002).

Derek L. Lisoski and Mark B. Tischler, “Solar Powered Stratospheric 

Research Aircraft—Flight Test and System Identifications,” paper 

presented at the RTO SCI Symposium held in Madrid, Spain, and 

published in RTO MP-11.

NASA, “Global Stratospheric Change: Requirements for a Very-High 

Altitude Aircraft for Atmospheric Research,” NASA CP-10041 (1989).

NASA Facts, “ERAST: Environmental Research Aircraft and Sensor 

Technology,” FS-2002-08-020 DFRC (2002).
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875

Thomas E. Noll, et al., (Mishap Investigation Board) “Investigation of 

the Helios Prototype Aircraft Mishap,” vol. 1, NASA (2004).

Richard L. Puthoff, “Fabrication and Assembly of the ERDA/NASA 

100-Kilowatt Experimental Wind Turbine,” NASA TM-X-3390 (1976).

Richard L. Puthoff, et al. “Installation and Checkout of the DOE/NASA 

MOD-1 2000-kW Wind Turbine Generator,” NASA TM-81444 (1980).

Palmer C. Putnam, “Wind Power: Yesterday, Today, and Tomorrow,” 

Proceedings of a workshop sponsored by the U.S. Department of Energy 

and NASA Lewis Research Center held in Cleveland, OH, July 28–30, 

1981, NASA CP-2230 (1982).

S.M. Schoenung and S.S. Wegener, “Meteorological and Remote Sensing 

Applications of High Altitude Unmanned Aerial Vehicles,” paper 

presented at the Fourth International Airborne Remote Sensing 

Conference and Exhibition/21st Canadian Symposium on Remote 

Sensing, Ottawa, Ontario, Canada (June 21–24, 1999).

D.A. Spera, et al., “Preliminary Analysis of Performance and Loads Data 

From the 2-Megawatt Mod-1 Wind Turbine Generator,” paper prepared for the Fourth Biennial Conference and Workshop on Wind 

Energy Conversion Systems, sponsored by the Department of Energy, 

Washington, DC, Oct. 29–31, 1979, NASA TM-81408 (1979).

Edward H. Teets, Jr., et al., “Atmospheric Consideration for Uninhabited 

Aerial Vehicle (UAV) Flight Test Planning,” NASA TM-1998-206541 

(1998).

Ronald Thomas, et al., “Plans and Status of the NASA-Lewis Research 

Center Wind Energy Project,” NASA TM-X-71701 (1975).

Ronald Thomas and Richard M. Donovon, “Large Wind Turbine 

Generators,” NASA TM-73767 (1978).

Ronald Thomas and Joseph M. Savino, “Status of Wind-Energy 

Conversion,” NASA TM-X-71523 (1973).

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13

Larry A. Viterna, “The NASA-LeRC Wind Turbine Sound Prediction 

Code,” presented at Second DOE/NASA Wind Turbine Dynamics 

Workshop, Cleveland (1981).

Larry A. Viterna and David C. Janetzke, “Theoretical and Experimental 

Power From Large Horizontal-Axis Wind Turbines,” NASA TM-82944 

(1982).

Books and Monographs:

Virginia P. Dawson, Engines and Innovation: Lewis Laboratory and 

American Propulsion Technology, ch. 10, NASA History Office, 

SP-4306 (Washington, DC: NASA, 1991).

Richard P. Hallion and Michael H. Gorn, On the Frontier: Experimental 

Flight at NASA Dryden (Washington, DC: Smithsonian Institution 

Press, 2004).

Alexander Laufer, et al., Shared Voyage: Learning and Unlearning from 

Remarkable Projects, NASA History Series, NASA SP-2005-4111 

(Washington, DC: NASA, 2005).

Palmer C. Putnam, Power from the Wind (New York: Van Nostrand 

Reinhold Company, 1948).

Robert W. Righter, Wind Energy in America: A History (Norman, OK: 

University of Oklahoma Press, 1996).

David A. Spera, ed., Wind Turbine Technology: Fundamental Concepts of 

Wind Turbine Engineering (New York: ASME Press, 1994).
THIS PAGE INTENTIONALLY BLANK
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879

Index

201 Aircraft, 23n39

202 Aircraft, 23n39

367-80 Aircraft, 20

500E Helicopter, 165

707 Aircraft, 4, 20, 65, 123, 740, 804

727 Aircraft, 737, 745, 802, 804

737 Aircraft, 124, 616, 737, 745, 771

747 Aircraft: 747-400 winglet use, 89, 123–

24; composite materials, 804; engine 

component improvements, 745; NASA’s 

aeronautics role and, 787; Shuttle carrier 

aircraft, 661; simulators testing of PCA, 

715; TAVs launched from, 337; titanium 

construction, 616; winglet use on, 89, 122, 

123–24

757 Aircraft, 124, 715, 756, 795, 807–8

767 Aircraft, 124, 615

777 Aircraft: 777-300ER aircraft, 795; fly-bywire systems, 722, 753; GE90 jet engine, 

742, 743, 747; materials for, 620; noise 

from, 771; supercritical wing, 115; titanium construction, 616

787 Aircraft: electrical flight control surfaces and actuators, 722; engine for, 769, 

795; materials for, 571, 620; noise from, 

771; range of, 620; titanium construction, 616

A

A&P Technology, Inc., 768, 769

A-2 rocket, 635–36

A4D-1 Skyhawk, 45–46

A-4 Skyhawk, 65

A-12 Blackbird, 210, 570

A-17A aircraft, 734

Abbett, Mike, 444

Aberdeen Proving Ground, Maryland, 14

Ablation, heat of: measurement of, 305, 581; 

satellites, conditions for, 305, 581

Ablation cooling, 286–88, 417–18, 581–82, 

591, 606–8

Accelerometers, 367, 376, 534

Ackeret, Jakob, 93, 182

Ackeret theory, 15

Acoustically Treated Nacelle program, 740

Acoustical Society of America (ASA): engine 

noise research, 736–37; research results, 

dissemination of through, 207; sonic 

boom conferences, 211, 217

Active cooling concepts, 420, 589–90, 605

Active-element computers, 468

Active flight control, 651–53, 805. See also

Control Configured Vehicle (CCV)

Active laminar flow control, 806–8

Active Oscillation Suppression (AOS) system, 389

ACTOMP program, 556

ACTON program, 556

Actuators, 384–85

Adams, Michael J., 299, 373

Adaptive Engine Control System (ADECS) 

program, 711–12, 757

Adaptive flight control system (AFCS), 372–

74, 375, 376, 379

Adaptive grid, 434

Ad Hoc Group for Structural Analysis, 474

Advanced Concepts Simulator, 715

Advanced Control Technology for Integrated 

Vehicles Experiment (ACTIVE) project, 

258, 716–17

Advanced Fighter Technology Integration 

(AFTI) F-16 project: aircraft modifications 
NASA’s Contributions to Aeronautics

880

for, 672–73, 675; asynchronous digital 

flight control system, 672, 680–82, 683; 

Automated Maneuvering Attack System 

(AMAS), 672, 673, 674–75, 676; benefits of 

program, 632; control laws, 680–82; digital 

flight control system, 672–73, 674, 680–82; 

F-16 CCV as basis for, 648, 670, 672; flighttesting, 673–79; forward-looking infrared 

(FLIR) device, 673, 675; goals of, 672; 

ground collision avoidance system, 675; 

helmet-mounted target designation sight, 

673, 675; Pilot Activated Recovery System, 676; power-by-wire testbed, 676–79; 

publication of technical papers on, 679–

80; research activities, 630; Stores Management System (SMS), 672, 681; strike 

technology testbed, 675–76; terrain system with moving map, 675–76; VoiceControlled Interactive Device (VCID)/

Voice Command System, 672–73, 674, 675

Advanced Research and Technology, Office 

of, 212, 473, 654

Advanced Subsonic Combustion Rig 

(ASCR), 799

Advanced Subsonic Technology (AST) program, 69, 786, 790, 796–97, 798

Advanced Supersonic Technology (AST) 

program. See Supersonic Cruise Research 

(SCR)/Supersonic Cruise Aircraft 

Research (SCAR)/Advanced Supersonic 

Technology (AST) program

Advanced Tactical Fighter (ATF) program, 

126–27

Advanced Tiltrotor Transport Technology 

(ATTT) program, 167–68

Advanced Turboprop (ATP) project, 744, 

747–50, 786, 788, 789, 790–91, 799–803

Advancing Blade Concept (ABC) rotor, 152–

53, 157–58

Adverse yaw, 19–20, 409

Advisory Group for Aerospace Research & 

Development (AGARD), 213

Aerion Corporation, 249

Aero Digest, 106, 106n48

Aerodynamic damping, 363–67

Aerodynamic heating, 414–15

Aerodynamics Committee, NACA, 59

Aerodynamic stability frequencies, 378

Aerodynamic structures, 805–6

Aeroelastic effects: aeroservoelastic effects, 

496, 525, 528–29; aircraft structure, 387, 

390–92, 397–98, 482–83, 521–22; B-47 aircraft, 20, 390, 409–10; B-52 Stratofortress, 

390–91, 392; B-58 Hustler, 392; FB-111 

aircraft, 392; rotary wing aircraft, 149–

50; SR-71 Blackbird, 391–92; swept and 

delta wing aircraft, 390; X-29 aircraft, 496, 

528–29, 694; XB-70 Valkyrie, 392

Aeroelastic Rotor Experimental System 

(ARES), 164

Aeroelastic tailoring, 521, 524–25, 694, 695

Aeroflightdynamics Directorate, 170

Aeronautical Research Associates of Princeton (ARAP), 213, 223

Aeronautical Research Institute of Sweden, 

212, 222

Aeronautical Research Laboratory, 186

Aeronautical Research Radar Complex, 151

Aeronautics and Space Technology, Office 

of, 219–20

Aeronautics Research Mission Directorate, 770

Aeronautics System Division, 220

Aeroservoelastic effects, 496, 525, 528–29

Aerospace America, 340

Aerospace Industries Association, 248

Aerospace Maintenance and Regeneration 

Group (AMARC), 672
881

Index 

Aerospaceplane (ASP) program, 334

Aerospace Vehicles Panel, Air Force Scientific Advisory Board, 585

Aerospatiale Dauphine helicopter, 165

Aerothermodynamic Elastic Vehicles (AEV), 

308–10, 594, 595

Aerothermodynamics. See also hot structures: ablation cooling, 286–87, 417–18, 

581–82, 591, 606–8; active cooling concepts, 420, 589–90, 605; aerodynamic heating, 414–15; blunt-body design, 291, 311, 

415; ceramic tiles, 419; Dryden research 

activities, 495–96; heat-sink structures, 

415–17, 579, 581, 607; LI-900 thermal protection, 311, 313–17, 600–601, 604, 605; 

LI-1500 thermal protection, 310–11, 313, 

599–600, 601, 604; LI-2200 thermal protection, 317; localized heating, 416; materials selection and speed, 569–73; passive 

cooling, 606; reusable surface installation 

(RSI), 312–23, 596–605; satellites, conditions for, 305, 579, 581–82; speed limits, 572–73; Structural-Thermal-Optical 

(STOP) program, 501–2; thermal barrier 

coatings (TBC), 610–13; tiles, thermalprotection, 278, 288, 310–23, 419, 582, 

598; tiles, thermal-protection, bonding 

of, 319, 320–23, 609; tiles, thermal-protection, coatings for, 315–16; tiles, thermal-protection, densification of, 320–21; 

tiles, thermal-protection, gaps between, 

318, 319–20; tiles, thermal-protection, 

installation of, 319–23, 419; tiles, thermalprotection, shaping of, 314–15, 321; tiles, 

thermal-protection, testing of, 316–19; 

tiles, thermal-protection, thermal expansion and, 315; transpiration cooling, 579–

80; water-wall cooling, 307, 589–90

Aerothermodynamic Structural Vehicle 

(ASV), 308–10, 594–95

AeroVironment, Inc., 853, 855, 856, 859, 

863, 864, 868, 871

AFT-100 aircraft, 221

Agility, 651

AGM-129 FSW stealthy cruise missile, 22

Agusta Helicopter Company, 168

Aileron reversal, 409–10

Ailerons: aeroelastic effects, 390; flight control coupling, 409; flight control with, 633; 

swept wing aircraft, 19–20, 21

AIM-9 Sidewinder air-to-air missile, 378

Air, compressed, 324, 573

Airborne Real-Time Imaging System 

(ARTIS), 861

Air-breathing vehicles, 279

Airbus A310 aircraft, 615

Airbus A319 aircraft, 124

Airbus A320 aircraft, 124, 616, 722, 753

Airbus A330 aircraft, 124, 616

Airbus A340 aircraft, 124, 616, 752

Airbus A350 aircraft, 616

Airbus A380 aircraft, 571, 616, 643, 722, 795

Airbus Industrie, 692, 722

Air combat simulation, 703

Aircraft: challenges of developing, 361–62; 

evolution of, 3, 361–62; flight envelope 

limitations, 368, 632; handling qualities 

criteria, 145–46; nicknames for, 3–4; safety 

of, 506, 632; speed limits, 572–73; turbojet 

engine development and, 3, 90; unstable 

aircraft, 379–80, 383, 526, 631, 632, 633, 

753; weight of, 461–62, 463

Aircraft and Armament Experimental 

Establishment (A&AEE), 638n16

Aircraft carrier landings, 33, 33n60

Aircraft design. See also area rule; computational fluid dynamics (CFD): aerodynamic predictions, 397–98; bullet 
NASA’s Contributions to Aeronautics

882

shapes, 97; commercial aircraft, 208; 

DC-4 Generation, 4, 4n1; documentation 

and communication of design and system requirements, 683–84; evolution of, 

3–4, 64; induced drag, 117, 118; NASA’s 

role in, 65; noise reduction and, 770–72; 

optimization of, 533; profile drag, 117; 

refinement and efficiency improvements, 

64–65, 89–90; safety factor, 387; stability 

and control and, 633; statically unstable 

aircraft, 633; supersonic aircraft design, 

93, 96; transonic aircraft, 93–94, 362–63, 

444–46; transonic drag, 94–100

Aircraft Energy Efficiency (ACEE) program: 

Advanced Turboprop (ATP) project, 744, 

747–50, 786, 788, 789, 790–91, 799–803; 

advancements through, 758; Composite 

Primary Aircraft Structures project, 744, 

804–5; Energy Efficient Engine (E Cubed) 

project, 742, 744, 745–47, 786, 788, 790–

91, 792–95, 817; Energy Efficient Transport (EET) project, 744, 805–6; Engine 

Component Improvement project, 744–45, 

746; Laminar Flow Control (LFC) project, 

744, 784, 786, 806–8; purpose of, 743–44, 

788; winglets development, 116–17, 120, 

121, 123

Aircraft Engine Research Laboratory, 137, 

786–87. See also Glenn Research Center/Lewis Research Center/Lewis Flight 

Propulsion Laboratory/Aircraft Engine 

Research Laboratory

Aircraft handling qualities criteria, 145–46

Aircraft noise. See noise

Aircraft structure: aerodynamic structures, 

805–6; aeroelastic effects, 387, 390–92, 

397–98, 409–10, 482–83, 521–22; aeroelastic tailoring, 521, 524–25; aeroservoelastic 

effects, 496, 525, 528–29; Computational 

Structural Mechanics (CSM) workshop, 

497; electro-optical flight deflection measurement system, 521–22, 527–28; flutter, 306, 309, 387–89, 482–83, 525–26, 

585–86, 595; redundant structures, 463–

64; research activities, meeting about, 

473; structural loads, 385–86, 390; structural resonances, 374–79; testing facilities, 553–55

Aircraft Synthesis (ACSYNT), 493

Air defense missions, 187–88

AiResearch, 335, 503

Airfoil design and theory. See also supercritical wing: evolution of, 109; NACA airfoil 

sections, 29n56, 109, 115

Air Force, U.S. (USAF): Aeronautical 

Systems Division, 337, 526; Aerospace 

Maintenance and Regeneration Group 

(AMARC), 672; Aerospaceplane (ASP) 

program, 334; AFTI/F-16 aircraft, 679; air 

defense missions, 187–88; Arnold Engineering Development Center (AEDC), 

17n30, 158, 280, 291, 760, 787; ASSET 

vehicles, 308–10, 418–19; Atlas ICBM, 

581, 582; ballistic missile accuracy, 284–

85; biofuel research, 817; Bomarc missile, 

289; BWB research, 71; C-17 Globemaster 

III, 127, 722; Century series fighters. See

Century series fightersconcurrent development, 322; Digital Electronic Engine 

Control (DEEC) program, 709–10, 754–

56; energy maneuverability concept, 651; 

Exceptional Service Medal, 107; F-22 Raptor, 126–27, 712; F-102/F102A Delta Dagger, 102–3; F-104 Starfighter, 104; F-111 

program, 61–62; fighter aircraft program, 61; fly-by-wire systems, development of, 379–80; fly-by-wire systems, 

research activities, 645–48; fly-by-wire 
883

Index 

systems, transition to, 632; Goodyear 

Electronic Digital Analyzer (GEDA), 

393, 394; hydrogen fuel research, 811–

13; HyTech research program, 345; Integrated Propulsion Control System (IPCS) 

program, 754; International Test Organization (ITO), 700–701; KC-135 aerial 

tanker, 20, 120, 122–24, 204, 806; Lightweight Fighter (LWF) program, 370, 647, 

664; missile tests, 288; NASTRAN, development of, 474; NASTRAN, Users’ Colloquia, 480; National Aero-Space Plane Joint 

Program Office, 338–39; National FullScale Aerodynamics Complex, 158, 166; 

Navaho cruise missile, 289; 1954 Interceptor initiative, 50; Noise and Sonic Boom 

Impact Technology (NSBIT) program, 

229–30, 231; Precision Aircraft Control 

Technology (PACT), 647–48; Project 680J 

Survivable Flight Control System (SFCS), 

379–80, 384, 646–48, 658n59, 660; scramjet research, 334–35; Self-Repairing Flight 

Control System (SRFCS), 713–14; Shuttle study, 450; side stick flight controller, 

658n59; sonic boom research, 194–203, 

207–8; sonic booms, concerns and complaints about, 185–88; Space Division 

(SD), 337; stability augmentation system research, 365; STOL transport competition, 114; strain gages, 386; Strategic 

Defense Initiative (SDI), 279, 336–40; 

structural resonance research, 378; Suntan project, 811–13; supersonic aircraft 

research, 182; synthetic fuel research, 

816–17, 816n114; Test Pilot School. See

Test Pilot School, USAF Transonic Aircraft 

Technology (TACT) program, 113, 386; 

V-22 Osprey, 166; Vietnam war aircraft 

losses, 644n32; Wright Air Development 

Center, Flight Control Laboratory, 639; 

Wright Laboratory, 541; X-15 hypersonic 

flight, 290–91; X-20 Dyna-Soar program, 

303–4, 400; X-29 aircraft, 526; X-31 aircraft, 707; X-51A vehicle, 345; YF-12 

Blackbird project, 517; YF-102/YF102A 

Delta Dagger, 100–102

Air Force Academy, 216

Air Force Aero-Propulsion Laboratory, 335

Air Force Air Material Command, 59

Air Force Flight Dynamics Laboratory 

(AFFDL): Advanced Fighter Technology Integration (AFTI) F-16 project, 670; 

Control Configured Vehicle (CCV) F-16 

project, 669–70; Digital Flight Control 

for Tactical Aircraft (DIGITAC), 382–83, 

684–86; electronic flight control system, 

372; FLEXSTAB program, 544–45; flutter suppression research, 388; fly-by-wire 

system research, 647; Highly Maneuverable Aircraft Technology (HiMAT), 686–89; 

JB-47E program, 645–46; reentry vehicle 

research, 404, 406, 415; supercritical wing 

refinement, 113; Transonic Aircraft Technology (TACT) program, 113, 386; X-24C 

computed surface streamlines, 299, 448–

49; X-29 aircraft, 692

Air Force Flight Test Center (AFFTC): 

F-100A research, 44; inertia table, 396; 

limit-cycle oscillation research, 368; simulators and flight simulation, 403–4, 406–7; 

sonic boom research, 195–96, 201–3; sonic 

booms and Century series fighters, 185

Air Force Flight Test Center Museum, 686

Air Force Institute of Technology, 542

Air Force Museum, 648, 679, 709

Air Force Plant 42, 196

Air Force Research Laboratory (AFRL): 

Advanced Control Technology for 
NASA’s Contributions to Aeronautics

884

Integrated Vehicles Experiment (ACTIVE) 

project, 716–17; Air Vehicle Directorate, 

388; Electrically Powered Actuator Design 

(EPAD) program, 384

Air Force Scientific Advisory Board, 585

Air Force Systems Command (AFSC), 29n56

Air France Concorde. See Concorde Supersonic Transport (SST)

Airlines, profitability of, 743–44, 788

AIRLOADS program, 556

Air Mobile Airborne Division, 139

Air pollution, 736. See also emissions

Airport noise. See noise

Air Proving Ground, 44

Air resonance, 149

AJ-37 Viggen, 391, 651n49

Alford, William J., Jr., 36, 60–61, 61n115, 

638n16

Allee, Cody, 706, 707

Allen, Harry Julian “Harvey”: blunt-body 

theory, 54, 285–86, 291, 311, 415, 440–41, 

579, 599; flattop delta aircraft, 56; nickname for, 441

AlliedSignal Engines, 796

Allison Engines, 796, 800–801

Allison/Rolls-Royce, 762, 797

Allison turbo shaft engine, 800–801

All-moving tail research, 34–35, 35n65, 362–

63, 363n4

Alloy 454, 614–15

Almroth, B.O. “Bo,” 551

Alonso, Juan, 494

Alper, Marshall E., 474

Alumina, 590

Aluminum: ablation cooling and, 417; aircraft built with, 310, 571, 602, 606–9, 

620; Alloy 454, 614–15; bonding to, 609; 

chemical milling, 569; composite materials compared to, 804; high-speed aircraft 

and, 569; sonic booms and, 210; Space 

Shuttle orbiter built with, 310, 602, 606–

9; structural analysis for building with, 

464; superalloy addition, 575, 583; temperature and strength of, 569

American Helicopter Society (AHS): Annual 

Forum Proceedings, 138; charter of, 135; 

Grover E. Bell Award, 169–70; Journal, 

138; rotary wing noise research, 163, 

164–65

American Institute of Aeronautics and 

Astronautics (AIAA): Astronautics & Aeronautics, 599; research results, dissemination of through, 138, 207; Sylvanus Albert 

Reed Award, 116

American Society of Mechanical Engineers, 

759

American Technology Alliances (AmTech), 

852n33, 853, 855

America’s Cup yacht race, 124–25

Ames-Dryden AD-1 oblique wing testbed, 

63–64, 515, 522–26, 530

Ames Research Center/Ames Aeronautical 

Laboratory: Advanced Concepts Simulator, 715; Advanced Tiltrotor Transport 

Technology (ATTT) program, 167–68; 

Advanced Turboprop (ATP) project, 800; 

Aeronautics Research Mission Center, 

transformation from, 166; aircraft handling qualities criteria, 145–46; Air Force 

F-111 program, 61–62; Army research 

presence at, 139, 147; blunt-body problem solution, 442; computational fluid 

dynamics (CFD), 234–36, 437–38, 451; 

computers, analog, 468; computers, ERA 

1103 digital, 437; computers, IBM digital, 

437–38; computers, Illiac IV, 438; CONMIN program, 492–93, 498, 512; delta 

wing development, 53; ENSAERO code, 
Index 

885

493–94; finite element analysis activities, 

491–94; flattop delta research, 55–56; 

FLEXSTAB program, 518, 544–45; Flight 

Simulator for Advanced Aircraft (FSAA), 

151; Full-Scale Tunnel, 147, 153, 156–58; 

high-altitude aircraft workshop, 850, 851; 

Highly Maneuverable Aircraft Technology 

(HiMAT), 687, 689; Hypersonic Weapons 

Research and Development Supporting 

System (HYWARDS), 300–301; hypersonic wind tunnel, 283; Illiac IV computer, 325; Intelligent Flight Controller 

(IFC), 720; Intelligent Flight Control Systems (IFCS) project, 719; JF-100C Super 

Sabre, 402n62; Joint Sponsored Research 

Agreement (JSRA), 852n33; M-2 lifting 

body, 56; NASTRAN, development of, 474; 

NASTRAN, Users’ Colloquia, 536; National 

Full-Scale Aerodynamics Complex, 158, 

165–66; National Rotorcraft Technology 

Center, 168–69; OVERFLOW code, 452, 

453; pitch-up research, 25–26; Propulsion Controlled Aircraft (PCA) program, 

715; rotary wing aircraft research, 137, 

147, 148, 149–50, 152–53, 155–61, 163–

64, 492; rotary wing noise research, 163; 

Rotor Systems Research Aircraft (RSRA) 

program, 155, 163–64, 165; Rotor Test 

Apparatus (RTA), 153; RSI coatings, 315; 

Science Mission Center, transformation 

to, 166; sensor technology development, 

861; sonic boom research, 214, 218, 223–

24, 233, 235; Space Shuttle program, 

56–57, 327; SST program, 66, 191; stability 

research activities, 637; Steep Approach 

program, 737; Structural Dynamics Laboratory, 553; tail configuration research, 

36–37; thermal-protection tiles, 316–17, 

319, 419; Tilt Rotor Research Aircraft 

(TRRA) program, 155, 156, 160–62, 165, 

166; Transonic Aircraft Technology (TACT) 

program, 386; Ultra Efficient Engine Technology (UEET) program, 762, 797; Unitary Wind Tunnel, 214, 235, 586; variable 

stability aircraft, 401; Vertical Motion 

Simulator (VMS), 151, 167; wind energy 

program, 829; wind tunnel studies, 25, 51, 

327, 524; X-29 aircraft, 526; YF-12 Blackbird project, 517

AmTech, 852n33, 853, 855

Analex Corporation, 539

Analog computers: active-element computers, 468; direct analog computers, 466–

68; flight simulation and, 392–94, 405–7, 

468; fly-by-wire systems, 648–49, 656, 

656n57, 660, 664; hybrid direct-analog/

digital computers, 468; indirect analog 

computers, 468

Anderson, Griffin, 336

Anderson, John D. Jr., 341–42

Angle of attack (AoA): aerodynamic damping and, 364, 366; arrow wing and, 69; 

deep stall and, 67n127; dihedral effects, 

410; fly-by-wire systems and angle-ofattack limiters, 665–66; oblique wing concept, 524–25; pitch-up and swept and delta 

wing aircraft, 21, 33, 34, 34n61, 36, 50, 

56; Shuttle reentry, 57; X-29 aircraft, 694, 

695, 696–98; X-31 aircraft, 701–2, 706–7

Anglo-French Concorde SST. See Concorde 

Supersonic Transport (SST)

Anhedral effects, 29, 34, 34n62

ANSYMP program, 545

Antelope Valley, 247

Antenna Design and Analysis, 549–50

Antipodal aircraft. See Silver Bird 

(Silbervögel)

Antispin parachute, 666n70, 695, 696
NASA’s Contributions to Aeronautics

886

Apollo 8 mission, 513

Apollo 9 mission, 513

Apollo 11 mission, 654

Apollo 13 mission, 513

Apollo program: ablation cooling, 417; centrifuge program, 400; Command and Service Module, 654; Display and Keyboard 

(DSKY), 655, 655n55; flight control systems, 644, 753; fly-by-wire systems, 653–

54, 663; fuel cells, 814; hydrogen fuel, 814; 

Inertial Management System (IMS), 655; 

Lunar Landing Training Vehicles (LLTVs), 

653–54; Lunar Module, vii, 654; moon 

landing mission, vii; pogo, 328; pogo 

oscillations, 513; reentry vehicles, 54; side 

stick flight controller, 658n59; silica fiber 

radome, 598; sonic boom research, 206, 

226; success of, 64

Applications Engineering process, 488

Applied Physics Laboratory, Johns Hopkins 

University, 332, 333

Apt, Milburn “Mel,” 44n87, 394–95, 407

Aramid fibers, 767

Arc-Heated Scramjet Test Facility (AHSTF), 

343

Arc-jet, 316–17, 415

Area rule: aircraft built with, 104; announcement of, 106–7, 106n48; application of 

to aircraft design, steps in, 100; benefits 

of, 89–90, 126; development of, 95n14, 

97–100; models, 88, 97, 99; naming of, 

99; recognition of, 107–8; research memorandum on, 99–100; secret status of, 106; 

transonic transport fuselage, 127

Area rule validation in flight, 100–106; F8U 

Crusader series, 103–4, 105; F9F-9/F11F-1 

Tiger, 104, 105, 106; F-102 Delta Dagger, 

51, 52; F-104 Starfighter, 104; YF-102 

Delta Dagger, 51, 89, 100–102, 103

Argonne National Laboratory, 540

Argyris, John, 471

Arlington, Virginia, 537

Arms race, 442

Armstrong, Neil, vii, 64, 276, 307, 591, 654

Armstrong-Whitworth flying wing, 70

Army, U.S. (USA): Aeroflightdynamics 

Directorate, 170; Air Mobile Airborne Division, 139; Armament Research & Development Command, 481–82, 540, 541; Atlas 

ICBM, 581, 582; Aviation Applied Technology Directorate, 146; aviation laboratory 

(AVLABS), 146, 150; Benet Laboratory, 

541; Frankford Arsenal, 542; Harry Diamond Laboratory, 541; missile tests, 288; 

NASA, relationship with, 139, 146, 147, 

152; NASTRAN Users’ Colloquia, 480, 

481–82, 540; National Rotorcraft Technology Center, 168–69; rotary wing aircraft 

research, 147, 150, 154–55, 166, 168–69, 

170–71; Rotorcraft Division, Army-NASA, 

166, 170; Rotor Systems Research Aircraft 

(RSRA) program, 154–55; Tactical Mobility Requirements Board (Howze Board), 

139; Tank Automotive Command, 549; 

Tilt Rotor Research Aircraft (TRRA) program, 155, 165, 166; Vietnam war aircraft 

losses, 644n32

Army Aeronautical Activity at Ames

 (AAA–A), 147

Army Air Corps (AAC), 141–42

Army Air Forces (AAF): Dart, sponsorship 

of, 9–10; glomb (guided glide bomb), 11; 

swept wing research, 13–14; technical 

intelligence interrogations, 14–15, 17

Army Air Mobility Research & Development 

Laboratory, U.S., 492–93, 540

Army Air Service Engineering Division, 118

Army Research Laboratory, 612
Index 

887

Army Research Office, 169

Army Structures Laboratory, 162–63

Army Tactical Missile System (ATACMS) 

rocket booster, 345

Army Transportation School, Fort Eustis, 

Virginia, 146

Arnold, Henry H. “Hap,” 14, 291

Arnold, Steven, 500

Arnold Engineering Development Center 

(AEDC), 17n30, 158, 280, 291, 760, 787

Arrowplane, 5

Arrow-Shaped Lifting Surface (Pfielförmiges Tragwerk), 6

Arrow wing, 65, 67–68, 69, 209, 225

Artificial feel devices, 635, 643

Aspect ratio of a wing, 119–20, 407

ASSIE-I program, 552

Astronautics & Aeronautics, 599

Asynchronous digital flight control system, 

672, 680–82, 683

Atar 9K-50 engine, 691

Atlanta, 206

Atlas intercontinental ballistic missile 

(ICBM), 284–85, 286, 288, 293–94, 579–82

Atmospheric Effects of Aviation project, 797

Atmospheric Effects of Stratospheric Aircraft project, 795–96

Atomic bombs, 284

Atomic Energy Commission (AEC), 206–

7, 213, 481

Augenstein, Bruno, 284

Augmented Longitudinal Control System 

(ALCS), 372

Ault, Stan, 861

Aurora Flight Sciences, 851, 853

Autogiros, 135, 137, 140, 141–42, 145

Automated Maneuvering Attack System 

(AMAS), 672, 673, 674–75, 676

A u t o n o m o u s F r e q u e n c y D o m a i n 

Ident ification (AU-FREDI), 504, 505

Autopilots, 631, 687–88, 706

Autothrottle system, 706

Auxiliary power units (APUs), 295, 503, 815

Avery, William, 332

Aviation Applied Technology Directorate, 

Army Transportation School, Fort Eustis, Virginia, 146

Aviation industry: airlines, profitability of, 

743–44, 788; certification of new aircraft, 

683–84; design of commercial aircraft, 

208; finite element structural analyses 

use by, 533–34; fly-by-wire systems, 632, 

722–23; NASTRAN use by, 478, 480–81; 

NASTRAN Users’ Colloquia, 480; supercritical wing use by, 114–15; winglet use 

by, 120–22, 806

Aviation laboratory (AVLABS), 146, 150

Aviation Partners, Inc., 124

Aviation Partners Boeing, 124

Aviation Week, 106, 106n48, 319, 639, 856

Aviation Week and Space Technology, 253, 

260–61

Avro 707C aircraft, 643

Avro Canada, 641–42, 642n25

Avro CF-105 Arrow, 641–42, 642n25, 643

Avro Vulcan bomber, 45, 639

B

B-1 bomber, 62, 62n118, 378; B-1A bomber, 

392; B-1B bomber, 392

B-2 Spirit, 722, 753

B-17 Flying Fortress, 736

B-26 Invader, 402n62

B-29 Superfortress, 29, 38, 92, 290, 736

B-36 bomber, 55

B-45 aircraft, 634–35

B-47 aircraft: aeroelastic effects, 20, 390, 

409–10; B-47B aircraft, 645, 736; B-47E 
NASA’s Contributions to Aeronautics

888

aircraft, 658n59; JB-47E aircraft, 645–46; 

side stick flight controller, 658n59

B-50 aircraft, 290

B-52 Stratofortress: aeroelastic effects, 390–

91, 392; DAST launch from, 389; flight 

control systems, 752; NB-52/NB-52B Stratofortress, 343, 369, 378–79, 568; vortex 

generators, 65; X-15 launch from, 197

B-57B bomber, 810, 811

B-58 Hustler: aeroelastic effects, 392; B-70 

as replacement, 191; cross-country flight 

by, 188; design of, 45, 53, 103; sonic boom 

research, 194, 195–98, 199, 200, 202, 203, 

204, 205, 209; sonic booms from, 188–89; 

supersonic flight by, 188, 189

B-70 aircraft, 191

Baals, Donald, 99

Babcock and Wilcox, 601

Backup Control System (BCS), 656, 657, 

658, 660

Baer-Riedhart, Jenny, 854–55

Balaclava, 828

Baldwin, Darrell H., 830

Baldwin-Lomax algebraic model, 341

Ball bearings, 307, 590

Ballistic Missile Defense Organization, 850

Ballistic missiles. See also intercontinental 

ballistic missiles (ICBMs): development 

of, 284; intermediate-range ballistic missile (IRBM), 286, 442, 571–72; V-2 ballistic missiles, 183–84, 280, 282, 571, 635–36

Ballistic shock waves, 183–85

Bamford, Robert M., 474

Banas, Ronald, 310, 598

Bare Reactor Experiment Nevada (BREN) 

Tower, 206–7

Barger, Raymond L., 223–24, 227

Barlow, Robert, 647

Barrowsville, Virginia, 634

Barter, Steve, 678

Base Amplitude and Direction Sensor 

(BADS), 247

Batch operation, 437–38

Battelle Memorial Institute, 230

Bauer, Jeffrey, 855

Baughman, John, 791, 793

Beam, Richard M., 474

Beamont, Roland, 23–24, 24n40

Bearingless Main Rotor (BMR), 157

Bearings, 307, 590

Beasley, R.M., 310, 598

Becker, John V.: aerothermodynamic 

research, 93, 418, 573; flat-bottom delta 

aircraft, 55; hypersonic aircraft and flight 

research, 282–83, 291; HYWARDS, 300, 

301; scramjet engines, 333; on Whitcomb’s 

research techniques, 113; X-20 Dyna-Soar 

program, 307

Beech C-45 aircraft, 199

Beechcraft, 481

Beechcraft Starship, 121

Beechcraft Super King Air 200 twin turboprop aircraft, 481

Beechcraft T-34C Mentor, 481

Beech L-23 sailplane, 248, 257

Beggs, James, 192

Belgium, 184

Bell, Lawrence D., 93, 94

Bell Aerospace Company, 479

Bell AH-1 Cobra FLITE configuration, 170

Bell Aircraft Corporation, 17n30, 290

Bell Award (AHS), 169–70

Bell BA 609 aircraft, 168

Bell BoMi bomber-missile, 290, 300

Bell Brass Bell boost-glider, 300, 301, 302–3

Bell H-13 helicopter, 150

Bell Helicopter Company/Bell Helicopter 

Textron: Design Analysis Methods for 
Index 

889

Vibration (DAMVIBS), 508, 509, 509n128; 

helicopter design and production, 136; 

RITA membership, 169; tilt rotor design, 

153

Bell L-39 aircraft, 22–23, 23n37, 23n39

Bell Lunar Landing Research Vehicles 

(LLRVs), 653–54

Bell Model 222 rotary wing aircraft, 156

Bell P-39 Aircobra, 23n37

Bell P-63 Kingcobra fighters (L-39 aircraft), 

22–23, 23n37, 23n39

Bell RoBo rocket bomber, 300, 301, 303

Bell UH-1 Huey, 147

Bell V-22 Osprey Tilt Rotor, 136, 155, 161, 

166

Bell X-1 aircraft: Collier Trophy, 93; development and manufacture of, 108, 290, 

494; engines for, 294; flight control systems, 362–63, 752; gliding landings, 402–

3; lift-to-drag ratio (L/D), 403; Round One 

research aircraft, 55n106; supersonic 

capabilities, 94; tail configuration, 411; 

X-1A aircraft, 40–41, 40n74, 43, 292, 363, 

407; X-1B aircraft, 40n74, 298; X-1D aircraft, 40n74; Yeager’s flight, 40–41, 292, 

363, 407

Bell X-2 aircraft: accident with, 44n87, 394–

95, 407, 410; control coupling, 410; design 

of, 34; development and manufacture of, 

17n30, 290; dihedral effects, 410; flight 

control systems, 298; inertial coupling, 

44n87; pressure suit for, 297; simulator, 

394–95; tail configuration, 411

Bell X-5 aircraft, 17n30, 58, 59–60, 413

Bell XS-1 aircraft: design of, 3, 28, 34, 

34n62, 36, 184; falling body test, 38, 43; 

flight control systems, 635; inspiration 

for, 14n22; sonic boom from, 182, 182n3; 

supersonic flight by, 182, 183; testing of, 

35; variable-geometry wing tests, 59; wingspan-to-fuselage length ratio, 37; XS-1 

FSW planform, 21; Yeager’s flight, 182, 

182n3, 283, 635

Bell XS-2 aircraft, 34, 34n62

Bell XV-1 Convertiplane, 134, 147

Bell XV-15 Tilt-Rotor Research Aircraft 

(TRRA), 155, 156, 160–62, 165, 166

Bendix nose skid, 592

Bendix ramjets, 289

Benet Laboratory, 541

Benscoter, Stanley, 467

Berg, Moe, 332–33

Beryllia, 587

Beryllium, 607, 620–21

Beta-21S, 617–18, 621

Betz, Albert, 15

Beyond the Limits gallery, National Air and 

Space Museum, 427, 432

B.F. Goodrich, 481, 538

Big Boom, Project, 195

Bikle, Paul, 296

Bi-Level Integrated System Synthesis 

(BLISS), 511

Billig, Frederick, 332

Biofuel, 817

Biplanes, 4–5

Birchenough, Arthur, 830

Birds, 385, 386

Bisplinghoff, Raymond, 473

Blackburn Buccaneer, 638n16

BLAde SIMulation (BLASIM), 547

Blade-vortex interaction (BVI) noise, 163, 

164–65

Blended wing-body (BWB): Boeing X-48B 

Blended Wing Body UAV, 2; BWB-450, 

71; development of, 70–72; heavy-lift military missions, 71, 71n131; hybrid laminar flow control, 808; performance of, 
NASA’s Contributions to Aeronautics

890

415; thermodynamic characteristics, 415

Blériot, Louis, 385

Block Island, Rhode Island, 827

Blow-down design, 280, 283

Blue Angels, 104

Blue Ridge Electrical Membership Corporation, 835

Blunt-body theory: computational fluid 

dynamics and solution to, 439–44; development of, 54, 415, 441; limitations of, 

415; numerical solution for, 442–44; reentry vehicles, 54, 285, 291, 441, 599; shock 

waves, 286, 415, 440–42; Star Clipper 

spacecraft, 311; thermodynamic characteristics, 291, 311, 415, 579

BMW 801 command equipment, 633n1

Bobbitt, Percy “Bud,” 235, 246

Boccadoro, Charles H., 246

Bodies of revolution, 184

Boeing 367-80 aircraft, 20

Boeing 707 aircraft, 4, 20, 65, 123, 740, 804

Boeing 727 aircraft, 737, 745, 802, 804

Boeing 737 aircraft, 124, 616, 737, 745, 771

Boeing 747 aircraft: 747-400 winglet use, 

89, 123–24; composite materials, 804; 

engine component improvements, 745; 

NASA’s aeronautics role and, 787; Shuttle 

carrier aircraft, 661; simulators testing of 

PCA, 715; TAVs launched from, 337; titanium construction, 616; winglet use on, 

89, 122, 123–24

Boeing 757 aircraft, 124, 715, 756, 795, 

807–8

Boeing 767 aircraft, 124, 615

Boeing 777 aircraft: 777-300ER aircraft, 

795; fly-by-wire systems, 722, 753; GE90 

jet engine, 742, 743, 747; materials for, 

620; noise from, 771; supercritical wing, 

115; titanium construction, 616

Boeing 787 aircraft: electrical flight control surfaces and actuators, 722; engine 

for, 769, 795; materials for, 571, 620; noise 

from, 771; range of, 620; titanium construction, 616

Boeing B-17 Flying Fortress, 736

Boeing B-29 Superfortress, 29, 38, 92, 290, 

736

Boeing B-47 aircraft: aeroelastic effects, 

20, 390, 409–10; B-47B aircraft, 645, 736; 

B-47E aircraft, 658n59; JB-47E aircraft, 

645–46; side stick flight controller, 658n59

Boeing B-50 aircraft, 290

Boeing B-52 Stratofortress: aeroelastic 

effects, 390–91, 392; DAST launch from, 

389; flight control systems, 752; NB-52/

NB-52B Stratofortress, 343, 369, 378–79, 

568; vortex generators, 65; X-15 launch 

from, 197

Boeing Business Jet, 124

Boeing Company: Advanced Control Technology for Integrated Vehicles Experiment 

(ACTIVE) project, 717; biofuel research, 

817; BWB research, 71–72; columbium 

coating, 585; composite materials for 

aircraft parts, 804; Cooperative Test 

Organization/VECTOR, 705–6; F-16XL 

modifications, 261; financial problems, 

217; FLEXSTAB program, 518; high-altitude aircraft workshop, 851; High-Speed 

Civil Transport (HSCT), 231, 232, 243; 

hydrogen fuel cell aircraft, 815; Intelligent Flight Control Systems (IFCS) project, 719; McDonnell-Douglas merger, 127; 

NASA contract work, 120; NASTRAN 

Users’ Colloquia, 480; nose cap design, 

587, 588–89; Numerical Propulsion System Simulation (NPSS) program, 760; 

P&W turbo fan engines on jetliners, 710; 
Index 

891

Phantom Works, 71; QSP proposal, 245; 

RITA membership, 169; Shaped Sonic 

Boom Demonstration (SSBD), 246; sonic 

boom research, 213, 224; SST program, 

contract for, 211, 217–18; SST program, 

data to support, 519; SST program, proposal for, 210; STOL transport competition, 114; supersonic business jet (SSBJ) 

concept, 249; Supersonics Project, 261; 

Trans Atmospheric Vehicles (TAV), 337; 

Ultra Efficient Engine Technology (UEET) 

program, 762, 797; Vertol Division, 136; 

wind energy program, 829, 835, 838–39, 

840, 841; winglet design, 124; winglet use 

by, 121–22, 124

Boeing F-15 aircraft, 710. See also McDonnell-Douglas F-15 Eagle

Boeing FLEXSTAB, 392

Boeing Helicopters, 508, 509–10. See also

Boeing-Vertol entries

Boeing JB-47E aircraft, 645–46

Boeing KC-135 aerial tanker: aerial refueling, 253; design of, 20; sonic boom 

research, 204; winglet use on, 120, 122–

23, 806

Boeing NB-52/NB-52B Stratofortress, 343, 

369, 378–79, 568

Boeing PB2-1S Superfortress (B-29), 29

Boeing Scientific Research Laboratories, 

445

Boeing Spanloader, 70

Boeing SST concept, 63, 68

Boeing-Vertol, 540

Boeing-Vertol Bearingless Main Rotor 

(BMR), 157

Boeing-Vertol CH-47 Chinook, 151, 162, 

163; CH-47C Chinook, 152; CH-47D Chinook, 509–10

Boeing-Vertol tilt rotor concept, 153

Boeing-Vertol XCH-62 Heavy Lift Helicopter (HLH) program, 158

Boeing WC-135B Stratotanker, 204

Boeing X-20 Dyna-Soar: cancellation of, 

304, 308, 400, 592; centrifuge program, 

400; contributions from program, 307–8; 

design confidence, 307–8; development of, 

300–305; flight control system, 372; flutter-testing panels, 585–86; hot structures, 

305–8, 418, 582–92; justification of, 302–

5; landing skids, 592; materials for, 305–8; 

missiles to launch, 301–2; nose cap, 306, 

308, 587–89, 598; Round Three research 

aircraft, 55–56; Round Two research aircraft, 55n106; schematic drawing, 583; 

simulator, 405–6; windows, 307

Boeing X-48B Blended Wing-Body UAV, 2, 

71–72

Boeing X-51A vehicle, 279, 345–46

Boeing XB-52 aircraft, 20

Boeing YC-14 transport, 114–15, 670–72

Bomarc missile, 289

Bombardier CSeries aircraft, 752

BoMi bomber-missile, 290, 300

Bond, Alan, 124–25

Bongo, Operation, 197–98

Bongo II, Operation, 199–200

Bonneville Power Administration, 837–38

Boom Event Analyzer Recorder (BEAR), 

229–30, 238

Boomfile, 229, 238

Boone, North Carolina, 835

Boosted (hydraulic) control-surface actuators, 384, 590, 631, 634–35, 643, 644

Boost-glider, 55

Boria, 315

Boron, 574, 583, 614

Boron, California, 182, 202

Borosilicate glass coatings, 315–16, 599–600
NASA’s Contributions to Aeronautics

892

Boscombe Down, 638n16

Bossart, Karel J. “Charlie,” 284

Boston, 536

Bosworth, John, 719

Boundary condition, 410, 431

Boundary layer: boundary layer control 

(BLC) experiments, 25, 147; laminar flow 

control research, 235, 495, 744, 784, 786, 

806–8; sonic boom research and, 196

Bowles, Mark, 802

Bradshaw, Peter, 341

Brass Bell boost-glider, 300, 301, 302–3

BREN Tower, 206–7

Breuhaus, Waldemar “Walt,” 401

Brewer, Jack D., 36

Bright, George W., 648

British Aeronautical Research Committee, 

118

British Aerospace (BAe), 690

British Aircraft Corporation BAC 1-11, 

67n127

British Airways Concorde. See Concorde 

Supersonic Transport (SST)

Brooklyn Polytechnic Institute, 332–33

Brooks, Peter, 4

Brown, Clinton E., 67

Brugh, Robert L. “Bob,” 485, 536

Brush, Charles F., 825

BUCKY program, 507, 550

Bullet, Project, 103–4

Bulzan, Dan, 799, 816–17

Bureau of Reclamation, 839

Burgess Company, 5

Burgess-Dunne seaplanes, 5

Burke, Melvin E., 654, 655

Burning, Pieter, 452

Burroughs Corporation Illiac IV computer, 

325, 438

Burstein, Adolph, 46, 46n90

Busemann, Adolf, 5–6, 15, 15n27, 93, 98, 99

Butler, Thomas G., 473, 474, 476, 532

Butler Analyses, 539

BWB-450 aircraft, 71

C

C-5 aircraft, 337, 407, 867

C-17 Globemaster III transport, 127, 384, 

715, 722, 795

C-45 aircraft, 199

C-47 transport, 195, 411

C-131B aircraft, 204

C-141 aircraft, 324, 658n59

Cadmium, 570

Cadmium-plated tools, 570

California Institute of Technology (Caltech), 

12, 17n30, 502

CALSPAN (Cornell Aeronautical Laboratory), 401–2, 402n62

CALSPAN NT-33 aircraft, 642

CALSPAN VISTA Lockheed-Martin NF-16, 

402n62

Calverton, New York, 528

Cambridge, 736

Campbell, John, 63

Canada, 641

Canadair Challenger 600 aircraft, 115

Canadair Sabre, 24

Canards, 224, 391, 392, 519, 526, 647–48

Carbides, formation of, 574–75. See also

silicon carbide; titanium carbide; tungsten carbide

Carbon, 614

Carbon braid engine fan casing, 768, 769

Carbon-carbon composite, 318, 576–78, 

618–20

Carbon composites, 619–20

Carbon dioxide emissions, 790, 790n22, 

797–98, 799, 813

Carbon fibers, 576, 578, 620, 768,
Index 

893

 769, 866, 868

Carbon-phenolic composite, 576

Carderock Division, Naval Surface Weapons Center, 157

CARES. See Ceramics Analysis and Reliability Evaluation of Structures (SCARE/

CARES)

Carl, Marion, 30, 297

Carlson, Harry W.: SCAT-15 arrow wing, 67, 

209; sonic boom research, 193, 207, 218; 

sonic boom research, conference and symposia presentations, 210, 211, 212; sonic 

boom research, minimization concepts, 

223–24, 225–26; sonic boom research, 

wind tunnel studies, 208–9, 208n66

Carlswell Air Force Base, 673

Carter, Alan, 518

Case, William R., 476

Castigliano, Carlo Alberto, 466

Castle Bravo nuclear test, 284–85

Caterpillar Tractor Company, 612

Caustics, 226–27, 228

Center for High-Speed Commercial Flight, 

230

Center for Rotorcraft Innovation (CRI), 

169–70

Center for Turbulence Research, Stanford 

University, 341

Center of gravity (cg), 363, 632

Central Intelligence Agency (CIA), 312, 582

Centrifuge, 398–400

Centurion, 850, 856, 859, 862, 864–66

Century series fighters. See also Convair 

F-102 Delta Dagger; Lockheed F-104 

Starfighter; McDonnell F-101 Voodoo; 

North American F-100 Super Sabre; 

Republic F-105 Thunderchief: area rule 

aircraft, 103; damper systems, 365, 366; 

design of, 401, 408–9; management of 

program for, 53n103, 61; roll coupling 

(inertial coupling), 408–9; sonic booms 

from, 185, 187; speed of, 290; tail configuration, 34

Ceramics: ceramic fiber mats, 310–11, 598; 

ceramic matrix composite (CMC) components, 499, 798; ceramic tiles, 419; ceramic 

zirconia, 598; environmental barrier coatings, 612; limitations of, 598; thermal barrier coatings, 610; for X-20 Dyna-Soar, 

587–88

Ceramics Analysis and Reliability Evaluation of Structures (SCARE/CARES), 499–

500; CARES/LIFE, 499–500; CARES/PC, 

499; CMC CARES (CCARES), 499

Ceria, 587

Cerium, 574

Cessna 150 aircraft, 204

Cessna 152 aircraft, 55

Cessna 182 Skylane, 720

Cessna supersonic business jet (SSBJ) 

concept, 249

CF-105 Arrow, 641–42, 642n25, 643

CFL criterion, 436

CFM International, 803

Challenger, 329–30

Chambers, Joseph R., 125

Chamis, Christos, 536

Champine, Robert, 30

Chance Vought nose cap, 306, 587–89

Charles, Prince of Wales, 371

Chemical milling, 569

Chemical vapor deposition (CVD) coatings, 613

Cheranovskiy, Boris Ivanovich, 5

Cheung, Samsun, 234–35

Chevron exhaust nozzle, 808–9

Chicago, 200, 206

Chlorine, 570
NASA’s Contributions to Aeronautics

894

Christian Science Monitor, 107

Chromium, 306, 575, 583, 614

Chrysler Corporation, 498

Circulation Control Rotor (CCR), 156–57

Civil Aeronautics Administration (CAA), 135

Civil aviation. See aviation industry

Civil Tilt Rotor Development Advisory Committee (CTRDAC), 167

CL-400 aircraft, 812–13

CL-1200 Lancer, 34n61

Clark, Larry L., 111

Clark, Thomas L., 475

Clark Atlanta University, 71

Clarke, Mike, 647

Clark MC-2 full-pressure suit, 276, 297, 298

Clayton, New Mexico, 833

Clean Air Act, 792

Clemson University, 542

Cleveland, 105, 137, 191

Climactic Impact Assessment Program 

(CIAP), 789

Climate change research, 789–90, 790n22, 

818

Clinton, William J., 243

Clipper Wind Power, 848

Closed-form analytical solutions, 433, 434

Clough, R.W., 471

Cobalt, 306, 575, 583

Cobalt-based alloys, 573, 574, 575, 610

Coca-Cola bottle design, 93, 106–7

Coen, Peter G., 235, 246, 248, 249, 251, 254

Cold jet hydrogen peroxide reaction

 control, 298

Cole, Julian, 445

Collier Research Corporation, 512

Collier Trophy: Gulfstream business jets, 

121; McDonnell-Douglas innovations, 127; 

NACA for engine and aerodynamic technology, 736; supersonic flight by X-1, 96; 

turboprop propulsion system, 748, 800; 

Whitcomb’s work, 52, 93–94, 107

Colorado, 200

Colorado Springs, 537

Columbia. See Space Shuttle Columbia

Columbium, 312, 584, 585, 587, 595, 602–3

Columbium alloy, 306, 308

Combined-cycle engines, 337

Combustion-Heated Scramjet Test Facility 

(CHSTF), 343

Combustor liners, 798

Combustors, 794–95, 798–99

Command and Service Module, 654

Command Augmentation Systems (CAS), 

366–67, 382

Commercial aircraft. See aviation industry; Supersonic Commercial Air Transport (SCAT) program

Commercial Remote Sensing program, 861

Committee on Technology, National Science 

and Technology Council, 171–72

Common Object Request Broker Architecture (CORBA), 760, 761

Community and Structural Response Program, 197–99

COmposite Blade STRuctural ANalyzer 

(COBSTRAN), 546–47

Composite materials: aircraft parts, 804–5; 

engine fan casings, 767–69; engine parts, 

798; Long Range Planning Study for Composites (RECAST), 804; obstacles to use of, 

804; UAV construction, 866, 868

Composite Primary Aircraft Structures project, 744, 804–5

Composite structures, 534

Compressed air, 324, 573

Compressible flows, 12, 14, 15

Compressors, 746–47, 793–94, 795

Computational fluid dynamics (CFD): 
Index 

895

application of to aircraft design, 428; benefits of, 340, 341; blunt-body problem solution, 439–44; CFL criterion, 436; computer 

codes, 451–53; development of, 340, 427–

28, 453; digital computer development 

and, 436–39; direct numerical simulation (DNS), 342, 439, 450–51; fluid dynamics theory and, 428–33, 434; limitations 

of, 340–42; MacCormack’s method, 446–

47; NASA’s role in development of, 428, 

453; NASP and, 279; Navier-Stokes equations solutions, 438–39, 446–47; propulsion system research, 803; Shuttle study, 

450; sonic boom research, 234–37, 246, 

248, 253, 260; three-dimensional flow-field 

calculations, 447–49; time-marching process, 443–44, 445–46; transonic flow and 

airfoils, 444–46; upwinding, 445

Computational fluid dynamics (CFD) grid 

points: computer technology development 

and, 438–39; direct numerical simulation 

(DNS), 342, 439, 450–51; Euler equations, 

438; flow-field calculations, 433, 434, 435; 

flow-field calculations, three-dimensional, 

446–47; grid generation, 434; numerical 

accuracy, 446–47

Computational simulations. See also simulators and flight simulation: analog computers and, 392–94, 405–7, 468; digital 

computers, 405–7; hybrid computers, 406–

7; Space Shuttle orbiter, 325–26

Computational structural analysis. See finite 

element structural analyses

Computational Structural Mechanics (CSM) 

workshop, 497

Computer-aided design (CAD), 493, 532, 550

Computer codes for computational fluid 

dynamics, 451–53

Computerized Structural Analysis and 

Research (CSAR), 486

Computer Programs Abstracts Journal, 489

Computers: active-element computers, 468; 

direct analog computers, 466–68; equations of motion and, 393; flight control 

system use of, 631, 633n1; Goodyear Electronic Digital Analyzer (GEDA), 393, 394; 

indirect analog computers, 468; supercomputers, 438–39, 446; women mathematicians as, 465, 494

Computers, analog: active-element computers, 468; direct analog computers, 466–

68; flight simulation and, 392–94, 405–7, 

468; fly-by-wire systems, 648–49, 656, 

656n57, 660, 664; hybrid direct-analog/

digital computers, 468; indirect analog 

computers, 468

Computers, digital: batch operation, 437–

38; development of, 436–39, 468–70; flight 

simulation and, 405–7; fly-by-wire systems, 659

Computers, hybrid: direct-analog/digital 

computers, 468; flight simulation and, 

406–7

Computer Sciences Corporation (CSC), 475, 

478, 538

Computer Software Management and Information Center (COSMIC), 477, 488–90, 

490n62, 499, 508

Concorde Supersonic Transport (SST): aluminum and speed, 569; American research 

and, 208; demise of, 249; design of, 3, 45, 

56, 66; economic condition, 229; flight 

control systems, 643; ozone layer damage 

from, 789; service to U.S., 227–28; sonic 

booms from, 227–28

Concurrent development, 322

Congress, U.S.: Civil Tilt Rotor Development 

Advisory Committee, 167; fuel efficiency 
NASA’s Contributions to Aeronautics

896

initiatives, 788, 792; High-Speed Research 

(HSR) program, 232; Joint Planning and 

Development Office (JPDO) authorization, 

171n53; ozone layer research, 789; SCR/

SCAR/AST program funding, 221; sonic 

boom complaints, 187; SST program funding, 218, 219; supersonic research, 219

Conical camber, 51–53

CONMIN program, 492–93, 498, 511, 512

Conservation of energy principle, 432, 433, 

443

Consortium for the Design and Analysis of 

Composite Materials, 500

Continenal Airlines, 123

Continuity equation, 430, 431, 433, 438, 443

Continuous Flow Hypersonic Tunnel 

(CFHT), 283

Contracted research, 153–54

Control-augmentation (CAS) mode, 659, 

660

Control Configured Vehicle (CCV): British 

Jaguar Active Control Technology (ACT), 

670, 690–91, 692; F-16 CCV, 648, 669–70; 

French Dassault Mirage 3NG, 691; German CCV F-104G, 670, 689–90, 692; Japanese CCV T-2, 670, 691–92; purpose and 

benefits of, 651–53, 692

Control coupling, 409–10

Control Data Corporation (CDC): 7600 computer, 324, 325; computers, 469; NASTRAN Users’ Colloquia, 538

Controllable Twist Rotor (CTR) concept, 

153, 156

Control laws, 381–82, 389, 402, 661, 662, 

665, 680–82, 696

Control stick feedback, 370–71, 635, 638–

39, 643

Control-surface actuators: electrical flight 

control surfaces and actuators, 384–85, 

631, 635; electrohydrostatic actuator, 385, 

676–79, 722; electro-mechanical actuator, 385; hydraulically actuated flight control surfaces, 384, 590, 631, 634–35; speed 

of, 383

Control surface commands, 365, 367

Convair Aerospaceplane proposal, 334

Convair B-36 bomber, 55

Convair B-58 Hustler: aeroelastic effects, 

392; B-70 as replacement, 191; crosscountry flight by, 188; design of, 45, 53, 

103; sonic boom research, 194, 195–98, 

199, 200, 202, 203, 204, 205, 209; sonic 

booms from, 188–89; supersonic flight 

by, 188, 189

Convair C-131B aircraft, 204

Convair delta wing aircraft, 46, 46n89, 50

Convair F-102 Delta Dagger: area rule and, 

51, 52, 103; design of, 52–53, 103, 409; 

development and manufacture of, 290; 

F-102A Delta Dagger, 102–3; inertial coupling (roll coupling), 409; management of 

program for, 53n103, 61

Convair F-102 Delta Dagger (YF-102): area 

rule and, 51, 89, 100–102, 103; handling 

qualities, 53; sonic booms, 185; YF-102A 

prototype, 89, 101–2, 103

Convair F-106 Delta Dart: design of, 45, 53, 

409; handling qualities, 50; inertial coupling (roll coupling), 409; purpose of, 103; 

sonic booms, 185, 196, 199, 201, 204, 261; 

variable stability aircraft, 402n62

Convair NC-131H Total In-Flight Simulator (TIFS), 402n62

Convair XF2Y-1 Sea Dart, 45, 103

Convair XF-92A aircraft, 45, 46n90, 49–51, 

100

Convair XP-92 aircraft, 46

Convair YF-102 Delta Dagger: area rule and, 
Index 

897

51, 89, 100–102, 103; handling qualities, 

53; sonic booms, 185; YF-102A prototype, 

89, 101–2, 103

Conventional Take-Off and Landing (CTOL) 

aircraft, 120

Convertible Engine Systems Technology 

(CEST) program, 158, 162

Cook, William L., 476

Cooper, Robert, 337, 339

Cooperative Test Organization, 705–6

Cooper-Harper Handling Qualities Rating 

Scale, 146

Coors Porcelain Company, 590

Copper Canyon research program, 337–38

Copper chill plate, 614

Copper heat shield, 579

Corliss, William, 99

Cornell Aeronautical Laboratory, 401–2, 

402n62

Cornell University, 211, 217, 226

Corona spy satellite, 305

Corrigan, Bob, 846

COSMIC. See Computer Software Management and Information Center (COSMIC)

Couch, James P., 830

Counter Intelligence Agency (CIA), 811

Coupled Eulerian Lagrangian Finite Element (CELFE), 514

Courant, Richard, 435–36, 471

Courant Institute, New York University, 

111, 471

Cowart, Robert A., 253

Cowling design, NACA, 734, 735–36, 809

Craig, Ray, 695

Cranfield Aerospace, Ltd., 71

Cranked arrow wing, 69

Crash-test activities, 152

Cray computers: Cray-1 computer, 325, 438; 

Cray X-MP computer, 438, 446; X-24C 

computed surface streamlines, 449

Cristobalite, 600

Crocco, Arturo, 6

Crossfield, A. Scott, 26, 27–28, 30, 42, 50

Crow, David Ed., 793, 794

Cudnik, Jason, 260

Culebra, Puerto Rico, 827, 833

Cunningham, Herbert J., 474

Curtin, Robert, 856

Curtis, Al, 518

Curtiss, Glenn, 107

Curtiss XP-46 aircraft, 35

Cut-and-try research technique, 113–14

Cyber 205 computer, 446

Cylindrical Optimization of Rings, Skin and 

Stringers (CORSS), 513

Czysz, Paul, 572

D

D-36 columbium alloy, 306, 585, 591

D-558-1 Skystreak, 3, 28–29, 65

D-558-2 Skystreak, 21; engines for, 294; 

inertial coupling experiment, 39; noserise problems, 414; pitch-up problems, 

28–32, 413; pressure suit for, 297; Round 

Two research aircraft, 55n106; tail configuration, 36, 411, 413; wing-stall research, 

414

Daimler-Benz, 701, 706

Dallas, 206

Dallas Convention Center Heliport/Vertiport, 166

Damage-tolerant fan casing, 766–69

Damper systems, 364–67, 412

Dana, William H., 658, 674, 679, 714

Darden, Christine M., 224–25, 229, 235, 

241–42

Dassault Falcon 50 aircraft, 115

Dassault Mirage series: development of, 53; 
NASA’s Contributions to Aeronautics

898

Mirage 3NG, 691; Mirage 2000 aircraft, 

669; Mirage G, 62; Mirage I, 45

Dassault Rafale, 692

Dassault supersonic business jet (SSBJ) 

concept, 249

Data buses, 684–85

David G. Clark Company pressure suits, 

276, 297, 298

David Taylor Naval Ship Research & Development Center, 481, 484, 540, 541

David Taylor Research Center, 541, 542

Da Vinci, Leonardo, 428, 429

Davis-Monthan Air Force Base, 672

Dawson, Virginia, 802

Day, Richard E., 394, 407, 409, 410

Dayton, 638n15

DC-2 aircraft, 6

DC-3 aircraft, 411

DC-4 Generation, 4, 4n1

DC-8 aircraft, 4, 737, 740, 817

DC-9 aircraft, 745

DC-10 aircraft: engine component improvements, 745; Sioux City airliner crash, 757, 

766; winglet research, 119, 120, 122, 123

Decoupled (unconventional) control modes, 

652, 669–70

Deep-stall conditions, 67, 67n127, 413, 

666n70

Deets, Dwain A., 654

Defense, Department of: aircraft safety, 

506; aircraft technology and systems to 

support, 171, 172; flight control systems, 

research activities on, 362–63; Jason advisory group, 227; software development, 

489; SST program, 191, 210–11

Defense Advanced Research Projects Agency 

(DARPA): computational fluid dynamics, 

340; Convertible Engine Systems Technology (CEST) program, 158, 162; Copper 

Canyon research program, 337–38; hypersonic aircraft, 230; NASA, relationship 

with, 139; powder metallurgy, 617; Quiet 

Supersonic Platform (QSP), 244–48, 249; 

Trans Atmospheric Vehicles (TAV), 337; 

X-29 aircraft, 526, 692; X-31 aircraft, 

701; X-51A vehicle, 345; X-Wing Stoppable Rotor program, 157, 158, 163–64, 165

Degree of freedom: unit displacement 

method, 469, 470

De Haven, Hugh, 9n13

De Havilland, Geoffrey, 27

De Havilland D.H. 108 Swallow, 27

Del Frate, John, 855, 858, 858n48, 862

Delta wing aircraft. See swept and delta 

wing aircraft

Denmark, 828, 846–47, 848

Denver, 206

Desert Storm, 62

Design Analysis Methods for Vibration 

(DAMVIBS), 508–11

Deutsche Aerospace, 699–700, 706

Deutsches Museum, 707

Deutsch Metal Components, 539

DFS 346, 36

DFS DM-1, 13

Dielmann, Howard, 476

Differential Maneuvering Simulator (DMS), 

666–67

Digital Array Scanned Interferometer 

(DASI), 861

Digital computers: batch operation, 437–

38; development of, 436–39, 468–70; flight 

simulation and, 405–7; fly-by-wire systems, 659

Digital Electronic Engine Control (DEEC) 

program, 707–10, 754–57, 758–59, 766

Digital Electronic Flight Control System 

(DEFCS), 711
Index 

899

Digital Flight Control for Tactical Aircraft 

(DIGITAC), 382–83, 684–86

Digital flight control system (DFCS), 655, 

672–73, 674, 680–82

Digital fly-by-wire (DFBW) flight control 

systems. See also Digital Fly-By-Wire F-8 

program and aircraft: actuators, 384–85; 

development of, 380, 753; military use of, 

383–84; redundant systems, 384; software 

programs, 381–82, 659–60, 671–72, 683–

84; space flight program use of, 632, 653–

54; testing of, 380–83

Digital Fly-By-Wire F-8 program and aircraft: aircraft modifications for, 380, 654–

56, 658, 661; AP-101 computers, 659, 660; 

Backup Control System (BCS), 656, 657, 

658, 660; benefits of program, 380–81, 632, 

662–64; control laws, 381–82, 661, 662; 

exhibit showing, 663; flight control modes, 

659–60; flight-testing, 647, 657, 661–62, 

753; follow-on research testing, 381–82; 

gain, 656; implementation of Phase I, 655–

58; implementation of Phase II, 658–62; 

Iron Bird simulator, 655, 656–57, 659, 660; 

Linear Differential Variable Transformers 

(LDVTs), 656; performance of, 663–64; 

preflight safety procedures, 383; redundant systems, 380, 655–56, 658, 659; Resident Backup System (REBUS), 662; side 

stick flight controller, 658, 658n59; simulated Shuttle landing flights, 661; software programs, 381–82, 659–60; transport 

delay, 661; YF-16 design validation, 665

Digital Matrix Abstraction Programming 

(DMAP), 476

Digital Systems Validation Handbook (FAA), 

684

Digital terrain system with moving map, 

675–76

Dihedral effects, 19, 34n62, 390, 410–11

Direct analog computers, 466–68

Direct Connect Supersonic Combustion Test 

Facility (DCSCTF), 343

Direct control system, 640, 653, 659, 660

Direct current power generation system, 

677–78

Directionally solidified (d.s.) castings, 

614–15

Directional stability: flattop delta aircraft, 

55; fuselage length and, 37–38; inertial 

coupling (roll coupling) and, 38–39, 40–41; 

static, 408, 409, 410–11, 648, 651; swept 

and delta wing aircraft, 19; wedge-shaped 

tail and, 283

Direct numerical simulation (DNS), 342, 

439, 450–51

Discoverer II vehicle, 305, 582

Discretization of structures, 466, 469, 

470–72

Display and Keyboard (DSKY), 655, 655n55

Distinguished Service Medal, 107

Distortion Estimation System (DES), 

717–18

Divone, Louis, 825

Donlan, Charles, 59, 603

Doolittle, James H., 737

Dornberger, Walter, 290, 300

Dornier Do-31 aircraft, 648–49

Dorsch, Robert, 331

Doublet-lattice model, 482–83, 526

Douglas, Frank J., 475

Douglas A4D-1 Skyhawk, 45–46

Douglas A-4 Skyhawk, 65

Douglas Aircraft Company. See also McDonnell-Douglas Corporation: delta wing 

development, 46–47; High-Speed Civil 

Transport (HSCT), 231, 232, 233; winglet use by, 121–22
NASA’s Contributions to Aeronautics

900

Douglas B-26 Invader, 402n62

Douglas C-47 transport, 195, 411

Douglas D-558-1 Skystreak, 3, 28–29, 65

Douglas D-558-2 Skyrocket: engines for, 

294; inertial coupling experiment, 39; 

nose-rise problems, 414; pitch-up problems, 28–32, 413; pressure suit for, 297; 

Round Two research aircraft, 55n106; 

tail configuration, 36, 411, 413; wing-stall 

research, 414

Douglas D-558 FSW planform, 21

Douglas DC-2 aircraft, 6

Douglas DC-3 aircraft, 411

Douglas DC-8 aircraft, 4, 740

Douglas F4D-1 Skyray, 48n94

Douglas F5D-1 Skylancer, 591

Douglass, R.R., 843–44

Douglas X-3 “Stiletto,” 3, 4, 37, 40, 41–42, 

43, 440

Douglas XF4D-1 Skyray, 32–33

Dow, Norris, 291, 292

Dow Corning, 314

Drake, Hubert, 63

Draper, Charles Stark, 38, 284

Draper Laboratory, Massachusetts Institute 

of Technology (MIT), 654

Drinkwater, Fred, 641

Driver, Cornelius “Neil,” 220

Drones for Aerodynamic and Structural 

Testing (DAST), 388–89

Dryden, Hugh L., 101

Dryden Flight Research Center (DFRC)/

Flight Research Center. See also HighSpeed Flight Research Station/HighSpeed Flight Station: ablation cooling 

research, 417; Advanced Fighter Technology Integration (AFTI) F-16 project, 673–

79; Advanced Turboprop (ATP) project, 

800; aeroservoelastic effects, 496; Boeing 

X-48B testbeds, 72; Centurion, 866; computers at, 468; Digital Electronic Engine 

Control (DEEC) program, 708–10, 754–

56; Digital Fly-By-Wire F-8 program and 

aircraft, 655, 657; Electrically Powered 

Actuator Design (EPAD) program, 384; 

Environmental Research Aircraft and 

Sensor Technology (ERAST) program, 

849, 853, 855, 856, 862, 866, 867, 868–

69; finite element analysis activities, 494–

96; FLEXSTAB program, 544–45; flutter 

suppression research, 388; fly-by-wire 

research activities, 383, 653, 654, 753; 

General Dynamics F-16XL research, 69, 

72, 237, 242–43, 245; General Purpose 

Airborne Simulator, 402n62; Grumman 

F-14A Tomcat testing, 63; Helios Prototype, 867, 868–69; high-altitude aircraft 

workshop, 850; Highly Maneuverable Aircraft Technology (HiMAT), 686–89; High 

Temperature Loads Calibration Laboratory, 553; hypersonic flight research, 299; 

inertial coupling research, 408–9; Inlet 

Spillage Shock Measurement (ISSM), 

246; Integrated Propulsion Control System (IPCS) program, 754; Intelligent 

Flight Control Systems (IFCS) project, 

719; laminar flow control research, 495, 

784, 808; limit-cycle oscillation research, 

368–69; Lunar Landing Research Vehicles 

(LLRVs), 653; NASTRAN development, 

474; Performance Seeking Control (PSC) 

program, 712; pilot-induced oscillations 

research, 381; Propulsion Controlled Aircraft (PCA) program, 714–15, 757, 758; 

rotary wing aircraft research, 164, 165, 

169–70; Self-Repairing Flight Control 

System (SRFCS), 713–14; Shuttle concept development, 57; simulators at, 296; 
Index 

901

solar cells, 862; sonic boom research, 195–

97, 201–4, 227, 227n133, 236–39, 242–43, 

245, 251, 252–53, 255–57, 258–60; Structural Analysis Routines (STARS), 556; 

structural resonance research, 376–77, 

378–79; supercritical wing program, 115, 

805–6; supercritical wing testing, 112–13; 

supersonic flight research, 193; synthetic 

fuel research, 816–17; Thermal Loads 

Laboratory, 553–55; thermal research, 

495–96; tilt rotor aircraft, 160; Transonic 

Aircraft Technology (TACT) program, 

386; variable-geometry wing research, 

63; X-15 aircraft, 294, 395–97; X-29 aircraft, 526, 694–98, 699; X-31 aircraft, 

701, 704–5, 707; X-43 Hyper-X vehicle, 

344–45; X-Wing Stoppable Rotor testing, 

164; YF-12 Blackbird project, 517; YF-12 

Experiment Symposium, 519

Duits-Nederlandse Wind Tunnel (DutchGerman Wind Tunnel), 164

Dulles Airport, 166, 638n15

Dunagan, Steven, 861

Dunne, John, 4–5

DuPont, 316, 320–21, 767

DuPont, Anthony “Tony,” 335, 337, 338, 339

Dutch-German Wind Tunnel (Duits-Nederlandse Wind Tunnel), 164

Dutch roll motions, 19, 50, 410–12, 637, 

637n11

DYLOFLEX program, 545

Dynacs Engineering, 539

Dynamic Analysis System, 153

Dynamic Engineering, Inc. (DEI), 481

Dynamic Simulation of Controls & Structure (DISCOS), 548–49

Dyson, Ken, 700

E

Eagle Engineering/Aeronautics, 235, 238, 

246, 249

EAL program, 498

Earth Resources Survey ER-2 aircraft, 850, 

851, 861

Earth Science Enterprise, 849

EBASCO Services, 538

Eckert, J. Presper, 436

E Core engine, 803

Eddy viscosity, 341

Edison, Thomas A., 91, 98

Edisonian research technique, 113

Edwards, George, 56

Edwards, Thomas A., 234–35

Edwards Air Force Base. See also Dryden 

Flight Research Center (DFRC)/Flight 

Research Center: Air Force Flight Test 

Center (AFFTC), 44, 185, 195–96; deepstrike fighter competition, 69; fly-by-wire 

testing, 753; Helios Prototype, 866n61, 

867; High-Speed Flight Research Station (HSFRS), 184, 185; landings at, 403; 

Pathfinder, 858; Shaped Sonic Boom 

Demonstration (SSBD), 247; sonic boom 

research, 193, 195–97, 200, 201–4, 206–7, 

227, 229–30, 251; sonic booms over, 226; 

Survivable Flight Control System program (F-4 SFCS program), 646–47; X-2 

aircraft, 394–95

Eggers, Alfred J., Jr.: B-70 design, 191; 

blunt-body theory, 55, 285–86, 291, 

311, 415; flattop delta aircraft, 54–55, 

56, 56n107; hypersonic tunnel, 283; 

HYWARDS, 300–301

Eglin Air Force Base, 187

Ehernberger, L.J., 236

Eindecker (Monoplane), 34

Eisenhower, Dwight D., 191, 288, 582
NASA’s Contributions to Aeronautics

902

Electric airplane, 815

Electrical flight control surfaces and actuators, 384–85, 631, 635, 722

Electrically operated trim tabs, 633, 634–35

Electrically Powered Actuator Design 

(EPAD) program, 384–85

Electrohydrostatic actuators (EHAs), 385, 

676–79, 722

Electro-mechanical actuator (EMA), 385

Electron beam physical barrier vapor deposition (EB-PVD) coatings, 610, 612, 613

Electronic damper systems, 364–67, 412

Electronic engine control systems, 631, 632, 

633n1

Electronic flight control system. See flight 

control systems

Electro-optical flight deflection measurement system, 521–22, 527–28

Elevators, 362–63, 633, 635

Elevons, 9, 50, 278, 391

Elliptic behavior, 443

Elliptic partial differential equations, 435, 

435n11

Emissions: Advanced Subsonic Technology 

(AST) program, 796–97, 798; Atmospheric 

Effects of Aviation project, 797; carbon 

dioxide emissions, 790, 790n22, 797–98, 

799, 813; HSCT aircraft, 795–96; nitrogen oxides (NOx) emissions, 762, 790, 794, 

795–99, 813; ozone layer damage from, 

789–90, 795, 798; reduction in, 785–86, 

788–89; regulations and standards, 745, 

791–92, 794, 796, 797, 798, 799; Ultra Efficient Engine Technology (UEET) program, 762–66; Zero CO2 project, 813

Emissions Reduction project, 763

Enabling Propulsion Materials program, 

767

Endplates, 118

Energy, conservation of, 432, 433, 443

Energy, Department of (DOE), 439, 824, 

825, 836–39, 846, 847

Energy Efficient Engine (E Cubed) project, 742, 744, 745–47, 786, 788, 790–91, 

792–95, 817

Energy Efficient Transport (EET) project, 

123, 744, 805–6

Energy maneuverability concept, 651

Energy method, 466

Energy Research and Development Administration (ERDA), 824, 825

Energy Technology Division, 830

Enevoldson, Einar K., 658, 661

Engine Component Improvement project, 

744–45, 746, 747

Engineering Dynamics, Inc., 551

Engine inlet unstarts, 336, 754

Engine mufflers, 737, 740

Engines. See also turbine blades; specific 

types of engines: combustor liners, 798; 

combustors, 794–95, 798–99; component parts, improvements to, 794, 798; 

compressors, 746–47, 793–94, 795; damage-tolerant fan casing, 766–69; design 

of, computer programs for, 746–47, 748; 

design of, NASA initiatives for, 791; digital engine control systems, 754–57, 758–

59; electronic engine control systems, 631, 

632, 633n1; emissions standards, 745; 

Energy Efficient Engine (E Cubed) project, 742, 744, 745–47, 786, 788, 790–91, 

792–95; materials for, 584; noise research, 

736–37; polymer composite engine fan 

casings, 767–69; Structural Tailoring of 

Engine Blades (STAEBL), 545–46; Ultra 

Efficient Engine Technology (UEET) 

program, 613, 762–66, 786, 790, 797–98; 

unstarts, 336, 708–9, 754
Index 

903

Engines, aircraft noise and improvements to, 737–40; Acoustically Treated 

Nacelle program, 740; Quiet Clean Short 

Haul Experimental Engine, 741–43; 

Quiet Engine program, 740, 741; Steep 

Approach program, 740

Engine Structural Dynamics Program, 555

England. See Great Britain

English Electric Lightning, 32

Enhanced Fighter Maneuverability (EFM) 

X-31 aircraft, 699–705

ENIAC, 436, 468–69

Eniwetok Atoll, 439

ENSAERO code, 493–94

Enterprise. See Space Shuttle Enterprise

Envia, Edmane, 770, 771, 772

Environmental barrier coatings (EBCs), 

612–13

Environmental impact statement (EIS), 226

Environmental Protection Agency (EPA): 

emissions standards, 745, 791–92, 794; 

Environmental Science Services Administration, 206

Environmental Research Aircraft and Sensor Technology (ERAST) program. See 

also Unmanned aerial vehicles (UAVs): 

alliance organizational structure, 853–

55, 872; budget and funding, 853, 854–55; 

goals of, 849–50, 852, 861, 872; high-altitude aircraft, requirements and capabilities of, 851–52; high-altitude aircraft, 

workshop on, 850–52; Joint Sponsored 

Research Agreement (JSRA), 849, 852–

55, 852n33, 872; ownership of hardware 

developed, 854; participants and partners, 

853–55; report requirements, 853; science 

missions, 859, 861; sensor technology 

development, 861; success of, 855, 871–72

Environmental Science Services

 Administration, 206

Epstein, Alan H., 750–51

Equations. See also Euler equations; NavierStokes equations: continuity, 430, 431, 

433, 438, 443; elliptic partial differential, 435; hyperbolic partial differential, 

435n11, 436, 443; integral, 433; linear partial differential, 431, 435–36; of motion, 

393; nonlinear partial differential, 430–

31, 432–33; parabolic partial differential, 

435n11, 436; partial differential, 430–31, 

432–33, 434–36, 435n11, 446–47

ER-2 aircraft, 850, 851, 861

Error Correction Information System 

(ECIS), 479

ESA 3560 HF ablative blanket, 607

Ethell, Jeffrey, 805

Ethylene, 346

Ettinger, Robert C., 647

Euler, Leonhard, 430–31

Euler equations: blunt-body problem solution, 443–44; continuity equation, 430, 

431, 433, 438, 443; ENSAERO code, 

493; friction and, 431; LAURA code, 451; 

momentum equation, 430, 431, 433, 438, 

443; transonic flow and airfoils, 445

Eurofighter, 692

European countries interest in NASTRAN, 

481, 484

Everest, Frank K. “Pete,” 51, 102, 297

Exceptional Service Medal, 107

Exhaust nozzle chevron, 808–9

Exhaust nozzles, thrust-vectoring, 715–16, 

717

Experimental Aircraft Program (EAP), 692

Experimental Air Vehicle (EAV), 720

Experimental Clean Combustor program, 

794–95

External tanks, 326, 327, 329, 578
NASA’s Contributions to Aeronautics

904

Extremely short takeoff and landing 

(ESTOL), 705–7

Exxon, 538

F

F-2 fighter, 692

F4D-1 Skyray, 48n94

F4H-1 Phantom II, 32, 33n60, 34, 195

F-4 Phantom, 644n32

F-4 Survivable Flight Control System program, 379–80, 384, 646–48, 658n59, 660

F4U-5 Corsair, 401

F5D-1 Skylancer, 591

F-5 fighter: F-5E Shaped Sonic Boom Demonstration, 245–48, 250, 252; sonic boom 

research, 245; tail configuration, 32, 413

F6F-5 Hellcat, 401, 637

F-8A fighter, 112–13

F-8 Digital Fly-By-Wire aircraft. See Digital Fly-By-Wire F-8 program and aircraft

F-8 Super Cruise Wing project, 655

F8U Crusader series. See also Digital Fly-ByWire F-8 program and aircraft: area rule 

and redesign of, 103–4, 105; digital flight 

control system, 380–81, 753; F-8C, 380–

81, 654–55, 753; F-8 Supercritical wing, 

89; F-8 Super Cruise Wing project, 655; 

F8U-1, 32, 37, 43; F8U-3, 194; XF8U-1 

Crusader, 37, 363

F9F-2 Panther, 638–39, 642, 658n59

F9F-5 Panther, 33

F9F-6 Cougar, 33

F9F-7 aircraft, 33n60

F9F-9 Tiger, 104, 105, 106

F11F-1 Tiger, 32, 104, 105

F-14 Tomcat, 44, 62, 413; F-14A Tomcat, 63

F-15 aircraft, 710. See also McDonnellDouglas F-15 Eagle

F-15 Eagle: Advanced Control Technology 

for Integrated Vehicles Experiment 

(ACTIVE) project, 716, 717; Digital Electronic Engine Control (DEEC) program, 

709–10, 754–56; Dryden research activities, 496; engines, 713; F-15A, 44, 711, 

713; F-15E Strike Eagle, 69; fly-by-wire 

system, 383–84; Highly Integrated Digital Electronic Control (HIDEC), 711; High 

Stability Engine Control (HISTEC) system, 718; Intelligent Flight Control Systems (IFCS) project, 718–21; Propulsion 

Controlled Aircraft (PCA) program, 713, 

757; Self-Repairing Flight Control System 

(SRFCS), 713–14; simulator, 407; sonic 

boom research, 238–39, 246–48, 252–53, 

256, 258–60, 261; tail configuration, 413; 

TF-15, 715

F-15 Eagle (F-15B/NF-15B): Advanced 

Control Technology for Integrated Vehicles Experiment (ACTIVE) project, 716, 

717; exhibit showing, 721; High Stability 

Engine Control (HISTEC) system, 718; 

Intelligent Flight Control Systems (IFCS) 

project, 718–21; Lift and Nozzle Change 

Effects on Tail Shocks (LaNCETS), 258–

60, 261; modifications for experiments, 

258–59; Quiet Spike device, 252–53, 256; 

Short Takeoff and Landing Maneuver 

Technology Demonstrator (S/MTD), 715–

16; standard features, 259

F-16 Fighting Falcon: deep-stall conditions, 666n70; development of. See General Dynamics YF-16 Lightweight Fighter 

(LWF)Digital Electronic Engine Control 

(DEEC) program, 710, 756; F-16A, 378, 

380, 669; F-16B, 380, 669; F-16C, 669; F-16 

CCV, 648, 651–52, 669–70; F-16XL-1, 808; 

F-16XL-2, 784, 808; F-16XL SCAMP, 69, 

69n129, 72, 237, 242–43, 245, 261; F-16XL 
Index 

905

SCAMP model, 807; fly-by-wire systems, 

380, 383–84, 669; pitch rocker technique, 

666n70; production of, 669; success of, 

668–69; tail configuration, 413, 666n70

F-18 Hornet, 237, 251, 255–57, 496, 676

F-22 Raptor, 126–27, 384, 402, 413, 712, 

713, 722

F-35 Lightning II Joint Strike Fighter, 384, 

413, 678, 713, 722

F-84 aircraft, 333

F-86 Sabre: accidents with, 24, 24n43; 

development of, 19n32, 23n39; F-86A, 

26; F-86D Sabre, 102; F-86F, 26; Korean 

war use of, 4, 19, 26; pitch-up problems, 

24–26, 36, 37, 413; reputation of, 23–24; 

sonic booms, 186–87; stability research 

activities, 637; tail configuration, 36, 413; 

variable stability aircraft, 401; XP-86 

Sabre prototype, 18, 635; YF-86D low-tail 

testbed, 36–37

F-89 Scorpion, 102

F-94 Starfire, 102

F100 engine model derivative (EMD) 

engine, 711

F-100 Super Sabre: development and manufacture of, 290; F-100C Super Sabre, 

409; F-100D Super Sabre, 409; flight control systems, 635; JF-100C Super Sabre, 

402n62; landing simulator, 403; sonic 

boom research, 194, 196; tail configuration, 413; YF-100A Super Sabre, 34, 363; 

YF-100 Super Sabre, 37, 40, 41, 42, 185

F-100 Super Sabre (F-100A): design of, 41, 

42–43, 393–94; roll-coupling accident, 41, 

42, 44, 393; roll coupling research, 408–

9; wingspan-to-fuselage length ratio, 37

F100 turbojet engine, 710, 712–13, 717, 

754–56

F-101 Voodoo: development and 

manufacture of, 290; flight control system, 

372; pitch-up problems, 33–34, 413; sonic 

boom research, 194, 199; sonic booms, 

185; tail configuration, 33–34, 36, 413; turbofan engines, 741

F-102 Delta Dagger: area rule and, 51, 52, 

103; design of, 52–53, 103, 409; development and manufacture of, 290; F-102A 

Delta Dagger, 102–3; inertial coupling (roll 

coupling), 409; management of program 

for, 53n103, 61

F-102 Delta Dagger (YF-102): area rule and, 

51, 89, 100–102, 103; handling qualities, 

53; sonic booms, 185; YF-102A prototype, 

89, 101–2, 103

F-103 interceptor fighter, 569; XF-103 interceptor fighter, 289, 290, 333

F-104 Starfighter: design of, 104, 401, 409, 

440; development and manufacture of, 

290; F-104A Starfighter, 37, 440; F-104B, 

409; F-104G, 688; F-104G (German CCV), 

670, 689–90, 692; heating, structural, 413; 

inertial coupling, 43, 409; landing of, 405; 

landing simulator, 403–4; nickname for, 4, 

104; pitch-up problems, 34n61, 413; reaction controls, 298; sonic boom research, 

195–96, 199, 200, 204, 205, 206–7; speed 

of, 413; supersonic flight by, 3, 104; use 

of, 104; wingspan-to-fuselage length ratio, 

407; XF-104 Starfighter, 185

F-105 Thunderchief: design of, 333, 409; 

development and manufacture of, 290; 

F-105B Thunderchief, 32, 44; inertial coupling (roll coupling), 409; sonic booms, 

185, 216; tail configuration, 37, 413

F-106 Delta Dart: design of, 45, 53, 409; 

handling qualities, 50; inertial coupling 

(roll coupling), 409; purpose of, 103; 

sonic booms, 185, 196, 199, 201, 204, 261; 
NASA’s Contributions to Aeronautics

906

variable stability aircraft, 402n62

F-107 aircraft, 372, 642, 658n59

F-111 aircraft: adaptive flight control systems (AFCS), 373–74, 375; design of, 

61–62, 62n118; digital engine control system, 754; F-111A variable-sweep aircraft, 

113, 386, 404; F-111F, 61; fly-by-wire system, 639; management of program for, 61; 

mission-adaptive wing (MAW), 386; sonic 

boom research, 238

F117 engine, 795

F-117 stealth fighter, 70, 384, 721–22

F119 engine, 712, 713

F135 engine, 713

F404 engine, 802

F/A-18 Hornet: Active Oscillation Suppression (AOS) system, 389; autothrottle system from, 706; design of, 669; engines for, 

706n152; F/A-18A Hornet, 44, 383–85; F/A18B Hornet, 386; tail configuration, 413

FAA. See Federal Aviation Agency/Administration (FAA)

Faget, Maxime “Max,” 56, 291, 602, 603

Fail-operational, fail-operational, fail-safe 

(FOFOFS) systems, 367, 384

Fail-operational, fail-operational (FOFO) 

system, 379–80

Fail-operational, fail-safe (FOFS) systems, 

366

Fairey F.D.2 aircraft, 45, 53

Falling body test, 38, 43, 92

Fallon, 247

Fanjet engines, 747

Farmer, Crofton, 851

Farnborough, England, 161, 643

FASTOP program, 527

FB-111 aircraft, 392

Fedden, Roy, 35n65

Federal Aviation Agency/Administration 

(FAA): aircraft noise standards, 739–40, 

771; aircraft safety, 506; aircraft technology and systems to support, 171, 172; 

certification of MD-11, 123; certification 

programs, 123, 683–84, 723; Civil Aeronautics Administration (CAA), 135; Digital 

Systems Validation Handbook, 684; emissions regulations, 745; International Helicopter Study Team (IHST), 172; NASA, 

relationship with, 139; National Rotorcraft 

Technology Center, 168; polymer composite engine fan casings, 768; recertification 

of DC-10, 123; rotary wing noise research, 

163, 164; Skyshield air defense exercises, 

188; SST program, 65–66, 191–92, 210–

11; Supersonic Transport Development 

Office, 199; tilt rotor aircraft, civil transportation use of, 166–67; unmanned aerial vehicles (UAVs) issues, 854

Federal Aviation Agency/Administration 

(FAA) and sonic booms: noise level standards, 250; noise standards workshop, 

248, 249; prohibition of by Concorde, 227–

28; reactions to, 261; research activities, 

199–201, 206, 207–8, 251; restrictions on, 

248–49; supersonic flight ban and, 220–

21, 220n112, 224

Federal Trade Commission (FTC), 486

Federal Wind Energy Program, 823–25, 827, 

828–29, 831, 833, 840, 843, 846–49

Federation of American Scientists, 227

Ferri, Antonio, 213, 215–16, 217, 222, 223, 

332–33, 334

Fiberglass, 804, 845

Fiber-optic controllers, 385, 685, 686

Finite-difference solutions, 434–36

Finite Element Composite Analysis Program 

(FECAP), 556

Finite Element Modeling of Piping Systems 
Index 

907

program, 552

Finite element structural analyses. See also

NASA Structural Analysis System (NASTRAN): Ad Hoc Group for Structural Analysis, 474; advancements in, 152, 530–34; 

aviation industry use of, 533–34; benefits of, 534–35; computational analysis 

or simulation, 515, 516; discretization of 

structures, 466, 469, 470–72; finite element method (FEM), 471–72, 485, 510; 

flight-testing, 514, 515, 534–35; HiMAT, 

515, 520–22, 530; laboratory testing, 515; 

materials behavior and, 532–33; modeling 

process, 531–32, 534–35; NASA’s contributions to, 535, 556; oblique wing concept, 515, 522–26, 530; optimization of 

aircraft design, 533; pre- and post-processing challenges, 531–32; rotary wing 

aircraft, 162–63, 515, 519–20, 529–30; software programs, 158, 490n62, 502, 544–

52; Space Shuttle studies, 327–28, 506; 

technology development for, 486; testing 

facilities, 553–55; theory, 515, 516; validation of, 530, 534–35, 553–55; X-29 aircraft, 515, 526–29, 530; YF-12 Blackbird 

project, 515, 516–19, 529

Finite element structural analyses, projects and contributions of Centers, 490–91, 

491n64; Ames Research Center, 491–94; 

Dryden Flight Research Center, 494–96; 

Glenn (Lewis) Research Center, 496–500, 

532; Goddard Space Flight Center, 500–

502; Jet Propulsion Laboratory, 502–5; 

Johnson Space Center, 505–7; Langley 

Research Center, 507–12, 532; Marshall 

Space Flight Center, 512–14

Finnegan, Patrick M., 830

Fischel, Eduard, 636

Fischer-Tropsch process, 816n114, 817

Fischetti, Thomas L., 105

Fisher Body Company, 91

Flaperons, 694, 697

Flat-bottom delta aircraft, 55–57, 301, 302

Flat-top delta aircraft, 55–56, 56n107, 

300–301

Flechner, Stuart G., 119, 120

Fletcher, Edward, 331

Fletcher, James, 116

Flexibility matrix, 471

FLEXSTAB program, 392, 518, 544–45

Flight control coupling, 409–10

Flight control surfaces. See also control-surface actuators; specific surfaces: electrical flight control surfaces and actuators, 

384–85, 631, 635, 722; hydraulically actuated flight control surfaces, 384, 590, 631, 

634–35, 643, 644

Flight control systems. See also digital flyby-wire (DFBW) flight control systems; 

fly-by-wire (FBW) systems: adaptive flight 

control system (AFCS), 372–74, 375, 376, 

379; advancements in, 278, 752–53; aerodynamic damping and, 363–67; artificial 

feel devices, 635, 643; Command Augmentation Systems (CAS), 366–67, 382; 

computer use for, 631, 633n1; control 

stick feedback, 370–71, 635, 638–39, 643; 

design of, frequency spectrum of challenges, 377, 378; direct control system, 

640, 653, 659, 660; electronic feedback 

flight control research, 636–37; evolution 

of, 632–35, 633n1; fly-by-light fiber-optic 

digital flight control system, 685, 686; 

gain, 656; hydraulically actuated flight 

control surfaces, 384, 590, 631, 634–35, 

643, 644; hypersonic flight, 298; limit-cycle 

oscillation, 367–69, 377, 378; mechanical linkages, 525, 631, 633, 637–38, 639, 
NASA’s Contributions to Aeronautics

908

655; Minneapolis Honeywell MH-96 

flight control system, 298–99, 372–74; 

model-following systems, 366–67; pilotinduced oscillations (PIO), 369–72, 378, 

381, 409, 412, 661; redundant systems, 

365–67; research activities, 362–63; resonant motions, 374–79, 513; Self-Repairing Flight Control System (SRFCS), 

713–14; side stick flight controller, 298, 

638, 638n15, 642, 658, 658n59, 664, 665, 

668; Space Shuttle orbiter, 367; stability 

augmentation systems, 364–66, 399, 636–

37; total loop gain, 367–69, 372; transport 

delay, 661; Vietnam war and, 644–45; X-1 

aircraft, 362–63; X-2 aircraft, 298

Flight envelope limitations, 368, 632

Flight gain ground rule, 376–77

Flight Operations, 494

Flight Research Center. See Dryden Flight 

Research Center (DFRC)/Flight Research 

Center

Flight Simulator for Advanced Aircraft 

(FSAA), 151

Flight simulators. See simulators and flight 

simulation

Flittie, Kirk, 856

Floating point operations per second 

(FLOPS), 341, 437, 438–39

FLO codes, 445–46

Florida, 226

Flourine, 570

Flow direction fences, 414

Fluid dynamics, 428–33. See also computational fluid dynamics (CFD)

Flush Air Data System, 707

Flutter, 306, 309, 482–83, 525–26, 585–86, 

595

Fly-by-light fiber-optic digital flight control 

system, 685, 686

Fly-by-wire (FBW) systems: acceptance of, 

379; actuators, 384–85; advancements in, 

631; advantages of, 631–32; aeroservoelastic effects, 496; analog computers, 648–

49, 656, 656n57; angle of attack limiters, 

665–66; asynchronous digital flight control system, 672, 680–82, 683; Augmented 

Longitudinal Control System (ALCS), 

372; aviation industry use of, 632, 722–

23; backup systems, 656, 657, 658, 660, 

662, 688; benefits of, 650–53, 654, 721–23; 

certification of new aircraft, 683–84; control stick feedback, 370–71; development 

of, 379–80, 631, 632, 635–37, 753; digital 

flight control system, 655, 672–73, 674, 

680–82; early applications, 637–43; events 

that supported need for, 643–45; nonaerospace applications, 631, 723; redundant 

systems, 379–80, 655–56, 658, 659, 671–

72; research activities at Dryden, 383, 

753; research activities at Edwards, 646–

47; research activities at Langley, 638–39; 

research activities by Air Force, 645–48; 

research activities in Europe, 648–50; software programs, 381–82, 659–60, 671–72, 

683–84; space flight program, transfer of 

technology from, 632, 653–54; Survivable 

Flight Control System program (F-4 SFCS 

program), 379–80, 384, 646–48, 658n59, 

660; transition to, 632; transport delay, 661

Fly-by-wire (FBW) systems, digital: actuators, 384–85; development of, 380, 753; 

military use of, 383–84; redundant systems, 384; software programs, 381–82, 

659–60, 671–72, 683–84; space flight 

program use of, 632, 653–54; testing of, 

380–83

Fly-by-wire ship control system certification program, 723
Index 

909

Flying Laboratory for Integrated Test and 

Evaluation (FLITE), 170

Flying-quality criteria, 145–46

Flying wings (Nurflügeln), 70

Focke-Wulf FW 190 fighter, 633n1, 636, 

703n144

Focke-Wulf Ta 183 aircraft, 36

Fokker, Anthony, 385

Fokker Eindecker (Monoplane), 34

Folland Gnat, 32

Forest Service, U.S., 200

Forman, Royce, 506

Fort Eustis, Virginia, 146

FORTRAN programming language, 437, 

469, 476

Fort Worth, 667, 674, 678

Forward-looking infrared (FLIR) device, 

673, 675

Forward-swept wing (FSW), 21–22, 524, 

527, 693–94

France: Concorde Supersonic Transport 

(SST). See Concorde Supersonic Transport 

(SST)Control Configured Vehicle (CCV), 

691; delta wing development, 45, 53; Paris 

Air Show, 161; Supersonic Commercial Air 

Transport (SCAT) program, 190; V-2 missile shock waves, 184; variable-geometry 

wing research, 62

Frankford Arsenal, 542

Franklin Institute, 142

Free body diagrams, 464–65

Freon gas, 149, 150

Frequencies and Modes of Shells of Revolution (FAMSOR), 506

Friction, 431–32

Friedrichs, K.O., 435–36

Fuel: alternative, 785, 815, 817; biofuel, 

817; Fischer-Tropsch process, 816n114, 

817; hydrogen, liquid, 809–14, 815–16; oil 

prices, 743, 788, 790–91, 816, 840–41; resonant motions from, 513; synthetic, 815–

17, 816n114; transfer between tanks, 632

Fuel cells: for aircraft, 785, 814–15; hydrogen-air fuel cell system, 864, 869, 871; 

hydrogen fuel cell aircraft, 815; hydrogen-oxygen regenerative fuel cell system 

(RFCS), 863–64, 871; regenerative fuel 

cell system (RFCS), 863–64, 871, 872; for 

UAVs, 856, 863–64, 869, 871

Fuel efficiency programs, 785–86, 790–91, 

817. See also Aircraft Energy Efficiency 

(ACEE) program

Full Authority Digital Engine Control program, 755

Full-cost accounting methods, 166

Fullerton, Gordon, 715

Fundamental Aeronautics Program, 254

FW 190 fighter, 633n1, 636, 703n144

G

Gain, 656

Galileo, 428, 429

Galileo probe, 277, 278, 573

Garrett AiResearch, 335

Garrett Turbine Company, 612

Garrison, Charles P. “Pete,” 646

Gas-metal-arc (GMA) welding, 575

GASP 2.0 / 3.0, 341

Gas-tungsten-arc (GTA) welding, 575

Geared Turbofan engine, 750, 751–52, 800–

801, 802

Gedser wind turbine, 827, 828, 846–47

Geisler, Frank, 645

Gemini program: ablation cooling, 417; 

centrifuge program, 400; flight control 

systems, 653; fuel cells, 814; reentry 

vehicles, 54; side stick flight controller, 

658n59; spacesuits, 297; Titan II missile/
NASA’s Contributions to Aeronautics

910

booster, 328

General American Research Division, 215

General Applied Science Laboratories 

(GASL), 215–16, 333, 336, 338, 343

General Atomics Aeronautical Systems, 853

General Aviation Manufactures Association, 248

General Dynamics, 480

General Dynamics F-16 Fighting Falcon: 

deep-stall conditions, 666n70; development of. See General Dynamics YF-16 

Lightweight Fighter (LWF)Digital Electronic Engine Control (DEEC) program, 

710, 756; F-16A, 378, 380, 669; F-16B, 380, 

669; F-16C, 669; F-16 CCV, 648, 651–52, 

669–70; F-16XL-1, 808; F-16XL-2, 784, 

808; F-16XL SCAMP, 69, 69n129, 72, 237, 

242–43, 245, 261; F-16XL SCAMP model, 

807; fly-by-wire systems, 380, 383–84, 669; 

pitch rocker technique, 666n70; production of, 669; success of, 668–69; tail configuration, 413, 666n70

General Dynamics F-111 aircraft: adaptive flight control systems (AFCS), 373–

74, 375; design of, 61–62, 62n118; digital 

engine control system, 754; F-111A variable-sweep aircraft, 113, 386, 404; F-111F, 

61; fly-by-wire system, 639; management 

of program for, 61; mission-adaptive wing 

(MAW), 386; sonic boom research, 238

General Dynamics FB-111 aircraft, 392

General Dynamics Tactical Fighter

 Experiment (TFX), 61

General Dynamics YF-16 Lightweight 

Fighter (LWF), 371; angle of attack limiters, 665–66; flight-testing, 667–68; flutter 

suppression research, 388, 389; fly-bywire system, 379–80, 647, 664–69; pilotinduced oscillations, 370; roll stick force 

gradient, 667; side stick flight controller, 658, 664, 665, 668; simulator studies, 

666–67; strain gages, 385–86; structural 

resonance research, 378; success of, 668; 

unintentional first flight, 667

General Electric (GE): ablation cooling, 286, 

581; Advanced Subsonic Technology (AST) 

program, 796; combustor concepts, 794–

95, 798–99; Cooperative Test Organization/VECTOR, 705–6; E Cubed technology 

use, 742, 747, 792–95; high-bypass fanjet 

engines, 741–43; Numerical Propulsion 

System Simulation (NPSS) program, 760; 

open-rotor engines, 803; Quiet Engine 

program, 737; thermal barrier coatings, 

612–13; Ultra Efficient Engine Technology 

(UEET) program, 762, 797; wind energy 

program, 829, 835, 836, 840, 841, 848

General Electric CF6-50C engine, 745

General Electric CF6 engine, 744

General Electric compressor, 793, 795

General Electric E Core engine, 803

General Electric F404 engine, 802

General Electric GE90-115B jet engine, 795

General Electric GE90 jet engine, 742, 743, 

793

General Electric GE F404 engine, 706n152

General Electric GEnx engine, 769, 795

General Electric GP-7200 engine, 795

General Electric J47 jet engine, 736

General Electric J79 turbojet, 104

General Electric J85 engine, 686

General Electric Mark 3 nose cone, 288

General Electric mullite RSI, 313, 601, 604

General Electric RVX-1 nose cone, 288, 582

General Electric RVX-2 nose cone, 288, 294

General Electric TF30 turbofan engines, 754

General Electric Twin Annular Pre-mixing 

Swirler (TAPS) combustor, 798–99, 817
Index 

911

General Electric Ultra-High Bypass 

 Turbofan engine, 750–51

General Electric Unducted Fan (UDF) 

engine, 749, 750, 801–2

General Motors, 481, 538

General Purpose Airborne Simulator, 

402n62

General Purpose Structural Analysis program, 474–75. See also NASA Structural 

Analysis System (NASTRAN)

General Rotorcraft Aeromechanical Stability Program (GRASP), 492

Geophysical Finite Element Simulation 

(GeoFEST), 556

George, Albert R., 211, 217, 222, 223

George Washington University, 542

Georgia Institute of Technology, 169, 719

Germany: A-2 rocket, 635–36; A-4b (winged 

V-2 derivative), 280, 281; CCV F-104G, 

670, 689–90, 692; Cooperative Test Organization/VECTOR, 706; Dutch-German 

Wind Tunnel (Duits-Nederlandse Wind 

Tunnel), 164; fly-by-wire system research, 

648–49; forward-swept wing aircraft, 21, 

22; high-speed flight program, 280–81; 

International Test Organization (ITO), 

700; oblique wing concept, 522; pitch-up 

problems, 27, 27n51; rotary wing noise 

research, 164; side stick flight controller, 

electrical, 638n15; swept and delta wing 

aircraft, 7, 13, 13n20, 14–15, 15n27, 17, 

72–73, 93; synthetic fuel process, 817; 

technical intelligence interrogations, 

14–15, 17; turbojet engines, 92; V-2 ballistic missiles, 183–84, 280, 282, 571, 635–

36; variable-geometry wing research, 63; 

wind energy, 828; X-31 aircraft, 699–700, 

706, 707

Gessow, Alfred, 144

Giesing, J.P., 483

Gifu Air Base, 692

Gigaflops, 341, 438–39

G II aircraft, 404–5

Gimbaled nose sensor, 298

G-induced loss-of-consciousness (g-loc), 33

Giner, 864

Glacier National Park, 201

Glazer, Jack, 852n33

Glenn, John H., Jr., 103–4

Glennan, T. Keith, 192

Glenn Research Center/Lewis Research 

 Center/Lewis Flight Propulsion Laboratory/Aircraft Engine Research Laboratory: Advanced Subsonic Combustion Rig 

(ASCR), 799; Advanced Tiltrotor Transport Technology (ATTT) program, 167–

68; Advanced Turboprop (ATP) project, 

800; aeronautics expertise, 787; Aircraft 

Engine Research Laboratory, 137, 786–

87; Air Force F-111 program, 61–62; AIRLOADS program, 556; ANSYMP program, 

545; Army research presence at, 139, 147; 

BWB research, 71; COmposite Blade 

STRuctural ANalyzer (COBSTRAN), 546–

47; Computational Structural Mechanics (CSM) workshop, 497; computers at, 

438; Coupled Eulerian Lagrangian Finite 

Element (CELFE), 514; Digital Electronic 

Engine Control (DEEC) program, 709–

10, 754; engine design, computer programs for, 746–47, 748; engine design 

and aircraft speed, 105–6; engine fan 

case improvements, 767; finite element 

analysis activities, 496–500, 532; Glenn 

Research Center, 137; Greenlab Research 

Facility, 817; Helicopter Transmission 

Technology program, 158; higher-order 

theory for functionally graded materials 
NASA’s Contributions to Aeronautics

912

(HOTGFM), 512; High Stability Engine 

Control (HISTEC) system, 717–18; hydrogen fuel research, 809–14; impact analysis 

research, 497; Integrated Composite Analyzer (ICAN), 556; jet engine innovations, 

105–6; Lewis Research Center, 137; Low 

Emissions Alternative Power (LEAP) program, 872; Micromechanics Analysis Code 

with Generalized Method Cells (MAC/

GMC), 512; NASTRAN, development of, 

474, 497; NASTRAN, Users’ Colloquia, 

536, 539; noise research, 738; Numerical Evaluation of Stochastic Structures 

under Stress (NESSUS), 498–99, 514; 

Numerical Propulsion System Simulation 

(NPSS) program, 759–62; propulsion system research, 786–87; Quiet Engine program, 737; rotary wing aircraft research, 

156, 158, 162; rotary wing noise research, 

163; Spin Rig, 555; SST program, 66, 191; 

Structural Performance and Redesign 

(SPAR), 495, 498, 512; Structural Tailoring of Engine Blades (STAEBL), 545–

46; supersonic combustion testing, 331; 

thermal barrier coatings research, 613; 

TRansfer ANalysis Code to Interface Thermal Structural (TRANCITS), 546; TRASYS Radiative Heat Transfer program, 

550; Ultra Efficient Engine Technology 

(UEET) program, 762, 797; Ultra-High 

Bypass Turbofan engine, 750–51; wind 

energy program, 824–25, 827, 829, 831, 

836–39, 841, 844–49; Wind Power Office, 

829–30; women mathematicians, 465; 

YF-12 Blackbird project, 517

Gliding landings, 402–3

Global Disaster Information Network, 861

Global Positioning System (GPS), 242, 706, 

856

Global warming research, 789–90, 790n22, 

818

Glomb (guided glide bomb), 11–12

Gloster Javelin, 45

Gluhareff, Michael E., 8–11, 13n20

Gluhareff-Griswold “Dart” concept, 8–11, 

9n12

Gnoffo, Peter, 451, 451n26

Goddard Space Flight Center (GSFC): 

ACTOMP program, 556; ACTON program, 

556; Dynamic Simulation of Controls & 

Structure (DISCOS), 548–49; finite element analysis activities, 500–502; Integrated Modeling of Optical Systems 

(IMOS), 548; NASTPLT program, 556; 

NASTRAN, development of, 152, 473–74, 

476, 501–2; NASTRAN, maintenance of, 

478; NASTRAN, use of, 477; NASTRAN, 

Users’ Colloquia, 480, 536; NPLOT program, 547; Structural-Thermal-Optical 

(STOP) program, 501–2; Ultra Efficient 

Engine Technology (UEET) program, 762, 

797; VIEW program, 502

Goldburg, Arnold “Bud,” 445

Goldendale, Washington, 837–38

Goldstein, Howard, 315

Goodrich, 481, 538

Goodyear: Goodyear Electronic Differential Analyzer (GEDA), 393, 394; Micromechanics Analysis Code and tire design, 

500; X-20 Dyna-Soar landing skids, 592

Goodyear blimp Mayflower, 204

Gossamer Albatross, 856

Gossamer Penguin, 856

Göthert, Bernard, 29n56

Göttingen aerodynamics, 17

Goward, C.W., 611

GP-7200 engine, 795

Graham, David H., 245–46
Index 

913

Graphite, 576–77, 587–88, 594

Graphite Car, 578

Graphite composite, 804

Graphite-epoxy composite, 576

Gray, George, 18

Great Britain: aeronautical community, ties 

to Langley Research Center, 60–61; Aircraft and Armament Experimental Establishment (A&AEE), 638n16; all-moving 

tail research, 35n65; Concorde Supersonic 

Transport (SST). See Concorde Supersonic 

Transport (SST)delta wing development, 

45, 53; fly-by-wire system research, 649; 

Jaguar Active Control Technology (ACT), 

670, 690–91, 692; Royal Air Force (RAF), 

186, 188, 643; Supersonic Commercial Air 

Transport (SCAT) program, 190; swept 

wing aircraft, 27, 28; turbojet engines, 92; 

V-2 missile shock waves, 183–84; variablegeometry wing research, 62; wind energy, 

828, 848

Great Lakes Industrial Technology Center 

(GLITeC), 500

Green, J.R., 641

Greenhouse gases, 790

Greenlab Research Facility, 817

Grid points: computer technology development and, 438–39; direct numerical 

simulation (DNS), 342, 450–51; Euler 

equations, 438; flow-field calculations, 

433, 434, 435; numerical accuracy, 446–

47; three-dimensional flow-field calculations, 447–49

Griffith, John, 30

Grindle, Thomas, 253

Griswold, Roger W., 9, 9n13

Ground collision avoidance system, 675

Ground Demonstrator Engine (GDE), 345

Ground resonance, 144–45

Ground vibration tests (GVT), 376, 526

Grover E. Bell Award (AHS), 169–70

Groves, Al, 701

Grumman: FASTOP program, 527; NASTRAN Users’ Colloquia, 480, 538; Shaped 

Sonic Boom Demonstration (SSBD), 

245–47

Grumman F6F-5 Hellcat, 401, 637

Grumman F9F-2 Panther, 638–39, 642, 

658n59

Grumman F9F-5 Panther, 33

Grumman F9F-6 Cougar, 33

Grumman F9F-7 aircraft, 33n60

Grumman F9F-9 Tiger, 104, 105, 106

Grumman F11F-1 Tiger, 32, 104, 105

Grumman F-14 Tomcat, 44, 62, 413; F-14A 

Tomcat, 63

Grumman G II aircraft, 404–5

Grumman Gulfstream, 325

Grumman TBM Avenger, 695n125

Grumman X-29 aircraft: aeroelastic effects, 

496, 528–29, 694; aeroelastic tailoring, 

694, 695; angle of attack capabilities, 

694, 695, 696–98; benefits of, 697–98; contract for, 692–93; control laws, 696; control surfaces, 694–95; difference between 

two X-29s, 693–94; exhibit showing, 698–

99; finite element analysis, 460, 515, 526–

29, 530; flight-testing, 695–98; fly-by-wire 

system, 383, 692–98, 753; forward-swept 

wing design, 22, 527, 693–94; NASTRAN 

model, 460; redundant systems, 695; sensor selection thresholds, 698; strain gages, 

528; supercritical wing, 693, 694, 695, 

695n126; vortex flow control, 698; X-29A 

aircraft, 460

Grumman XF10F-1 Jaguar, 59, 60

Guggenheim Aeronautical Laboratory, California Institute of Technology (GALCIT), 
NASA’s Contributions to Aeronautics

914

12, 17n30, 502

Guided glide bomb (glomb), 11–12

Guidonia Laboratory, 6, 332–33

Gulfstream Aerospace Corporation: Quiet 

Spike device, 251–53, 256, 261; Shaped 

Sonic Boom Demonstration (SSBD), 246, 

249; sonic boom mitigation techniques, 

251–53; sonic boom simulator, 257; supersonic business jet (SSBJ) concept, 249

Gulfstream II business jet, 124

Gulfstream III, 121

Gulfstream II turboprop modification, 801

Gulfstream IV, 121

Gulfstream V, 121

Gulfstream X-54A aircraft, 255

Gurevich, M.I., 19n32

Guruswamy, Guru, 494

Gyroplane, 141

Gyros, 298, 364–65, 366, 367

H

H-13 helicopter, 150

Haering, Edward A., 236, 242, 248, 255

Hafnia, 574, 587, 588, 614

Haise, Fred, 370–71, 381

Halaby, Najeeb E., 192

Hall, Charles F., 51–53

Hall, David B., 476

Hall, Eldon, 810, 811, 814

Hallion, Richard P., 115, 595

Hamburger Flugzeugbau HFB-320 Hansa 

Jet, 22

Hamilton Standard, 742, 800–801, 839

Hamilton Standard SR-7A prop fan, 

800–801

Hamilton Sundstrand direct current power 

generation system, 677–78

Hamlin, Fred, 106n48

Hampton, Virginia, 90, 91, 126, 137, 

 146, 191
Index 

915

Henderson, Herbert, 217

Hennrich, Carl, 476

Henry, John, 335

Henschel Hs 129 aircraft, 636–37, 637n11

Henschel Hs 293 aircraft, 638n15

Herbst, Wolfgang, 703n144

Herbst Maneuver, 703, 703n144

Hersman, Walt, 647

Hicks, John, 855

High-altitude aircraft, 850–52. See also Environmental Research Aircraft and Sensor 

Technology (ERAST) program

High-Altitude Long-Endurance (HALE) 

vehicles, 496, 849, 859, 868, 871

High-Altitude Solar (HALSOL) aircraft, 850

High aspect ratio aircraft, 6–7, 13, 13n20, 

390

High aspect ratio wings, 120

High-bypass fanjet engines, 741–43

Higher-order theory for functionally graded 

materials (HOTGFM), 512

Highly Integrated Digital Electronic Control (HIDEC), 710–11, 712

Highly Loaded Turbomachinery project, 763

Highly Maneuverable Aircraft Technology (HiMAT), 22, 325, 515, 520–22, 530, 

686–89

High Performance Computing and Communications program, 759

High-Speed Aircraft Industrial Project 

(HISAC), 250

High-Speed Civil Transport (HSCT) aircraft, 

231–33, 236, 241, 242, 519, 795–96

High-Speed Flight Research Station/HighSpeed Flight Station. See also Dryden 

Flight Research Center (DFRC)/Flight 

Research Center: aerodynamic refinement 

and efficiency improvements, 65; conical camber research, 53; Convair XF-92A 

aircraft, 49; establishment of, 184; flight 

control systems, research activities on, 

363; organization of, 494–95; pitch-up 

research, 25; roll coupling research, 393; 

sonic boom research, 185

High-Speed Research (HSR) program, 232–

43, 519, 795–96, 797

High-Speed Tunnel (HST), 94–97, 127–28

High Stability Engine Control (HISTEC) 

system, 258, 717–18

High Temperature Loads Calibration Laboratory (Dryden Flight Research Center), 

553

High-temperature structures and materials, 569–73

High Temperature Tunnel (HTT), 342, 343, 

573

Hill, G.T.R., 5

Hillaker, Harry J., 69, 69n129, 668

Hiller Aircraft, 136

HiMAT (Highly Maneuverable Aircraft Technology), 22, 325, 686–89

HL-10 aircraft, 56, 360, 368–69, 378–79

Holloman Air Force Base, 200

Honeywell: DISCOS, 549; Numerical 

Propulsion System Simulation (NPSS) 

program, 760; Propulsion Controlled Aircraft (PCA) program, 715; Ultra Efficient 

Engine Technology (UEET) program, 762, 

797

Hoover, Herbert “Herb,” 634

Horizontal-axis wind turbine (HAWT) generators, 823, 825

Horten, Reimar, 70

Horten flying wing, 70

Horton, Walter, 70

Hot Section Technology (HOST) program, 

612

Hot structures: ASSET vehicles, 304, 308–9, 
NASA’s Contributions to Aeronautics

916

418–19, 592–96, 607; complexity of, 312–

13, 603; concept and characteristics of, 

305, 418–19; Dyna-Soar vehicles, 305–8, 

418, 582–92; nose cones, 286, 579, 580–

81; Space Shuttle orbiter, 310–13, 601–

3; thermal protection tiles compared to, 

278; titanium, 312

House Variable Intensity Boom Effect on 

Structures (House VIBES), 256

Houston, 217

Howze, Hamilton H., 139

Hrennikoff, Alexander, 471

Hubbard, Harvey H., 193, 210, 211, 229

Huckleberry Finn technique, 107

Huff, Dennis, 803

Hughes Falcon missiles, 102

Hughes Helicopter, 136

Hughes Offshore, 539

Hughes OH-6 helicopter, 163

Humes, Donald, 318

Hunter, Maxwell, 311, 599, 603

Hunton, Lynn, 218

Hütter, Ulrich, 826–27, 846

Hütter-Allgaier wind turbine, 826–27, 831, 

846

Huygens, Christiaan, 429

Hybrid computers, 406–7, 468

Hybrid laminar flow control (HLFC), 235, 

807–8

Hydraulically actuated flight control surfaces, 384, 590, 631, 634–35, 643, 644

Hydrocarbons, liquid, 817

Hydrogen, 617, 618

Hydrogen, liquid, 809–14, 815–16

Hydrogen-air fuel cell system, 864, 869, 871

Hydrogen-air supersonic combustion

 testing, 332

Hydrogen bomb, 439

Hydrogen fuel cell aircraft, 815

Hydrogen-oxygen regenerative fuel cell system (RFCS), 863–64, 871

Hydrogen peroxide reaction control, 298

Hyperbolic behavior, 443

Hyperbolic partial differential equations, 

435n11, 436, 443

Hyperelastic structures, 503–4

HyperSizer program, 512

Hypersonic aircraft and flight. See also

McDonnell Aerothermodynamic/elastic 

Structural Systems Environmental Tests 

(ASSET); North American X-15 aircraft; 

Space Shuttle orbiter: aircraft for, 279; 

challenges of, 361–62; delta wing development, 53–57; development of technology 

for, 278–79; flight control systems, 298; 

flutter, 306, 309, 388–89; heating, structural, 414–15; military use of, 279; record 

flights, 278, 293, 299, 642; reentry challenges, 285–89, 291–92; research activities, 231; scramjet engines, 279, 330–36, 

340, 342–46, 343n126; stability issues, 292; 

support for, 230–31; wing configuration, 

300–301

Hypersonic Research Engine (HRE), 335, 

337

Hypersonic speed, 276

Hypersonic Weapons Research and Development Supporting System (HYWARDS), 

300–301. See also Boeing X-20 Dyna-Soar

Hypersonic wind tunnels, 280–83, 573

Hyper-X vehicle, 343. See also X-43 HyperX vehicle

HYPULSE Shock Tunnel, 343

HyTech research program, 345

I

IBM. See International Business Machines 

(IBM)
Index 

917

Icing environment research, 162

Illiac IV computer, 325, 438

Illinois Center Plaza Hotel, 481

Ilyushin Badger, 102

Imaging sensors, 861

Impact analysis research, 497

Impact Dynamics Research Facility, 152

Incompressible flows, 12, 15

Inconel 625 nickel alloy, 345

Inconel X, 292, 293, 295, 416, 571, 592

Independence, Project, 823–24, 831

Indian Springs, 195

Indirect analog computers, 468

Induced drag, 117, 118

Industrial Application Centers, 488

Industry/University Cooperative Research 

Center (I/UCRC), 549

Inertial coupling (roll coupling), 37–45, 

40n74, 44nn86–87, 393, 394–95, 407–9

Inertial Management System (IMS), 655

Inertia table, 396

In-Home Noise Generation/Response System (IHONORS), 240

Iniki, Hurricane, 859

Inlet Spillage Shock Measurement (ISSM), 

246

Innis, Robert, 649

Institute for Computer Applications in Science and Engineering (ICASE), 446

Institute for Scientific Research, 719

Instrument flight rules (IFR), 151

Instrument meteorological conditions 

(IMC), 159

Integral equations, 433

Integrated Component Technology Demonstrations project, 764

Integrated Composite Analyzer (ICAN), 500, 

547, 556; ICAN/PART, 500

Integrated Design and Evaluation of 

 Advanced Spacecraft (IDEAS), 511–12, 
NASA’s Contributions to Aeronautics

918

version development, 484; 20-petaflop 

computer, 439

International Civil Aviation Organization 

(ICAO): emissions standards, 762, 791, 

796, 797, 798, 799; noise level standards, 

250; sonic booms, reaction to, 261

International Harvester

 Aerothermodynamic Structural Vehicles 

(ASV), 308–9, 595

International Helicopter Study Team 

(IHST), 172

International Nickel, 292

International Panel on Climate Change, 

United Nations, 790, 790n22

International Space Station (ISS), 243, 506

International Test Organization (ITO), 

700–701

Inventions and Contributions Board, 116

Inviscid flows, 438, 444–45, 448

Iron Bird simulator, 655, 656–57, 659, 660

Iron Cross simulator, 298

Iron-nickel–based alloys, 573, 574, 575

Ishmael, Stephen D., 662, 679, 695

Israel, 481

Ivan C. Kincheloe Award, 649

Izu, Douglas, 310, 598

J

J29 fighter, 23n39

J32 Lancen (Lance), 23n39

J35 Draken (Dragon), 45

J47 jet engine, 736

J57 turbojet, 100, 101, 102

J58 engines, 708, 709–10

J75 turbojet, 103

J79 turbojet, 104

J85 engine, 686

Jackass Flats testing, 206–7, 213

Jacobs, Eastman, 17n30, 109

Jacobs, Joseph, 826

Jacobs, Marcellus, 826, 827

Jacobs, Peter F., 119, 120

Jaguar Active Control Technology (ACT), 

670, 690–91, 692

Jameson, Antony, 341, 445–46, 494

Japan: CCV T-2, 670, 691–92; NASTRAN use 

by, 481, 484; Technical Research Development Institute (TRDI), 692

Japan Aerospace Exploration Agency 

(JAXA), 255

Jarvis, Calvin R., 654, 655

JAS-39 Gripen fighter, 706

JB-47E aircraft, 645–46

Jesperson, Dennis, 452

Jet engine aircraft: handling qualities, 634; 

hydraulically actuated flight control surfaces, 634–35; innovations in and aircraft 

speed, 105–6

Jet Propulsion Laboratory (JPL): Antenna 

Design and Analysis, 549–50; finite element analysis activities, 502–5; Geophysical Finite Element Simulation (GeoFEST), 

556; high-altitude aircraft workshop, 851; 

history of, 502–3; Integrated Modeling of 

Optical Systems (IMOS), 548; NASTRAN 

development, 474, 503; SAMIS program, 

506; System Identification/Parameter 

Identification (PID), 504–5, 504n113

Jettisonable assistance takeoff (JATO), 29

JF-100C Super Sabre, 402n62

John Deere, 539

Johns Hopkins University Applied Physics 

Laboratory, 332, 333

Johns Manville, 305, 314, 586, 600, 601

Johnson, Clarence “Kelly,” 34n61

Johnson, James, 474

Johnson, Kelly, 812–13

Johnson, Lyndon B., 66
Index 

919

Johnson, Richard L. “Dick,” 89

Johnson, Wayne, 159

Johnson Space Center/Manned Spacecraft 

Center: carbon-carbon composite study, 

577; finite element analysis activities, 

505–7; NASTRAN development, 474, 506; 

OVERFLOW code, 452, 453; Shuttle concept development, 56–57; structural test 

facility, 553; thermal-protection tiles, 318, 

319, 419; TRASYS Radiative Heat Transfer program, 550

Johnston, A.M. “Tex,” 22–23

Johnsville, Pennsylvania, 297, 398

Joint Acoustic Propagation Experiment 

(JAPE), 238–39

Joint Planning and Development Office 

(JPDO), 171, 171n53

Joint Sponsored Research Agreement 

(JSRA), 849, 852–55, 852n33, 872

Joint Strike Fighter (JSF) program, 675

Joint Strike Fighter/Integrated Subsystems 

Technology (J/IST) program, 676–77

Joint Task Force 2, 195

Jones, L.B., 223

Jones, Robert T.: delta wing development, 

12–13, 13n20, 46, 48, 51, 53, 67; oblique 

wing concept, 63, 523, 523n169; Perl’s 

association with, 17n30; swept and delta 

wing aircraft, 8, 11–17, 93, 108

Jordan, Gareth, 195–96

Joyce, David, 751

JP-7 fuel, 345–46

JT8D engine, 744

JT9D-7A engine, 745

JT9D engine, 611, 615, 744, 745

Junkers, Hugo, 70

Junkers Ju 88 bomber, 636

Junkers Ju 287 bomber, 21

Jupiter, 277, 288, 573. See also Galileo probe

Jupiter-C missile, 288

Jupiter intermediate-range ballistic missile 

(IRBM), 286, 287

K

Kalman, T.P., 483

Kaman Aerospace Corporation/Kaman Aircraft Corporation: helicopter design and 

production, 136; RITA membership, 169; 

wind energy program, 829, 835, 845

Kaman Circulation Control Rotor (CCR), 

156–57

Kaman Controllable Twist Rotor (CTR), 153

Kane, Edward J., 224

Kansas City, Missouri, 537

Kantrowitz, Arthur, 14, 14n23, 17n30, 286

Karlskronavarvet (KKRV), 839

Kartveli, Alexander “Sasha,” 333, 334, 569

Kauai, 858, 859

Kaufman, Stanley, 475

KC-135 aerial tanker: aerial refueling, 253; 

design of, 20; sonic boom research, 204; 

winglet use on, 120, 122–23, 806

Kellett KD-1 autogiro, 141

Kellett YG-1A autogiros, 141

Kellett YG-1B autogiros, 141

Kempel, Robert, 369

Kennedy, John F., 66, 109, 191, 193, 304, 644

Kennedy Space Center (KSC): finite element analysis activities, 491n64; mission 

of, 491n64; NASTRAN Users’ Colloquia, 

536; Space Shuttle Columbia, 321

Kenyon, George, 56

Kerosene, synthetic aviation, 815–16

Keyworth, George, 337

Kick angle, 527

Kiel probe, 701, 704, 705, 705n149

Kincheloe, Iven, 297, 298, 394

King Air 200 twin turboprop aircraft, 481
NASA’s Contributions to Aeronautics

920

Kleber-Colombes, 538

Kleinknecht, Kenneth, 323

Knight, William J. “Pete,” 64, 294, 418

Knöpfel, Rüdiger “Rudy,” 706

Knox, Fred, 701

Kock, Berwin, 517

Korean war: F-86 Sabres in, 4, 19, 26; MiG 

15s in, 26–27; rotary wing aircraft use, 

138–39

Korycinski, Peter, 55, 301, 307

Kotcher, Ezra, 13–14, 14n22, 94

Krier, Gary E., 647, 657, 658, 660, 661

Küchemann, Dietrich, 13n20, 93, 107

Küssner, H.G., 15

Kvaternik, Raymond G., 508

Kwajalein Atoll, 607

L

L-23 sailplane, 248, 257

L-39 aircraft, 22–23, 23n37, 23n39

L-1011 aircraft, 120, 122

Lambert Field, 646

Laminar Flow Control (LFC) project, 235, 

495, 744, 784, 786, 806–8

Lancaster, 196

Lanchester, Frederick W., 118

Lang, Karl, 701, 704, 705

Langley Aerothermodynamic Upwind 

Relaxation Algorithm (LAURA), 451–52, 

453

Langley Research Center/Langley Aeronautical Laboratory/Langley Memorial 

 Aeronautical Laboratory: Acoustically 

Treated Nacelle program, 740; Acoustics 

Division, 195, 234; Advanced Tiltrotor 

Transport Technology (ATTT) program, 

167–68; Advanced Turboprop (ATP) project, 800; aerodynamic refinement and 

efficiency improvements, 90, 128–29; 

Aeroelastic Rotor Experimental System 

(ARES), 164; Aeronautics Research Mission Center, transformation to, 166; Aeronautics System Division, 220; Air Force 

F-111 program, 61–62; all-moving tail 

research, 35; Arc-Heated Scramjet Test 

Facility (AHSTF), 343; Army research 

presence at, 139, 146, 147; AST Project 

Office, 220; Avro CF-105 Arrow development, 641–42, 642n25; Bi-Level Integrated System Synthesis (BLISS), 511; 

British aeronautical community, ties 

to, 60–61; BWB research, 71; columbium heat shields, 602–3; CombustionHeated Scramjet Test Facility (CHSTF), 

343; computational fluid dynamics and 

transonic flow and airfoils, 445–46; computational fluid dynamics CFL3D code, 

452–53; computational fluid dynamics LAURA code, 451–52, 453; computational fluid dynamics Shuttle study, 426; 

Computational Structural Mechanics 

(CSM) workshop, 497; computers, analog, 468; computers, IBM, 438; CONMIN 

program, 511; Continuous Flow Hypersonic Tunnel (CFHT), 283; Convair YF-102 

Delta Dagger testing, 100–101; delta wing 

development, 53; delta wing research, 

47–48, 48n94; Design Analysis Methods 

for Vibration (DAMVIBS), 508–11; Differential Maneuvering Simulator (DMS), 

666–67; Direct Connect Supersonic Combustion Test Facility (DCSCTF), 343; DM-1 

glider research full-scale tunnel studies, 

47; finite element analysis activities, 507–

12, 532; Finite Element Composite Analysis Program (FECAP), 556; flat-bottom 

delta research, 55–56; FLEXSTAB program, 544–45; Flight Research Branch, 
Index 

921

154; flutter suppression research, 388, 389; 

fly-by-wire research activities, 638–39; forward-swept wing research, 21–22; Free 

Flight Tunnel studies, 63; fuel research, 

815–16; fuel-saving aircraft structures, 

787; Gas Dynamics Branch, 332–33; glomb 

(guided glide bomb), 11; Ground Demonstrator Engine (GDE), 345; high-altitude aircraft workshop, 850; High-Speed 

Tunnel (HST), 94–97, 127–28; high-speed 

tunnel studies, 36; high-temperature tunnels, 342, 343, 573; hingeless rotor concept 

full-scale tunnel studies, 150; HL-10 lifting body research, 56; hypersonic flight 

research, 299; Hypersonic Research 

Engine (HRE), 335; Hypersonic Tunnel, 

282–83; Hypersonic Weapons Research 

and Development Supporting System 

(HYWARDS), 301; HYPULSE Shock 

Tunnel, 343; Impact Dynamics Research 

Facility, 152; inertial coupling research, 

42–43; Institute for Computer Applications in Science and Engineering (ICASE), 

446; instrument flight rules (IFR) investigation, 151; Integrated Design and Evaluation of Advanced Spacecraft (IDEAS), 

511–12; laminar flow control research, 

807; Low-Frequency Noise Facility, 

215; Lunar Landing Research Facility, 

152; Mach 6 tunnels, 343; Mach 10 tunnel, 343; multidisciplinary optimization 

(MDO) studies, 511; NASTRAN, development of, 152, 473–74; NASTRAN, management of, 477, 508; NASTRAN, Users’ 

Colloquia, 480, 508, 536, 540, 542; noise 

research, 163, 164–65; Programming 

Structural Synthesis (PROSS), 511–12; 

rotary wing aircraft full-scale wind tunnel 

studies, 140, 143, 144; rotary wing aircraft 

research, 137, 139–45, 147–52, 154–55, 

156, 161, 162–63, 164–65; Rotor Systems 

Research Aircraft (RSRA) program, 154–

55; SCAT-15F, 67; scramjet engines, 342–

44; scramjet research, 335–36, 343n126; 

SCR/SCAR/AST conference, 225–26, 228–

29; Shaped Sonic Boom Demonstration 

(SSBD), 246; slotted transonic tunnel, 93, 

95–96; sonic boom conferences, symposiums, and workshops, 209–10, 231, 233–

34; sonic boom research, 192–200, 203, 

208–9, 214–15, 217, 227, 239–40, 256–

58; sonic boom research, publication of, 

218–19, 221; sonic boom research simulators, 249, 257–58; sonic boom wind tunnel studies, 235; Space Shuttle concept 

development, 56–57; Space Shuttle orbiter 

heat shield, 312; SST program, 66–67, 191; 

Steep Approach program, 737; Structural 

Analysis of General Shells (STAGS), 514, 

551–52; Structural Performance and Redesign (SPAR), 498; structural test facility, 

553; ST-SIZE program, 512; supercritical wing program, 110–15, 805–6; Supersonic Pressure Wind Tunnel, 208, 208n66, 

214; supersonic tunnel, 14; swept wing 

research, 6, 8, 13–15; tail configuration 

research, 36; thermal-protection tile testing, 318; Transonic Aerodynamic Branch, 

94, 108; Transonic Aircraft Technology 

(TACT) program, 386; Transonic Dynamics Tunnel (TDT), 149–50, 161, 164, 388, 

389; Transonic Pressure Tunnel (TPT), 

109–10, 118, 119, 123, 127–28; Transonic 

Tunnel (High-Speed Tunnel), 95n16; Ultra 

Efficient Engine Technology (UEET) program, 762, 797; Unitary Wind Tunnel, 235, 

585–86; variable-geometry wing research, 

59, 60–61, 61n115, 63; Vehicle Technology 
NASA’s Contributions to Aeronautics

922

Directorate (VTD), 162–63; vortex lift 

research, 651n49; wind energy program, 

829; wind tunnel studies, 665, 702; winglet program, 117–19, 120, 122–23; X-15 

development, 291–93; X-43 Hyper-X vehicle, 343–45, 344n127; X-48B testing, 71; 

YF-12 Blackbird project, 517

Large Eddy Simulation (LES), 341–42

Larson, George, 117

Larson, Howard, 315

Larson, Nils, 259, 260

Lateral-directional stability, 19, 38–39

LAURA (Langley Aerothermodynamic 

Upwind Relaxation Algorithm), 451–52, 

453

Laureate Award in Aeronautics and Propulsion, 253

Laurel Award, 856

Lavochkin fighters, 36

Leading edge extensions, 414

Lean Pre-mixed, Pre-vaporized (LPP) combustor concepts, 798–99

Lear, Bill, 115

Lear 28/29 aircraft, 481

Lear 35/36 aircraft, 480

Lear 55 business jet, 481

Learjet, NASTRAN use by, 480–81

Learjet Model 28 aircraft, 121

Learjet Model 55 aircraft, 121

Learjet variable stability aircraft, 402n62

Leduc, René, 50

Lemay, Curtis E., 812

Lewis, David, 11

Lewis, George W., 14, 142

Lewis Flight Propulsion Laboratory. See

Glenn Research Center/Lewis Research 

Center/Lewis Flight Propulsion 

 Laboratory/Aircraft Engine Research 

Laboratory

Lewis Research Center. See Glenn Research 

Center/Lewis Research Center/Lewis 

Flight Propulsion Laboratory/Aircraft 

Engine Research Laboratory

Lewy, Hans, 435–36

LI-900 thermal protection, 311, 313–17, 

600–601, 604, 605

LI-1500 thermal protection, 310–11, 313, 

599–600, 601, 604

LI-2200 thermal protection, 317

Lichtenstein, Jacob H., 36

Liebeck, Robert H., 70

Lift and Nozzle Change Effects on Tail 

Shocks (LaNCETS), 258–60, 261

Lifting bodies research, 56, 417, 418

Lifting line theory, 12

Lifting surface predictions, 483

Lift-to-drag ratio (L/D), 118, 403, 415, 595

Light emitting diodes (LEDs), 521–22, 

527–28

Lightweight Fighter (LWF) program, 370, 

647, 664

Limit-cycle oscillation, 367–69, 377, 378

Lindbergh, Charles A., 591

Lindbergh Field, 102

Linear Differential Variable Transformers 

(LDVTs), 656

Linear partial differential equations, 431, 

435–36

Ling-Temco-Vought A-7D Corsair II DIGITAC, 382–83

Ling-Temco-Vought shock expansion tube, 

214–15

Link Trainer, 295, 392

Lippisch, Alexander, 5, 7, 8, 13, 13n20, 14, 

45, 46–47, 58

Lippisch DFS DM-1 aircraft, 13

Lippisch DM-1 aircraft, 47–48, 48n94

Lippisch P 13 aircraft, 13, 47
Index 

923

Liquid hydrogen, 809–14, 815–16

Liquid methane, 815–16

Liquid oxygen (LOX) lines, 513

Lithgow, Mike, 67n127

Little Boom, Project, 194–95

Littleman, Project, 199

Littleton Research & Engineering, 538

Load factors, 377

Loads Alleviation and Mode Suppression 

(LAMS) system, 392

Loads prediction, 461–62, 462n1, 465–66

Lockalloy, 620

Lockheed A-12 Blackbird, 210, 570

Lockheed C-5 aircraft, 337, 407, 867

Lockheed C-141 aircraft, 324, 658n59

Lockheed CL-400 aircraft, 812–13

Lockheed CL-1200 Lancer, 34n61

Lockheed coatings: LI-0042 coating, 315; 

LI-0050 coating, 315

Lockheed Corporation: Aerospaceplane proposal, 334; high-altitude aircraft workshop, 851; NASTRAN Users’ Colloquia, 

480; Palo Alto Research Laboratory, 551, 

600; sonic boom research, 237; SST proposal, 210; supersonic business jet (SSBJ) 

concept, 249; titanium hot structures, 312; 

wind energy program, 829, 832, 834; winglet use by, 121–22

Lockheed F-16, 710. See also General 

 Dynamics F-16 Fighting Falcon

Lockheed F-94 Starfire, 102

Lockheed F-104 Starfighter: design of, 104, 

401, 409, 440; development and manufacture of, 290; F-104A Starfighter, 37, 

440; F-104B, 409; F-104G, 688; F-104G 

(German CCV), 670, 689–90, 692; heating, structural, 413; inertial coupling, 

43, 409; landing of, 405; landing simulator, 403–4; nickname for, 4, 104; pitch-up 

problems, 34n61, 413; reaction controls, 

298; sonic boom research, 195–96, 199, 

200, 204, 205, 206–7; speed of, 413; supersonic flight by, 3, 104; use of, 104; wingspan-to-fuselage length ratio, 407; XF-104 

Starfighter, 185

Lockheed F-117 stealth fighter, 70, 384, 

721–22

Lockheed Have Blue aircraft, 721

Lockheed L-1011 aircraft, 120, 122

Lockheed Loads Alleviation and Mode Suppression (LAMS) system, 392

Lockheed Martin: QSP proposal, 245; 

Shaped Sonic Boom Demonstration 

(SSBD), 246; supersonic business jet 

(SSBJ) concept, 250; Ultra Efficient 

Engine Technology (UEET) program, 

762, 797

Lockheed Martin F-22 Raptor, 126–27, 384, 

402, 413, 712, 713, 722

Lockheed Martin F-35 Lightning II Joint 

Strike Fighter, 384, 413, 678, 713, 722

Lockheed Martin X-33 Research Vehicle, 

384

Lockheed Missiles and Space, 310

Lockheed NT-33A Shooting Star, 402

Lockheed P-80 Shooting Star, 634

Lockheed SR-71 Blackbird: accident with, 

392; aerodynamics of, 70; aeroelastic 

effects, 391–92; digital engine control system, 754; engine inlet unstarts, 708–9, 754; 

heating, structural, 413; JP-7 fuel, 346; 

simulator, 407; sonic boom research, 202–

3, 204, 205–6, 236–38, 239, 242–43, 245, 

261; spacesuit for, 297; speed of, 413; thermal barrier coatings, 611; thermal loads 

project, 516; titanium construction, 312, 

570–71, 602

Lockheed Star Clipper, 311, 599, 600
NASA’s Contributions to Aeronautics

924

Lockheed TF-104G aircraft, 688

Lockheed thermal protection, 310–11, 419, 

596–99, 602; borosilicate glass coatings, 

315–16, 599–600; LI-900 thermal protection, 311, 313–17, 600–601, 604, 605; 

LI-1500 thermal protection, 310–11, 313, 

599–600, 601, 604; LI-2200 thermal protection, 317; Space Shuttle tiles, 310–23, 

419

Lockheed U-2 spy plane, 297, 811, 812–13

Lockheed X-7 testbed, 278, 289, 290, 572

Lockheed X-17 missile, 287–88

Lockheed X-27 aircraft, 34n61

Lockheed XF-104 Starfighter, 185

Lockheed XH-51N helicopter, 150

Lockheed X-Wing Stoppable Rotor, 157, 158

Lockheed YF-12 Blackbird: digital engine 

control system, 754; Loads Alleviation 

and Mode Suppression (LAMS) system, 

392; sonic boom research, 202–3, 204, 

210, 227, 227n133; thermal loads project, 515, 516–19, 529, 554; YF-12A, 202–

23; YF-12C, 516, 708–9; YF-12 Experiment 

Symposium, 519

Lockheed YF-22 aircraft, 402

Lockheed YO-3A aircraft, 242–43

Loftin, Laurence K., Jr., 108, 111, 125

Long-Davis, Mary Jo, 254

Longhorn wing and winglet system, 121

Longitudinal stability: all-moving tail and, 

34–35, 35n65, 362–63, 363n4; inertial coupling (roll coupling) and, 38–39; static, 

648, 664, 669, 690, 692; swept and delta 

wing aircraft, 19, 21; T-tail aircraft, 36–37

Long Range Planning Study for Composites (RECAST), 804

Lord, W.T., 94

Los Alamos National Laboratory, 498, 541

Los Angeles, 206

Louisiana Tech University, 542

Love, Eugene S., 56

Low aspect ratio aircraft, 11–12, 13n20, 18, 

43, 54

Low-boom demonstrator, 249, 250–51, 255, 

261

Low boom/no boom technique, 255–57

Low Emissions Alternative Power (LEAP) 

program, 872

Low-Frequency Noise Facility, 215

LTV Aerospace, 576, 618

LTV YA-7D Digital Flight Control for Tactical Aircraft (DIGITAC), 684–86

Ludington-Griswold Company, 9

Ludox, 320–21

Lunar Landing Research Facility, 152

Lunar Landing Research Vehicles (LLRVs), 

653–54

Lunar Landing Training Vehicles (LLTVs), 

653–54

Lunar Module, vii, 654

Lux, David, 236

Lynntech, Inc., 864

Lyons, Valerie, 814

M

M2-F2 lifting body, 407

MacCormack, Robert, 446–47, 448

MacCormack’s method, 446–47

Mach, Ernst, 97, 182

Mach 2+ supersonic transport, 11

Mach angle of wing, 15–16, 18

Mach number: definition, 182; of wing, 5, 

12, 14, 18, 52, 53

Mack, Robert, 223–24, 225, 229, 235, 249

Mackall, Dale, 682, 683

MacKay, John, 331–32

Maclise, Dougal, 861

MacNeal, Richard H., 467, 475
Index 

925

MacNeal Schwendler Corporation (MSC): 

cofounder of, 467; finite-element heat 

transfer program, 502; MSC/NASTRAM, 

478, 484, 499, 546, 547; NASTRAN competitors, 486; NASTRAN development, 

475; NASTRAN maintenance, 479

Maglieri, Dominic J.: AST project, 220; 

Shaped Sonic Boom Demonstration 

(SSBD), 246; sonic boom conferences, 

210, 211, 212; sonic boom research, 193, 

229, 236; supersonic aircraft proposals, 

245–46; XB-70 records, 238

Magnesia, 587

Magnesia-stabilized zirconia, 611

Magnesium, 293, 571

Mahesh, Krishnan, 341

Makani Uwila Power Corporation, 842

Mallick, Donald L., 638

Management and Budget, Office of (OMB), 

601–2

Manching, 690

Maneuver load control, 652

Manke, John A., 381, 661

Manned Orbiting Laboratory (MOL), 303

Manned Spacecraft Center (MSC). See

Johnson Space Center/Manned Spacecraft Center

Marc Analysis & Research, 484, 546

Marching solutions, 435–36

Marconi Company, 671

MARC structural analysis code, 546

Marine Corps, U.S. (USMC): AFTI/F-16 aircraft, 679; V-22 Osprey, 166; Vietnam war 

aircraft losses, 644n32; X-31 aircraft, 707

Mariotte, Edme, 428–29

Marquardt ramjets, 289, 290

Marquardt scramjet flight-test program, 

335, 336

Marshall Islands, 439

Marshall Space Flight Center: ASSIE-I program, 552; Coupled Eulerian Lagrangian 

Finite Element (CELFE), 514; Cylindrical 

Optimization of Rings, Skin and Stringers (CORSS), 513; finite element analysis 

activities, 498, 512–14; impact analysis 

research, 497; NASTRAN, development 

of, 474, 514; NASTRAN, Users’ Colloquia, 

536; Program for Analysis of Nonlinear 

Equilibrium and Stability (PANES), 552; 

Shuttle concept development, 57; Structural Analysis of General Shells (STAGS), 

514, 551–52; structural test facility, 553; 

VLOADS launch loads and dynamics program, 513–14; wind energy program, 829

Martin, H.C., 471

Martin, Roy, 247

Martin B-57B bomber, 810, 811

Martin Baltimore, 475

Martin MA-25S ablative coating, 417–18

Martin Marietta TRASYS Radiative Heat 

Transfer program, 550

Martins, Joaquim, 494

Martin SV-5D (X-23) PRIME (Precision 

Recovery Including Maneuvering Entry), 

27n51, 278, 288, 304, 417, 418, 606–8

Martin X-24A (SV-5P) lifting body, 27n51, 

407, 607

Martin X-24B lifting body, 56, 404, 407

Martin X-33 Research Vehicle, 384

Martin X-38 aircraft, 27n51

Mascitti, Vincent R., 220

Mason, James B., 476

Massachusetts Institute of Technology 

(MIT), 471, 654

Master Builders, Inc., 500

Materials and Structures for High Performance project, 763

Matranga, Gene, 517
NASA’s Contributions to Aeronautics

926

Matrix method, 466, 470–71

Matrix Methods in Structural Mechanics 

conference, 474–75

Mattingly, Kenneth, 658

Mauchly, John, 436

Maxwell, James Clerk, 433–34

Mayo Graduate School of Medicine, 542

McAleese, James D., 474

McComas, Robert L., 474, 536

McConnell, Richard D., 476

McCormick, Caleb W., 475

McDonnell Aerothermodynamic/elastic 

Structural Systems Environmental Tests 

(ASSET): Aerothermodynamic Elastic 

Vehicles (AEV), 308–10, 594, 595; Aerothermodynamic Structural Vehicle (ASV), 

308–10, 594–95; flights in program, 309, 

594–95; hot structures, 304, 308–10, 418–

19, 592–96, 607; International Harvester 

Aerothermodynamic Structural Vehicles 

(ASV), 308–9, 595; nose cap, 594–95; performance of, 309, 596–97

McDonnell Aircraft Corporation, 592

McDonnell-Douglas 500E helicopter, 165

McDonnell-Douglas Aerospace, 494

McDonnell-Douglas BWB-450 aircraft, 71

McDonnell-Douglas C-17 Globemaster III 

transport, 127, 384, 715, 722, 795

McDonnell-Douglas Corporation: Advanced 

Control Technology for Integrated Vehicles Experiment (ACTIVE) project, 717; 

Boeing merger, 127; BWB transport, 

70–71; Collier Trophy, 127; lifting surface 

predictions, 483; NASA contract work, 

120; NASTRAN Users’ Colloquia, 480; 

sonic boom research, 237; STOL transport competition, 114; winglet research, 

123, 125

McDonnell-Douglas DC-8 aircraft, 737, 817

McDonnell-Douglas DC-9 aircraft, 745

McDonnell-Douglas DC-10 aircraft: engine 

component improvements, 745; Sioux City 

airliner crash, 757, 766; winglet research, 

119, 120, 122, 123

McDonnell-Douglas F-4 Phantom, 644n32

McDonnell-Douglas F-15 Eagle: Advanced 

Control Technology for Integrated Vehicles Experiment (ACTIVE) project, 716, 

717; Digital Electronic Engine Control (DEEC) program, 709–10, 754–56; 

Dryden research activities, 496; engines, 

713; F-15A, 44, 711, 713; F-15E Strike 

Eagle, 69; fly-by-wire system, 383–84; 

Highly Integrated Digital Electronic Control (HIDEC), 711; High Stability Engine 

Control (HISTEC) system, 718; Intelligent 

Flight Control Systems (IFCS) project, 

718–21; Propulsion Controlled Aircraft 

(PCA) program, 713, 757; Self-Repairing 

Flight Control System (SRFCS), 713–14; 

simulator, 407; sonic boom research, 238–

39, 246–48, 252–53, 256, 258–60, 261; tail 

configuration, 413; TF-15, 715

McDonnell-Douglas F-15 Eagle (F-15B/NF15B): Advanced Control Technology for 

Integrated Vehicles Experiment

 (ACTIVE) project, 716, 717; exhibit

 showing, 721; High Stability Engine Control (HISTEC) system, 718; Intelligent 

Flight Control Systems (IFCS) project, 

718–21; Lift and Nozzle Change Effects 

on Tail Shocks (LaNCETS), 258–60, 261; 

modifications for experiments, 258–59; 

Quiet Spike device, 252–53, 256; Short 

Takeoff and Landing Maneuver Technology Demonstrator (S/MTD), 715–16; standard features, 259

McDonnell-Douglas F-18 Hornet, 237, 251, 
Index 

927

255–57, 496, 676

McDonnell-Douglas F/A-18 Hornet: Active 

Oscillation Suppression (AOS) system, 

389; autothrottle system from, 706; design 

of, 669; engines for, 706n152; F/A-18A 

Hornet, 44, 383–85; F/A-18B Hornet, 386; 

tail configuration, 413

McDonnell-Douglas Harrier strike fighter, 

60

McDonnell-Douglas Helicopter Company, 

508, 510

McDonnell-Douglas MD-11 aircraft, 123, 

715, 757, 758, 806

McDonnell-Douglas Project 680J Survivable Flight Control System (SFCS), 379–

80, 384, 646–48, 658n59

McDonnell-Douglas YC-15 aircraft, 114, 127

McDonnell-Douglas YF-4E Phantom Survivable Flight Control System program, 

379–80, 384, 646–48, 658n59, 660

McDonnell F-4E Phantom II, 195

McDonnell F4H-1 Phantom II, 32, 33n60, 34

McDonnell F-101 Voodoo: development 

and manufacture of, 290; flight control 

system, 372; pitch-up problems, 33–34, 

413; sonic boom research, 194, 199; sonic 

booms, 185; tail configuration, 33–34, 36, 

413; turbo fan engines, 741

McDonnell XF-88 aircraft, 33

McKee, William F. “Bozo,” 192

McKinney, Linwood, 651n49

McLean, F. Edward, 67, 210, 211, 217, 220, 

222

McLellan, Charles, 292

McMurtry, Thomas C. “Tom,” 89, 112, 113, 

657, 658, 661, 714

McNamara, Robert S., 192, 211, 303–4, 

400, 592

MD-11 aircraft, 123, 715, 757, 758, 806

MDBOOM, 237

Mechanical linkages in flight control systems, 525, 631, 633, 637–38, 639, 655

Medal for Exceptional Scientific Achievement, 116

Medicine Bow, Wyoming, 839, 848

Mercury ballast, 595

Mercury program: ablation cooling, 417; 

centrifuge program, 400; flight control systems, 653; heat shields, 587; McDonnell 

Aircraft spacecraft, 592; reaction controls, 

298; reaction thrusters, 278; Redstone 

missile/booster, 636; reentry vehicles, 54; 

side stick flight controller, 658n59; spacesuits, 297

Messerschmitt-Bölkow-Blohm/

 Messerschmitt: advanced project group, 

14, 48n94; Military Aircraft Division, 

703n144; variable-geometry wing aircraft, 58–59

Messerschmitt Me 163 Komet (Comet), 7, 

13, 27, 27n51

Messerschmitt Me 262 aircraft, 27n51

Messerschmitt P 1101 aircraft, 58–59

Metallurgy: directionally solidified (d.s.) 

castings, 614–15; powder metallurgy, 

616–17; rapid-solidification technique, 

279, 617, 620–21

Metal matrix composites (MMCs), 617–18

Methane, 342, 815–16

Methyl methacrylate (Plexiglas), 310–11, 

598

Meyer, Corwin “Corky,” 23, 59, 104

Michel, Doug, 473–74

Microcraft X-43 vehicle, 279, 341, 343–45, 

344n127

Micromechanics Analysis Code (MAC), 500

Micromechanics Analysis Code with Generalized Method Cells (MAC/GMC), 512
NASA’s Contributions to Aeronautics

928

Middleton, Wilbur, 67

MiG aircraft. See Mikoyan and Gurevich 

entries

Mighty Mouse rockets, 102

Mikoyan, Anushavan “Artem,” 26

Mikoyan, Stepan, 26

Mikoyan and Gurevich MiG-15 aircraft: 

development of, 19n32, 23n39; Korean 

war use of, 4, 19, 26; pitch-up problems, 

26–27; tail configuration, 36

Mikoyan and Gurevich MiG-17 aircraft, 27

Mikoyan and Gurevich MiG-21 aircraft, 

32, 45

Mikoyan and Gurevich MiG-23/27 Flogger, 62

Miles M.52 aircraft, 35n65

Miller, Robert, 611, 612–13

Millikan, Clark, 14

Minneapolis, 206

Minneapolis Honeywell MH-96 Adaptive 

Flight Control System (AFCS), 298–99, 

372–74

Minneapolis Honeywell MH-96 flight control system, 376, 379

Minuteman intercontinental ballistic missile (ICBM), 286, 303

Mirage IIIE aircraft, 691

Mishap Investigation Board (MIB), 870–71

Missile Defense Agency, 850

Mission-adaptive wing (MAW), 385–86

Mistle (Mistletoe) aircraft, 636

Mitsubishi F-2 fighter, 692

Mitsubishi Regional Jet, 752

Mitsubishi T-2 aircraft, 691–92

Mod-0A wind turbine, 831, 833–35

Mod-0 wind turbine testbed, 827, 830, 831–

33, 846

Mod-1A wind turbine, 836

Mod-1 wind turbine, 831, 835–36, 845

Mod-2 wind turbine, 830, 831, 836–39, 840, 

843, 847, 848

Mod-3 wind turbine, 840

Mod-4 wind turbine, 840

Mod-5A wind turbine, 840, 841

Mod-5B wind turbine, 840, 841–44, 847

Mod-5 wind turbine, 831, 840

Mod-6 wind turbine, 840

MODE-ID program, 504

Model-following systems, 366–67

Modern Rotor Aerodynamic Limits survey, 

515, 519–20, 529–30

Modlin, C. Thomas, Jr., 474, 506

Moes, Timothy R., 236, 259

Moffett Field, 137

Molybdenum, 306, 575, 583, 587

Molybdenum alloy, 306, 308, 584

Molzahn, Leslie M., 253

Momentum equation, 430, 431, 433, 438, 

443

Monoplane, 387

Moon: mission to, 787; reentry from, 278, 

288; return to, 765; training to land on, 

653–54

Moore School of Electrical Engineering, 

University of Pennsylvania, 436

Morane brothers, 385

Morane Bullet, 34, 385

Moretti, Gino, 444

Morgan, David, 24

Morgan, Ray, 855, 856, 868–69

Morgenstern, John, 246

Moskalev, Alexandr Sergeevich, 9n12

MOSTAB computer code, 834

Motion, equations of, 393

MSC-040 concept, 602

MSC/NASTRAM, 478, 484, 499, 546, 547

Mufflers, 737, 740

Mullite reusable surface installation, 
Index 

929

 313, 604

Multhopp, Hans, 27n51

Multidisciplinary optimization (MDO) studies, 511

Munk, Max, 11, 12

Murman, Earll, 445

Muroc Army Airfield, 182, 184

Muroc Dry Lake, 23–24

Muroc Flight Test Unit, 494. See also Dryden 

Flight Research Center (DFRC)/Flight 

Research Center

Mussolini, Benito, 332–33

Mutterperl, William, 17n30

Myasishchev Bison, 102

Myers, Dale, 313, 603–4

Myers, Garry, 144

N

N-156 aircraft, 32

N aircraft concepts: N+1, 771, 772, 808; 

N+2, 771, 772; N+3, 771, 772; N2A, 808; 

N2B, 808

Nap-of-the-Earth (NOE) flight, 159, 170

NASA Structural Analysis System (NASTRAN): aeroelastic effects research, 392, 

521–22; capabilities of, 328, 472, 476–

77, 531, 535; commercial versions, 484–

87, 531; competitor organizations, 486; 

development of, 152, 497, 501–2, 503, 506, 

535, 556; distribution of, 477, 489, 508; 

documentation, 477; early use of, 477–

83; elements development, 478, 479, 485, 

501–2, 514; equations of, 328; error corrections, 478, 479; government agency use of, 

481–82, 540–41; maintenance of, 478–79; 

management of, 477, 508; nonaerospace 

applications, 481, 535, 538–39; organization-specific versions, 484–86; origins and 

development of, 152, 468, 472–77; price of, 

477; reference sources, 543; requirements 

of, 475; Space Shuttle studies, 327–28; 

success of, 486–87; Systems Management 

Office (NSMO), 477, 478, 480, 484, 508; 

team members, 475–76; upgrades to, 478, 

479, 482–83, 531, 535; user community 

profile, 478, 480–81, 535; Users’ Colloquia, 

479–82, 484, 491n64, 497, 508, 536–37; 

X-29 aircraft model, 460

NASGRO program, 506

NASTPLT program, 556

NASTRAN. See NASA Structural Analysis 

System (NASTRAN)

National Academy of Sciences, 211, 215

National Advisory Committee for Aeronautics (NACA): aerodynamic research 

and development, 17, 17n30; aeronautics 

contributions, viii–x, 108, 171–72, 735–

36; aeronautics staff, 146; airfoil sections, 

29n56, 109, 115; computation tasks, 137–

38; cowling design, 734, 735–36, 809; Distinguished Service Medal, 107; formation 

of, viii, 361; Jones as aide, 11; rotary wing 

aircraft contributions, 136–37; Rotating 

Wing Subcommittee (Helicopter Subcommittee), 142; Special Subcommittee on 

Aircraft Noise, 737; technical intelligence 

collection target, 17n30; transformation 

into NASA, 146, 361, 737

National Aeronautics and Space Act, 487, 

852, 852n33

National Aeronautics and Space Administration (NASA): aeronautics contributions, viii–x, 64–65, 817–18; aeronautics 

research, return to, 253–54; Army, relationship with, 139; budget for aeronautics, 

785; Executive Order to support aeronautics research, 171–72; full-cost accounting methods, 166; Headquarters meeting, 
NASA’s Contributions to Aeronautics

930

472–74; Medal for Exceptional Scientific 

Achievement, 116; Outstanding Mechanical Engineer Award, 828n6; rotary wing 

aircraft contributions, 136–37; Software 

of the Year Award, 761; Space Act Award, 

846; space initiatives, vii–viii, 787; transformation of NACA to, 146, 361, 737

National Aeronautics Association, 93, 107, 

116

National Aeronautics Research and Development Policy, 171–72

National Aero-Space Plane (NASP), 230–31, 

241, 279, 336–40, 342, 419, 615–21

National Aero-Space Plane Joint Program 

Office, 338–39

National Air and Space Museum: Beyond 

the Limits gallery, 427, 432; delta test 

model donation, 15n26; HiMAT vehicle, 

689; Udvar-Hazy Center, 166, 638n15

National Business Aircraft Association, 121

National Carbon, 587

National Full-Scale Aerodynamics Complex 

(NFAC), 158, 165–66

National Hypersonic Flight Research Facility (NHFRF), 299, 605

National Medal of Science, 116

National Museum of the United States Air 

Force, 638n15, 698–99

National Oceanic and Atmospheric Administration (NOAA), 206–7

National Opinion Research Center, University of Chicago, 199

National Park Service, 201

National Register of Historic Places, 127–28

National Research Council, 248

National Rotorcraft Technology Center 

(NRTC), 168–69

National Science and Technology Council, 

Committee on Technology, 171–72

National Science Foundation (NSF), 549, 

823–24, 825, 827

National Sonic Boom Evaluation Office 

(NSBEO), 203

National Sonic Boom Evaluation Program, 

203–5

National Technology Transfer Center 

(NTTC), 490

NATO (North Atlantic Treaty Organization), 

213, 365

Naumann, Eugene C., 508

Navaho cruise missile, 289, 290, 569–70

Naval Air Development Center (NADC), 297, 

398–400, 540

Naval Air Warfare Center, 384

Naval Ocean Systems Center, 541

Naval Research Laboratory, 227

Naval Ship Research & Development

 Center, 481, 484, 540, 541

Naval Surface Weapons Center, Carderock 

Division, 157, 541

Naval Technical Mission, 29n56

Naval Underwater Systems Center, 481, 

540, 541

Naval Weapons Center, 540

Navier, Claude-Louis-Marie-Henri, 431–32

Navier-Stokes equations: Beyond the Limits gallery display, 427, 432; concepts governed by, 324, 432–33, 446; ENSAERO 

code, 493; LAURA code, 451; origins of, 

431–32; solution to, 323, 432–33, 434, 

436, 438–39, 446–47; turbulence, 450–

51; X-24C computed surface streamlines, 

448–49

Navy, U.S. (USN): Cooperative Test Organization/VECTOR, 705–6; Douglas D-558-2 

Skyrocket, 28; electronic ship control 

systems, 631; F-8C Crusaders, 654–55; 

fighter aircraft program, 61; fly-by-wire 
Index 

931

ship control systems, 723; Full Authority Digital Engine Control program, 755; 

Grumman F9F-9/F11F-1 Tiger, 104, 105; 

helicopter specifications, 145–46; International Test Organization (ITO), 700–701; 

NASA, relationship with, 139; NASTRAN 

Users’ Colloquia, 480, 481, 541; National 

Aero-Space Plane Joint Program Office, 

338–39; National Rotorcraft Technology 

Center, 168; Pacific Missile Range Facility, 

867, 869; sonic boom research, 206; spacesuits, 297; stability augmentation system 

research, 365; swept wing research, 22–23; 

Talos missile, 289; Tilt Rotor Research Aircraft (TRRA) program, 155; Vanguard satellite project, 500; Vietnam war aircraft 

losses, 644n32; X-15 hypersonic flight, 

290; X-31 aircraft, 707

NB-52/NB-52B Stratofortress, 343, 369, 

378–79, 568

NC-131H Total In-Flight Simulator (TIFS), 

402n62

Nellis Air Force Base, 194–95

Netherlands: Dutch-German Wind Tunnel (Duits-Nederlandse Wind Tunnel), 

164; V-2 missile shock waves, 184; wind 

energy, 827

Neural network project, 718–21

Nevada, 202

New Mexico, 827

Newton, Isaac, 429–30, 432

Newton-Raphson Parameter Identification, 397

Newton’s fluid dynamics laws, 429–30; second law, 429–30; sine-squared law, 429

New York Times, 216

New York University: Courant Institute, 111, 

471; sonic boom research, 213, 217, 222

Next Generation Clean Aircraft Power 

(NEXCAP) program, 815

NFK Engineering Associates, 484

Nibe A wind turbine, 848

Nicholas-Beazley, 11

Nicholls, J. Arthur, 332

Nickel, 575, 583

Nickel-based alloys. See also René 41 alloy: 

Alloy 454, 614–15; fasteners from, 619; 

gamma-prime, 574; oxygen and, 575; 

stress to cause failure, 613–14; superalloys, 573–76; thermal barrier coatings 

(TBC), 610; welds on, 575

1954 Interceptor initiative, 50

Niobium, 575

Nippon Univac Kaisha Ltd. (NUK), 484

Nitrogen oxides (NOx) emissions, 762, 790, 

794, 795–99, 813

Nixon, Richard M., 66–67, 107, 116, 217, 

219, 313, 603

NKF Engineering, 538, 539

Noise. See also sonic booms: airport noise 

studies, 206; design of aircraft and, 770–

72; engine noise research, 736–37; exhaust 

nozzle chevron, 808–9; FAA regulations 

and standards, 739–40, 741, 771; openrotor engines, 791; research activities at 

Lewis, 738; rotary wing aircraft, 163, 164–

65, 167–68; turbojet engines, 737; turbopropeller engines, 800, 802

Noise and engine improvements, 737–40; 

Acoustically Treated Nacelle program, 

740; Quiet Clean Short Haul Experimental Engine, 741–43; Quiet Engine program, 

740, 741; Steep Approach program, 740

Noise and Sonic Boom Impact Technology 

(NSBIT) program, 229–30, 231

Nomex, 767

Nomex nylon, 316

Nonlinear partial differential equations, 
NASA’s Contributions to Aeronautics

932

430–31, 432–33

Nonplanar lifting system, 118

Nordic Windpower, 848

Northam, Burton, 336

North American A3J-1 Vigilante, 32

North American A-5 Vigilante, 639

North American Air Defense Command 

(NORAD), 188

North American Aviation/North American 

Rockwell: B-70 contract, 191; business jet 

concept, 228–29; Space Shuttle orbiter 

contract, 313, 604; SST proposal, 210; 

supercritical wing contract, 112

North American B-45 aircraft, 634–35

North American B-70 aircraft, 191

North American F-86 Sabre: accidents with, 

24, 24n43; development of, 19n32, 23n39; 

F-86A, 26; F-86D Sabre, 102; F-86F, 26; 

Korean war use of, 4, 19, 26; pitch-up 

problems, 24–26, 36, 37, 413; reputation 

of, 23–24; sonic booms, 186–87; stability research activities, 637; tail configuration, 36, 413; variable stability aircraft, 

401; XP-86 Sabre prototype, 18, 635; YF86D low-tail testbed, 36–37

North American F-100 Super Sabre: development and manufacture of, 290; F-100C 

Super Sabre, 409; F-100D Super Sabre, 

409; flight control systems, 635; JF-100C 

Super Sabre, 402n62; landing simulator, 

403; sonic boom research, 194, 196; tail 

configuration, 413; YF-100A Super Sabre, 

34, 363; YF-100 Super Sabre, 37, 40, 41, 

42, 185

North American F-100 Super Sabre 

(F-100A): design of, 41, 42–43, 393–94; 

roll-coupling accident, 41, 42, 44, 393; 

roll coupling research, 408–9; wingspanto-fuselage length ratio, 37

North American F-107 aircraft, 372, 642, 

658n59

North American JF-100C Super Sabre, 

402n62

North American P-51D Mustang, 37

North American P-51 Mustang, 92

North American Sabre 45 aircraft, 36

North American simulators, 296

North American SM-64 Navaho cruise missile, 289, 290, 569–70

North American T-2C Buckeye, 114–15

North American X-15 aircraft: ablation 

cooling research, 417–18; accident with, 

299, 373–74; centrifuge program, 399–

400; contracts for, 293; contributions 

from program, 299; design confidence, 

307, 308; development of, 278, 283, 290–

93; at Dryden, 294; Dutch roll motions, 

411–12; emergencies, practice for, 296; 

flight control system, 298–99, 372–74, 376, 

379, 642; flight-testing, 294–95; flight time, 

299; funding for, 291; heating, structural, 

414–15; heat-sink structures, 415–17; inertial coupling (roll coupling), 409; instrumentation development, 298; landing of, 

405; landings of, 403–4; lift-to-drag ratio 

(L/D), 403; limit-cycle oscillation research, 

368; materials for, 292, 293, 295, 571, 592; 

pressure suits, 276, 278, 297, 298; reaction 

controls, 298; record hypersonic flight, 

278, 293, 299, 642; redundant systems, 

365–66; roll damper, 412; Round Two 

research aircraft, 54, 55, 290; side stick 

flight controller, 642, 658n59; simulator, 

295–97, 395–97, 403–4, 405, 409, 411–

12; sonic boom research, 197, 226; speed 

of, 415, 642; Stability Augmentation Systems, 365–66, 399; stability issues, 292, 

411–12; structural resonances, 375–76; 
Index 

933

tail configuration, 278, 411; thermal barrier coatings, 611; thermal damage, 294–

95, 416, 418; thermal data from, 517–18; 

X-15A-2 aircraft, 64, 335, 406, 417–18

North American XB-70 Valkyrie: aeroelastic 

effects research, 392; SCAT program and, 

108; sonic boom research, 201–2, 201n49, 

203, 204, 205, 210; XB-70-1 testbed, 201, 

202, 204; XB-70-2 testbed, 201–2, 204; XB70A, 56n107, 191

North American XP-86 Sabre, 18, 635

North American YF-86D low-tail testbed, 

36–37

North American YF-100A Super Sabre, 34, 

363

North American YF-100 Super Sabre, 37, 

40, 41, 42, 185

North Atlantic Treaty Organization (NATO), 

213, 365

Northrop, John “Jack,” 70

Northrop A-17A aircraft, 734

Northrop B-2 Spirit, 722, 753

Northrop F-5 fighter: F-5E Shaped Sonic 

Boom Demonstration, 245–48, 250, 252; 

sonic boom research, 245; tail configuration, 32, 413

Northrop F-89 Scorpion, 102

Northrop Grumman/Northrop Corporation: 

flying wings concept, 70; NASTRAN Users’ 

Colloquia, 480; Shaped Sonic Boom Demonstration (SSBD), 245–47

Northrop HL-10 aircraft, 56, 360, 368–69, 

378–79

Northrop M2-F2 lifting body, 407

Northrop N-156 aircraft, 32

Northrop Q-ball, 296, 298

Northrop Snark cruise missile, 52n100

Northrop T-38 aircraft, 32, 238, 239, 241

Northrop X-4 aircraft, 27–28

Northrop YB-49 flying wing jet bomber, 

19, 637

Northrop YF-17 Lightweight Fighter (LWF), 

388, 389, 664, 669

Nose cones/caps: ablation cooling, 286–88, 

581–82; ASSET vehicles, 594–95; General 

Electric Mark 3 nose cone, 288; General 

Electric RVX-1 nose cone, 288, 582; General Electric RVX-2 nose cone, 288, 294; 

heat sinks, 579, 581; hot structures, 286, 

579, 580–81; ICBMs, 278, 285–88, 305, 

417, 440–42, 572, 579–82; Space Shuttle 

orbiter, 318, 577; transpiration cooling, 

579–80; V-2 ballistic missiles, 571; X-20 

Dyna-Soar, 306, 308, 587–89, 598

Nose sensor, 298

No-slip condition, 431

Notch filter, 376, 377, 378

NPLOT program, 547

NT-33 aircraft, 642

NT-33A Shooting Star, 402

Nuclear Metals, 617

Numerical accuracy, 446–47

Numerical Evaluation of Stochastic Structures under Stress (NESSUS), 498–99, 514

Numerical Propulsion System Simulation 

(NPSS) program, 759–62

Numerical solutions, 442–44, 446–47

Nurflügeln (flying wings), 70

O

Oahu, Hawaii, 827, 842

Obama, Barack, 818

Oberammergau, 14, 48n94, 59

Oberschleißheim, 707

Oblique wing concept, 63–64, 515, 522–26, 

523n169, 530

Oblique Wing Research Aircraft (OWRA) 

program, 526
NASA’s Contributions to Aeronautics

934

Oestricher, Philip, 370, 658, 663, 665, 667

Ohio Aerospace Institute, 512

Ohio State University, 437, 768, 845

Oil embargo, 116, 743, 787–88, 817

Oil prices, 743, 788, 790–91, 816, 840–41

Oklahoma City sonic boom study, 199–200

Old Dominion University, 71, 542

Ontario Hydro, 538

Open Channel Foundation, 486, 490, 

490n62

Open Channel Software (OCS), 489–90, 

490n62

Open-rotor engines, 750–51, 786, 788, 791, 

800–803, 817

Optical data links, 659

Optimum Trajectory Research Experiment 

(OPTRE), 662

Orbital Sciences Pegasus rocket/booster, 

343–45

Organization of the Petroleum Exporting 

Countries (OPEC), 116, 743, 787–88

Orient Express, 230, 339. See also National 

Aero-Space Plane (NASP)

Orion spacecraft, 506

Orlando, 536

Orton, George, 344

Oscillations: aerodynamic damping, 363–67; 

limit-cycle oscillation, 367–69, 377, 378; 

pilot-induced oscillations (PIO), 369–72, 

378, 381, 409, 412, 661; pogo oscillations, 

328, 513; resonant motions, 375

Outstanding Mechanical Engineer Award, 

828n6

OVERFLOW code, 452, 453

Ozone layer research, 789–90, 795, 798

P

P-39 Aircobra, 23n37

P-47 aircraft, 333

P-51D Mustang, 37

P-51 Mustang, 92

P-63 Kingcobra fighters (L-39 aircraft), 

22–23, 23n37, 23n39

P-80 Shooting Star, 634

P 1101 aircraft, 58–59

Pacemaker reentry vehicle, 600

Pacific Gas and Electric, 839

Pacific Missile Range Facility, 417, 867, 869

Painter, Weneth, 369

Palmdale, 196, 707

Pamidi, P.R., 485

Parabolic partial differential equations, 

435n11, 436

Parameter Identification (PID), 504–5, 

504n113

PAREDYM program, 510

Paris Air Show, 161, 707, 750

Paris engine noise and sonic boom conference, 213

Parker Aerospace, 677

Parker Pen Company, 616

Parks, William, 856

Parrish, Philip, 617

Partial differential equations, 430–31, 432–

33, 434–36, 435n11, 446–47

Passive cooling, 606

Passive laminar flow control, 806–8

Pasteur, Thomas, Jr., 36

Pathfinder, 850, 855–59, 861, 862, 863

Pathfinder Plus, 850, 859–60, 862

PATRAN finite element preprocessor, 514

Patuxent River Naval Air Station, 706, 707

PB2-1S Superfortress (B-29), 29

PCA-2 autogiro, 137, 140

PCBoom, 230, 237, 257

PCBoom4 modeling, 255

Peabody, Paul R., 475

Pegasus rocket/booster, 343–45
Index 

935

Pennsylvania State University, 169, 251, 

850–51

Performance Seeking Control (PSC) program, 712–13, 757

Performance Test Engine (PTE), 345

Perl, William, 17n30

Perseus, 850, 853

Petaflop computers, 341, 439

Peterson, Bruce, 369

Pfielförmiges Tragwerk (Arrow-Shaped Lifting Surface), 6

Phenolic fiberglass, 607

Phenolic glass, 581

Phillips, William Hewitt, 38–39, 41, 42, 

44n86, 408

Phoenix Integration, Inc., 493

Piasecki Helicopter, 136

Pilot Activated Recovery System, 676

Pilot-induced oscillations (PIO), 369–72, 

378, 381, 409, 412, 661

Pilot-induced oscillations (PIO) suppression filter, 661–62

Pilotless Aircraft Research Division, Wallops Island, Virginia, 39–40

Pima Air and Space Museum, 672

Pinckney, Shimer, 335

Pinkel, Irving, 331

Piper Colt, 199

Piston engines, 634, 736

Pitcairn PCA-2 autogiro, 137, 140

Pitcairn YG-2 autogiro, 141

Pitch rocker technique, 666n70

Pitch-up: Douglas D-558 Skyrocket program, 28–32; Douglas XF4D-1 Skyray, 

32–33; Grumman F9F-6 Cougar, 33, 

33n60; Lockheed F-104 Starfighter, 34n61, 

413; McDonnell F-101 Voodoo, 33–34; 

stick-kicker, 34, 34n61; swept and delta 

wing aircraft, 19, 20–28, 27n51, 32–33, 

33n59, 50, 53; tail configuration and, 

32–37, 33n60, 34nn61–62, 413

Pitot nose inlet, 29, 50

Pitot tube, 701, 704, 705

Plasma Jet Subsonic Splash Facility, 588

Plasma Rotating Electrode Process, 617

Plexiglas (methyl methacrylate), 310–11, 

598

Plotkin, Kenneth J., 226–27, 246

Plum Brook Station, 831, 846

Pogo oscillations, 328, 513

Polaris missiles, 303

Polentz, Perry P., 474, 518

Polhamus, Edward C., 60–61, 61n115, 

651n49

Polymer composite engine fan casings, 

767–69

Popson, Raymond, 59–60

Porter, Lisa J., 253–54

Portland, Oregon, 537

Poststall maneuvering, 702–3, 703n144

Powder metallurgy, 616–17

Powell, C.W., 647

Power, Thomas, 300

Power-by-wire AFTI F-16 testbed, 676–79

Powered Flight Control Units (PFCUs), 642–

43, 677

Prandtl, Ludwig, 15

Pratt & Whitney compressor, 794

Pratt & Whitney F100 engine model derivative (EMD) engine, 711

Pratt & Whitney F100 turbojet engine, 710, 

712–13, 717, 754–56

Pratt & Whitney F117 engine, 795

Pratt & Whitney F119 engine, 712, 713

Pratt & Whitney F135 engine, 713

Pratt & Whitney Geared Turbofan engine, 

750, 751–52

Pratt & Whitney J57 turbojet, 100, 101, 102
NASA’s Contributions to Aeronautics

936

Pratt & Whitney J58 engines, 708, 709–10

Pratt & Whitney J75 turbojet, 103

Pratt & Whitney JT8D engine, 744

Pratt & Whitney JT9D-7A engine, 745

Pratt & Whitney JT9D engine, 611, 615, 

744, 745

Pratt & Whitney PW1000G engine, 752

Pratt & Whitney PW2000 engine, 795

Pratt & Whitney PW2037 turbofan engines, 

756, 795

Pratt & Whitney Rocketdyne/Pratt & Whitney: Advanced Control Technology for 

Integrated Vehicles Experiment (ACTIVE) 

project, 716–17; Advanced Subsonic Technology (AST) program, 796; combustor 

concepts, 795, 798, 799; Digital Electronic 

Engine Control (DEEC) program, 709–

10; digital engine control system, 754–

56; E Cubed technology use, 747, 792–95; 

geared turbofan engine, 800–801, 802; 

HyTech research program, 345; merger 

of Pratt & Whitney and Rocketdyne, 345; 

Numerical Propulsion System Simulation (NPSS) program, 760; Propulsion 

Controlled Aircraft (PCA) program, 715; 

scramjet engines, 345; thermal barrier 

coatings, 611–13; turboprop engine development, 800–802; Ultra Efficient Engine 

Technology (UEET) program, 762, 797

Pratt & Whitney Technology for Advanced 

Low Nitrogen Oxide (TALON) combustor, 799, 817

Pratt & Whitney TF34 turbofan engines, 

154, 158, 162, 163

Precision Aircraft Control Technology 

(PACT), 647–48

Precision Recovery Including Maneuvering 

Entry (PRIME). See Martin SV-5D (X-23) 

PRIME (Precision Recovery Including 

Maneuvering Entry)

Prescribed Velocity Distribution (PVD) 

blades, 612

Presidential Advisory Committee (PAC), 

192, 210–11

Presidential Executive Order, 171–72

Pressure suits, 276, 278, 297

Prien, Austria, 47

Princeton University, 211, 217

Principia (Newton), 429

Probabilistic Structural Analysis Methods 

(PSAM), 498

Profile drag, 117

Program for Analysis of Nonlinear

 Equilibrium and Stability (PANES), 552

Programming Structural Synthesis 

(PROSS), 511–12

Project 680J Survivable Flight Control System (SFCS), 379–80, 384, 646–48, 658n59

Project Independence, 823–24, 831

Project Independence Interagency Solar 

Task Force, 823–24

Projekt 100LDU testbed, 649–50, 692

Propeller, folding, 91

Propulsion-Airframe Integration project, 

763–64

Propulsion Controlled Aircraft (PCA) program, 714–15, 757, 758, 766

Propulsion-Only Flight Control System, 711

Propulsion systems. See also engines: 

advancements in, 3, 735, 769–70; computational fluid dynamics (CFD) and, 803; 

computers and technology improvements, 

757–59; propulsion efficiency, 794

Propulsion Systems Integration and Assessment project, 763

Proteus, 850, 853

Puccinelli, Edward F., 476

Puerto Rico, 827, 833
Index 

937

Pulse-type maneuvers, 397

Purifoy, Dana, 679

Puthoff, Richard L., 828, 830

Putnam, Palmer Cosslett, 826, 827, 846

PW1000G engine, 752

PW2000 engine, 795

PW2037 turbofan engines, 756, 795

Pyle, Richard S., 475

Pyrolysis, 577

Q

Q-ball, 296, 298

Q-felt insulation, 305, 586–87, 589, 590

Quarles, Donald A., 106

Quartz, 600

Quesada, Elwood R. “Pete,” 66, 192

Quiet Clean Short Haul Experimental 

Engine, 741–43, 758

Quiet Engine program, 740, 741

Quiet Spike device, 251–53, 256, 258, 261

Quiet Supersonic Platform (QSP), 244–48, 

249

R

R&D Magazine, 499

R-4 helicopter, 135, 142

Radar systems, 159–60

Radiative cooling, 588

Ragusa, Vivian, 706

Rahn, Robert O. “Bob,” 33

Ramjet engines: capabilities of, 278, 279, 

289, 571; scramjet engines, 279, 330–36, 

340, 342–46, 343n126

Rand Corporation, 284

Raney, J. Philip, 536

Rapid-solidification technique, 279, 617, 

620–21

RAPTOR/Pathfinder UAV, 850

Rate gyros, 298, 364–65, 366, 367

Rausch, Vincent, 343

Rauscher, Manfred, 38

RAVFAC program, 502

Rayleigh, Lord, 440, 466

Rayon, 577

Raytheon, 246, 249

Reaction controls, 298

Reaction-cured glass, 315

Reaction Motors XLR11 engines, 294

Reaction thrusters, 278

Reactor Centrum Nederland & Hazameyer, 

538

Reagan, Ronald, 229, 230, 279, 336, 339

Reclamation, Bureau of, 839

Recoverable Orbital Launch System 

 (ROLS), 334

Redner, Keith, 475

Redstone missile/booster, 288, 636

Redundant structures, 463–64

Redundant systems, 671–72; Boeing YC-14 

transport, 671–72; Digital Fly-By-Wire F-8 

program and aircraft, 380, 655–56, 658, 

659; flight control systems, 365–67; fly-bywire systems, 379–80, 384, 655–56, 658, 

659, 671–72; Grumman X-29 aircraft, 695; 

North American X-15 aircraft, 365–66; 

unmanned aerial vehicles (UAVs), 855–56

Reeder, John P. “Jack,” 641

Reentry vehicles: Air Force Flight Dynamics 

Laboratory research, 404, 406, 415; bluntbody theory, 54, 285–86, 291, 599; Dryden 

research activities, 495–96

Reeves Electronic Analogue Computer 

(REAC), 42

Regenerative fuel cell system (RFCS), 863–

64, 871, 872

Reid, H.J.E., 142

Reinhardt, Walter, 325

Relaxation method, 435
NASA’s Contributions to Aeronautics

938

Relaxed stability aircraft, 647–48, 651

Remington Rand ERA 1103 digital computer, 437

Remotely piloted vehicle (RPV), 236

Remotely Piloted Vehicle Facility, 662

René 41 alloy: ASSET vehicles, 594; composition of, 583–84; Dyna-Soar vehicles, 306, 

307, 308, 582–84, 586, 590, 591, 592, 597; 

jet engine use, 584; Space Shuttle orbiter, 

602; strength of, 583; strength versus temperature, 597

Renegar, William W., 474

Republic Aviation, 333, 334, 569

Republic F-84 aircraft, 333

Republic F-103 interceptor fighter, 569; 

XF-103 interceptor fighter, 289, 290, 333

Republic F-105 Thunderchief: design of, 

333, 409; development and manufacture 

of, 290; F-105B Thunderchief, 32, 44; inertial coupling (roll coupling), 409; sonic 

booms, 185, 216; tail configuration, 37, 

413

Republic P-47 aircraft, 333

Republic XF-103 interceptor fighter, 289, 

290, 333

Research Airplane Projects Panel (RAPP), 

14n23

Research Applied to National Needs 

(RANN) project, 824

Research Applied to National Needs 

(RANN) Symposium, 827–28

Resident Backup System (REBUS), 662

Resonances, structural, 374–79, 513

Reusable surface installation (RSI), 312–

23, 596–605

Reuther, James, 494

Reynolds number flows, 438

Rhoades, Carrie, 260

Rhode Island, 827

Rich, Quick Mix, Lean (RQL) combustor 

concept, 798, 799

Richardson, Holden C., 5

Richardson, L.F., 434–35

Richwine, David, 248

Ritchie, Virgil S., 95

Ritz, Walter, 466

RM-1 engine, 706

Robbins, Karen Risa, 853

Robbins, William, 830

Robert C. Byrd National Technology Transfer Center (NTTC), 490

Robert J. Collier Trophy. See Collier Trophy

Robins, A. Warner, 67

RoBo rocket bomber, 300, 301, 303

Roché, Jean, 13

Rocket Assisted Take Off solid rocket 

motors, 736

Rocketdyne, 345. See also Pratt & Whitney 

Rocketdyne/Pratt & Whitney

Rockwell Aerospace/Rockwell International: International Test Organization 

(ITO), 700–701; NASTRAN, development 

of, 484; NASTRAN, Users’ Colloquia, 480; 

X-31 aircraft, 700–701, 706

Rockwell B-1 bomber, 62, 62n118, 378; B-1A 

bomber, 392; B-1B bomber, 392

Rockwell Highly Maneuverable Aircraft 

Technology (HiMAT), 22, 325, 686–89

Rockwell Sabreliner 65 aircraft, 115

Rockwell Space Shuttle contract: awarding 

of, 313; thermal-protection tiles, 318, 319, 

321, 419; wind tunnel studies, 326

Rockwell X-31 Enhanced Fighter Maneuverability (EFM) aircraft, 699–707

Rocky Mountains, 200

Rodden, W.P., 483

Rodgers Dry Lake, 182, 394, 403, 688, 858

Rokide coatings, 611
Index 

939

Roll coupling (inertial coupling), 37–45, 

40n74, 44nn86–87, 393, 394–95, 407–9

Roll damper, 364–67, 412

Rolling tail, 278

Rolls-Royce: Allison/Rolls-Royce, 762, 797; 

combustor concepts, 799; Numerical Propulsion System Simulation (NPSS) program, 760; turbine blades, 609

Rolls-Royce RB 108 engines, 64

Roll stability, 37

Roll stick force gradient, 667

Ronald, Terence, 621

Root, L. Eugene, 27n51, 48n94

Rose, Peter, 572

Rosecrans, Richard J., 474

Rosenberg espionage investigation, 17n30

Rotary wing aircraft and helicopters: aircraft handling qualities criteria, 145–46; 

airfoil design for, 144; Ames as lead center for research, 156; Army research on, 

147; birth and growth of industry, 135–36; 

changes in program priorities and, 165–

66; Design Analysis Methods for Vibration (DAMVIBS), 508–11; development of, 

135; flying-quality criteria, 145–46, 148–

49; ground resonance, 144–45; instrument 

flight rules (IFR), 151; Korean war use of, 

138–39; military interest in, 141, 142, 168–

69; NACA/NASA contributions, 136–37; 

National Aeronautics Research and Development Policy, 171–72; National Rotorcraft Technology Center, 168–69; piloting 

techniques, 145; tilt rotor aircraft, civil 

transportation use of, 166–68; Vietnam 

war aircraft losses, 644n32

Rotary wing aircraft and helicopter research 

activities: Advanced Tiltrotor Transport 

Technology (ATTT) program, 167–68; 

Advancing Blade Concept (ABC) rotor, 

152–53, 157–58; aeroelastic stability, 149–

50; Airloads Program, 170–71; at Ames, 

137, 147, 148, 149–50, 152–53, 155–61, 

163–64, 492; by Army, 147, 150, 154–55, 

166, 168–69, 170–71; Bearingless Main 

Rotor (BMR), 157; blade motion studies, 

141–42; blade pressure research, 144, 

147–48; Circulation Control Rotor (CCR), 

156–57; contracted research, 153–54; Controllable Twist Rotor (CTR) concept, 153, 

156; Convertible Engine Systems Technology (CEST) program, 158, 162; crashtest activities, 152; at Dryden, 164, 165, 

170; Executive Order to support, 171–72; 

finite element analysis, 152, 162–63, 515, 

519–20, 529–30; Flying Laboratory for 

Integrated Test and Evaluation (FLITE), 

170; Heavy Lift Helicopter program, 158; 

hingeless rotor concept, 150, 152–53; icing 

environment, 162; at Langley, 137, 139–

45, 147–52, 154–55, 156, 161, 162–63; at 

Lewis/Glenn, 156, 158, 162; by manufacturers, 146; nap-of-the-Earth (NOE) flight, 

159, 170; noise research, 163, 164–65, 167–

68; radar systems, 159–60; resources for 

research, 155–56, 165–66, 169, 171; results 

and findings, publication or distribution 

of, 138, 143, 144, 154; rotary dynamics, 

144–45; rotor efficiency, 159; rotor-flow 

studies, 146; Rotor Systems Research 

Aircraft (RSRA) program, 154–55, 163–

64, 165; safety research, 167–68, 172; Tilt 

Rotor Research Aircraft (TRRA) program, 

155, 156, 160–62, 165; tilt rotor simulations, 151; Vehicle Technology Directorate 

(VTD), 162–63; X-Wing Stoppable Rotor, 

157, 158, 163–64; X-Wing Stoppable Rotor 

program, 165

Rotating-Wing Aircraft Meeting, Franklin 
NASA’s Contributions to Aeronautics

940

Institute, 142

Rotating Wing Subcommittee (Helicopter 

Subcommittee), 142

ROTONET, 164–65

Rotor Aerodynamic Limits survey, 515, 519–

20, 529–30

Rotorcraft Aircrew Systems Concepts Airborne Laboratory (RASCAL), 170

Rotorcraft Centers of Excellence (RCOE) 

program, 168–69

Rotorcraft Division, Army-NASA, 166, 170

Rotorcraft Industry Technology Association 

(RITA), 169–70

Rotor Systems Research Aircraft (RSRA) 

program, 154–55, 163–64, 165

Rotor Test Apparatus (RTA), 153

Round One research aircraft, 53–54, 

55n106, 301. See also Bell X-1 aircraft; Bell 

X-2 aircraft; Douglas D-558-2 Skyrocket

Round Three research aircraft, 55–56, 

55n106, 301. See also Boeing X-20 

Dyna-Soar

Round Two research aircraft, 54, 55n106, 

290, 301. See also North American X-15 

aircraft

Royal Aeronautical Establishment, 94

Royal Aeronautical Society (RAeS), 333

Royal Aircraft Establishment (RAE): 

Hawker Hunter T.12 aircraft, 649; Jaguar Active Control Technology (ACT), 670, 

690–91, 692; Short S.C.1 aircraft report, 

641; swept and delta wing aircraft, 13n20

Royal Air Force (RAF): Skyshield air defense 

exercises, 188; sonic booms, 186; VC10K 

tanker, 643

Rudder, 411, 412, 632, 633, 637, 703–4

Ruhrstahl Fritz X missiles, 638n15

Runge-Kutta time integration, 445

Runyan, Harry L., 217, 473

Rural Electrification Act, 826

Russell, Philip, 851

Russia, 22, 827–28

Rutan, Burt, 120–21, 619

Rutan, Dick, 121n106

Rutkowski, Mike, 524

RVX-1 nose cone, 288, 582

RVX-2 nose cone, 288, 294

Ryder, Jim, 647

S

S-61 Sea King, 154, 163–64

S-72 Rotor Systems Research Aircraft 

(RSRA), 155

S-76 helicopter, 156

Saab, 706

Saab 201 aircraft, 23n39

Saab 202 aircraft, 23n39

Saab AJ-37 Viggen, 391, 651n49

Saab J29 fighter, 23n39

Saab J32 Lancen (Lance), 23n39

Saab J35 Draken (Dragon), 45

Saab JAS-39 Gripen fighter, 706

Saab Safir aircraft, 23n39

Saab seat restraint, 9n13

Sabre 45 aircraft, 36

Sabreliner 65 aircraft, 115

Safe Slide, Project, 201

Safety factor, 387

Safety of aircraft, 506, 632

Safir aircraft, 23n39

Saint Augustine, 246–47

Saint Germain, France, 14

Saint Louis, 197–99

SAMIS program, 506

SAMSAN program, 549

San Antonio, 537

Sandusky, 831

San Francisco, 536
Index 

941

Sänger, Eugen, 54, 300

Sänger-Bredt, Irene, 54, 300

Sänger-Bredt Silbervögel (Silver Bird), 54, 

300

Sargent & Lundy, 538

Satellites, 305, 579, 581–82

Saturn launch vehicle, 814

Saturn V launch vehicle, 328, 513

Savannah, 252, 253, 801

Savino, Joseph M., 827–28, 828n6

Scaled Composites, 619, 853

Scherr, S.J., 448

Schkolnik, Gerard, 253, 714, 719

Schlichting, Hermann, 15

Schlieren photography, 97, 239, 241

Schneider, Edward, 662, 663

Schramm, Wilson, 310, 598

Schwartz, Ira R., 213, 216, 217

Science and Technology, Office of, 230–31

Science Museum, 641

Scott, H.G., 611

Scott, William C., 474

Scott Air Force Base, 197–98

Scramjet engines, 279, 330–36, 340, 342–

46, 343n126, 620

Seat restraint, three-point, 9n13

Seawolf submarines, 723

Seebass, A. Richard, 211, 212, 217, 218, 222, 

223, 226–27, 244, 244n181

Seeck, Dietrich, 700, 701

Self-adaptive flight control system (AFCS), 

372–74, 375, 376, 379

Self-Repairing Flight Control System 

(SRFCS), 713–14

Semispan model, 118, 119

Serafini, John, 331

Sewell, Charles A. “Chuck,” 695, 695n125

SH-3 Sea King, 151

Shaker vanes, 519

Shang, Joseph S., 448

Shaped Sonic Boom Demonstration 

(SSBD), 244–48, 249, 250, 252, 258

Shaped Sonic Boom Experiment (SSBE), 

248

Sharkey, John, 855

Shell Oil Company, 503

Shepard, Alan, 636, 653

Shepherd, Kevin P., 240, 257

Ship control systems, 723

Shock waves. See also sonic booms: bluntbody theory, 286, 415, 440–42; threedimensional flow-field calculations, 448; 

transonic flow, 94, 97, 98–99, 111, 445; 

unstarts, 336, 708–9, 754

Short Brothers, 641

Short S.C.1 aircraft, 639–41

Short Take-Off and Landing (STOL) aircraft, 670–72; Air Force transport competition, 114; engine innovations for, 743; 

F-15B/NF-15B, 258, 715–16; research on, 

146

Short Takeoff and Landing Maneuver 

Technology Demonstrator (S/MTD), 258, 

715–16

Shrout, Barrett L., 211

Shuttle Training Aircraft (STA), 404–5

Sideslip, 39–40, 42–43, 408, 409–10, 681–

82, 700, 701

Side stick flight controller, 298, 638, 638n15, 

642, 658, 658n59, 664, 665, 668

Sight, helmet-mounted target designation, 

673, 675

Sikorsky, Igor I., 9, 135, 136

Sikorsky Advancing Blade Concept (ABC) 

rotor, 152–53, 157–58

Sikorsky Aircraft Corporation: Design Analysis Methods for Vibration (DAMVIBS), 

508, 509n128, 510; helicopter design and 
NASA’s Contributions to Aeronautics

942

production, 136; PAREDYM program, 

510; RITA membership, 169

Sikorsky R-4 helicopter, 135, 142

Sikorsky S-61 Sea King, 154, 163–64

Sikorsky S-72 Rotor Systems Research Aircraft (RSRA), 155

Sikorsky S-76 Spirit, 156

Sikorsky SH-3 Sea King, 151

Sikorsky UH-60 Blackhawk: Airloads Program, 170–71; RASCAL configuration, 

170; UH-60A Blackhawk, 510, 519–20, 

529–30

Sikorsky VS-300 helicopter, 9, 135, 136, 142

Sikorsky X2 helicopter, 158

Sikorsky XH-59A Advancing Blade Concept 

(ABC) helicopter, 152–53, 157–58

Sikorsky YR-4B helicopter, 142, 143–44

Silane, 344–45

Silbervögel (Silver Bird), 54, 300

Silica ablative coating, 417–18

Silica fiber insulation, 305, 598–99, 600–601

Silica thermal-protection tiles, 310–18, 

599–605

Silicides, 603

Silicon carbide, 315, 317, 587, 599

Silicon dioxide, 599–600

Siliconized graphite, 306, 308, 587, 588

Silicon tetraboride (SiB4), 315–16

Silver Bird (Silbervögel), 54, 300

Silverstein, Abe, 810, 811, 814

Simonov, Mikhail Petrovich, 650n46

Simplifications, 463

Simulators and flight simulation: accuracy 

of, 396–98, 406; analog computers and, 

392–94, 405–7, 468; centrifuge, 398–400; 

digital computer simulation, 405–7; evolution of, 392–96; finite element analysis, 

515, 516; hybrid computer simulator, 406–

7; importance of, 392; Iron Bird simulator, 

655, 656–57, 659, 660; Iron Cross simulator, 298; landing-pattern trainers, 402–5; 

Link Trainer, 295, 392; moving-base cockpits, 398; reaction controls, 298; Space 

Shuttle orbiter, 371; variable stability aircraft, 401–2; visual displays, 407; X-15 aircraft, 278–79, 295–97, 395–97, 403–4, 405, 

409, 411–12; X-20 Dyna-Soar, 405–6

SINDA codes, 546

Sine-squared law, 429

Single-crystal blades, 614–15

Single-stage-to-orbit (SSTO) vehicles, 279, 

337

Single-string damper systems, 365, 412

Sioux City airliner crash, 757, 766

Skantze, Lawrence “Larry,” 337, 338

Skoog, Mark, 679

Skyshield air defense exercises, 188

SkyTower, 853

Slender-beam theory, 466–67

Slender wing theory, 12, 15

Slots, fixed and retracting, 414

SM-64 Navaho cruise missile, 289, 290, 

569–70

Small Business Innovation Research (SBIR) 

grant, 768

Smidinger, Peter A., 474

Smith, Apollo M.O., 29n56, 48n94, 70

Smith, Beauchamp, 826, 827

Smith, Burwell, 826

Smith, Rogers, 662, 679, 695, 701

Smith, Ron, 524

Smith-Putnam wind turbine, 826, 827, 828, 

835, 846

Smithsonian Institution. See National Air 

and Space Museum

Smolka, James W. “Jim,” 253, 255, 259, 260, 

679, 714, 719

S. Morgan Smith Company, 826
Index 

943

Snark cruise missile, 52n100

Snecma, 803

Sobieszczanski-Sobieski, Jaroslaw, 511

Society of Automotive Engineers (SAE)-

Aerospace, 138

Society of Experimental Test Pilots, 649

Software programs. See also NASA Structural Analysis System (NASTRAN): 

Backup Flight System software, 661; Computer Programs Abstracts Journal, 489; 

Computer Software Management and 

Information Center (COSMIC), 477, 488–

90, 490n62, 499, 508; digital fly-by-wire, 

381–82, 659–60, 671–72, 683–84; documentation, 489; finite element structural 

analyses, 158, 490n62, 502, 544–52; NASAdeveloped software, 490; Open Channel 

Software (OCS), 489–90, 490n62; prices, 

489, 490n62; Robert C. Byrd National 

Technology Transfer Center (NTTC), 490

Solano County, California, 839

Solar cells, 862–63, 868, 872

Solar Energy Research Institute, 836

Sol-gel deposition coatings, 613

Solid-propellant rocket boosters, 326, 327, 

329, 606

Solid rocket fuels, 503–4

Solovyev, Yevgeny, 650, 650n46

Sonic bang, 184

Sonic booms: complaints and damage 

claims, 186–89, 190, 197–98, 199–200, 

204–6, 210–11, 216–17; conferences, symposiums, and workshops, 209–10, 211–

14, 217, 231, 233–34, 241–42; control and 

reduction of, 181, 211–12, 218, 221–26, 

231–32; database on, 229, 238; doubleboom sounds, 184–85, 186; duration of, 

188–89; impact of, 185–88; low-boom/noboom technique, 255–57; magnification of, 

187; National Sonic Boom Evaluation Program, 203–5; natural, 182–83; prediction 

of, 212–13; prohibition of by Concorde, 

227–28; Quiet Spike device, 251–53, 256, 

258, 261; secondary booms, 227–28; signature of, 196–97, 198, 222–23; signature 

of, evanescent waves, 256; signature of, 

N-shaped, 185, 186, 197, 206, 223, 232, 

255–56; SST program research, 192–207; 

super booms, 187; survey of civilians 

exposed to, 206, 207, 240–41, 242, 249

Sonic Boom Mitigation Project, 251, 261

Sonic boom research activities: by Air 

Force, 194–203, 207; at Ames, 214, 223–

24; arrow-wing configurations, 209, 225; 

by Boeing, 224; caustics, 226–27, 228; 

Community and Structural Response 

Program, 197–99; computational fluid 

dynamics (CFD), 234–37, 246, 248, 253, 

260; computer program development 

and use, 224–25, 226, 230, 237, 255, 257; 

at Dryden, 227, 227n133, 236–39, 242–

43, 245, 251, 252–53, 255–57, 258–60; at 

Edwards, 193, 195–97, 200, 201–4, 206–7, 

229–30, 251; by FAA, 199–201, 206, 207; 

flight-testing, 227; High-Speed Research 

(HSR) program, 232–43; at Jackass Flats 

test site, 206–7; at Langley, 192–200, 203, 

208–9, 214–15, 217, 227, 239–40, 249, 

256–58; Lift and Nozzle Change Effects 

on Tail Shocks (LaNCETS), 258–60, 261; 

low-boom demonstrator, 249, 250–51, 

255, 261; low-boom/no-boom technique, 

255–57; by Navy, 206; results and findings, publication or distribution of, 207–

8, 213–14, 218–19, 221, 229, 233, 243, 

243n179; Shaped Sonic Boom Demonstration (SSBD), 244–48, 249, 250, 252, 

258; Shaped Sonic Boom Experiment 
NASA’s Contributions to Aeronautics

944

(SSBE), 248; simulators, 214–16, 240, 249, 

257–58; SST program research, 192–207, 

210; Supersonics Project, 254–60; techniques for, 218; at Wallops Island Station, 

194, 198, 206; wind tunnel studies, 208–9, 

208n66, 214, 222, 235

Sonic Boom Symposium, 211

Sorensen, Hans, 222

Soulé, Hartley A., 14, 14n23

Sound, speed of, 182

South Kensington, London, 641

Southwest Research Institute (SwRI), 498, 

514

Soviet Union (USSR): aerodynamic research 

and development, 17, 17n30; arms race 

with, 442; delta wing development, 45–46; 

fly-by-wire system research, 649–50; Projekt 100LDU testbed, 649–50, 692; space 

race against, 632, 644, 786; Sputnik, 288, 

301, 361, 786, 818; Supersonic Transport 

study, 66; swept and delta wing aircraft, 

9n12; technical intelligence collection, 17, 

17n30; variable-geometry wing research, 

62; Vietnam war, 644

Space Act, 487, 852, 852n33

Space Act Agreement, 761

Space Act Award, 846

Spacecraft Design Division, NASA, 56–57

Space race, 632, 644, 786

Space Shuttle Challenger, 329–30

Space Shuttle Columbia: accident with, 578; 

delay in launch of, 322–23; descent of, 322; 

launch of, 323; record hypersonic flight, 

278; significance of return of, 289; structural testing, 329–30; thermal-protection 

tiles, 313, 319–23

Space Shuttle Enterprise: approach and 

landing evaluations, 661; pilot-induced 

oscillations (PIO), 370–71, 381, 661; 

structural testing, 329; wind tunnel studies, 327

Space Shuttle Main Engine (SSME), 337

Space Shuttle orbiter: aluminum construction, 310, 602, 606–9; AP-101 computers, 

659; approach and landing evaluations, 

661; Backup Flight System software, 661; 

computational fluid dynamics study, 426, 

448, 450; computational simulations, 325–

26; concept development, 56–57; contract 

for, 313, 604; control stick feedback, 370–

71; design development tests, 317–18; 

external tanks, 326, 327, 329, 578; finite 

element analysis, 327–28, 506; flight control system, 278; flight control systems, 

367; fly-by-wire system, 753; fuel cells, 

814; heating, structural, 415, 572; hot 

structures thermal protection, 310–13, 

601–3; hypersonic flow pattern, 426; insulating foam, 578; landing simulator, 404–5; 

nose cap, 318, 577; pilot-induced oscillations (PIO), 370–71, 381; pilot-induced 

oscillations (PIO) suppression filter, 661–

62; pogo (oscillation), 328; record hypersonic flight, 278, 299; Shuttle Orbiter 036 

series, 57; Shuttle Orbiter 040C design, 

57; side stick flight controller, 658n59; 

solid-propellant rocket boosters, 326, 327, 

329, 606; speed of, 415; strain isolator pad 

(SIP), 316, 319, 320, 321; structural verification study, 554–55; thermal modeling, 495; thermal protection, 576, 577–78; 

thermal-protection tiles, 278, 288, 310–23, 

419, 582, 598; thermal-protection tiles, 

bonding of, 319, 320–23, 609; thermal-protection tiles, coatings on, 315–16; thermalprotection tiles, densification of, 320–21; 

thermal-protection tiles, gaps between, 

318, 319–20; thermal-protection tiles, 
Index 

945

installation of, 319–23, 419; thermal-protection tiles, shaping of, 314–15, 321; thermal-protection tiles, testing of, 316–19; 

thermal-protection tiles, thermal expansion and, 315; three-dimensional shock 

wave shape, 448; titanium construction, 

608–9; wind tunnel studies, 323–30; wing 

configuration, 301; wing leading edge protection, 577–78

Space Shuttle program: approval of, 313, 

603; cancellation of, 765; concurrent 

development, 322–23; cost overruns, 229, 

602; Digital Fly-By-Wire F-8 program to 

support, 661; environmental impact statement (EIS), 226; Johnson Space Center 

activities, 506; sonic booms from, 226, 

226n130; spacesuits for, 297

Space station, 243, 311–12, 506

Spacesuits, 276, 278, 297

SPAD XIII fighter, 37

Spark discharge system, 214

Special Subcommittee on Aircraft Noise, 

737

Specific fuel consumption (SFC), 745–46, 

792–93

Spectralab, 862

Speed of sound, 182

Spera, David A., 830

Sperry Marine, 538

Sperry Support Services, 484, 656, 656n57, 

660

Spinoff, 512

Spin Rig (Glenn [Lewis] Research Center}, 

555

Spirit of St. Louis, 591

Splat-cooling, 617

Split-S maneuvers, 702, 702n139

Sputnik, 288, 301, 361, 786, 818

SR-71 Blackbird: accident with, 392; 

aerodynamics of, 70; aeroelastic effects, 

391–92; digital engine control system, 754; 

engine inlet unstarts, 708–9, 754; heating, structural, 413; JP-7 fuel, 346; simulator, 407; sonic boom research, 202–3, 

204, 205–6, 236–38, 239, 242–43, 245, 261; 

spacesuit for, 297; speed of, 413; thermal 

barrier coatings, 611; thermal loads project, 516; titanium construction, 312, 570–

71, 602

STA-099, 329–30. See also Space Shuttle 

Challenger

Stabilator tail, 34–35, 670

Stability augmentation system research, 

636–37

Stability augmentation systems (SASs), 

364–66, 399, 659, 660

Stability Management Control (SMC), 

717–18

Stabilizer, 26, 35, 362–63

Stack, John P.: aeronautics contributions, 

107; Anglo-American partnership and 

Swallow aircraft, 60, 61n115; area rule 

and, 99, 101; Bell X-1 aircraft, 108; Collier Trophy, 93, 96; SCAT program, 193; 

slotted transonic tunnel, 93, 95–96; supersonic aircraft research, 94, 108; Supersonic Transport Research Committee 

(STRC), 65–66

Stage 1 aircraft, 739

Stage 2 aircraft, 739, 741

Stage 3 aircraft, 739

Stage 4 aircraft, 739

Stainless steel, 210, 590

Stall conditions, 67, 67n127, 412–14, 

666n70, 711

Stall margin, 711n163

Stanford Research Institute (SRI), 203, 

205, 215
NASA’s Contributions to Aeronautics

946

Stanford University, 71, 244, 315, 341

Star Clipper, 311, 599, 600

Star Clipper spacecraft, 311, 599, 600

Starship, 121

Star Wars (Strategic Defense Initiative), 

279, 336–40

Static Nonlinear Analysis of Shells of Revolution (SNASOR), 506–7

Static stability: directional, 408, 409, 410–

11, 648, 651; longitudinal, 648, 664, 669, 

690, 692; relaxed stability aircraft, 647–

48, 651

Statically unstable aircraft, 633

Stealth aircraft, 70, 384, 721–22

Stecura, Stefan, 611

Steel, 569, 620

Steel, stainless, 210, 590

Steep Approach program, 740

Stefan-Boltzmann law, 585

Stennis Space Center: finite element analysis activities, 491n64; Finite Element Modeling of Piping Systems program, 552; 

mission of, 491n64

Stephenson, Geoffrey, 41

Stewart, James, 855

Stick kicker, 34, 34n61, 413

Stick shaker, 413

Stiffness matrix, 470–71

Stinger, 261

Stockpile Stewardship program, 498

Stokes, George Gabriel, 431–32

Stores Management System (SMS), 672, 681

Strain equations, 464–65

Strain gages, 370, 385–86, 528, 534, 554, 555

Strain isolator pad (SIP), 316, 319, 320, 321

Strake flaps, 526, 694

Strategic Air Command (SAC) missions, 

188, 197

Strategic Defense Initiative (SDI), 279, 

336–40

Stress distribution, 471

Stress equations, 464–65

Structural analysis. See also finite element 

structural analyses; NASA Structural Analysis System (NASTRAN): aircraft built 

without, 464; digital computer development and, 468–72; direct analog computers, 466–68; discretization of structures, 

466, 469, 470–72; energy method, 466; 

flexibility matrix, 471; free body diagrams, 

464–65; goal of, 461; hybrid direct-analog/

digital computers, 468; loads prediction, 

461–62, 462n1, 465–66; matrix method, 

466, 470–71; NASA’s contributions to, 461; 

principles of, 462–63; prior to computers, 462–66; redundant structures, 463–64; 

rotary wing aircraft, 152; simplifications, 

463; stiffness matrix, 470–71; stress distribution, 471; topics covered by, 462n1; unit 

displacement method, 469, 470

Structural analysis groups, 461

Structural Analysis of General Shells

 (STAGS), 514, 551–52; STAGS-C1, 551

Structural Analysis Routines (STARS), 496, 

525–26, 528, 556

Structural Dynamics Laboratory (Ames 

Research Laboratory), 553

Structural Dynamics Research Corporation, 550

Structural Performance and Redesign 

(SPAR), 495, 498, 512

Structural resonances, 374–79, 513

Structural Tailoring of Engine Blades 

(STAEBL), 545–46

Structural-Thermal-Optical (STOP) program, 501–2

ST-SIZE program, 512

Subsonic aircraft, 109, 117
Index 

947

Sukhoi: Experimental Design Bureau 

(OKB), 650n46; supersonic business jet 

(SSBJ) concept, 249

Sukhoi Su-7U aircraft, 649–50

Sukhoi Su-17/22 Fitter, 62

Sukhoi Su-24 Fencer, 62, 62n118

Sukhoi Su-27 Flanker, 692

Sukhoi Su-37 FSW prototype, 22

Sukhoi T-4 aircraft, 649

Sukhoi T-10 aircraft, 650, 650n46

Sullivan, Brenda M., 257

SunPower Corporation, 860, 862, 872

Suntan project, 811–13

Superalloys, 573–76, 602. See also cobaltbased alloys; nickel-based alloys

Superchargers, 736

Supercomputers, 438–39, 446

Supercritical wing: aviation industry use of, 

114–15; benefits and importance of, 89–90, 

115–16, 126, 805–6, 817; development of, 

65, 68, 109–12, 786, 805; F-8 Supercritical wing, 89; models, 110–11; naming of, 

111; patent for, 114; publication of findings, 111, 113; recognition of, 116; testing of, 112–14; X-29 aircraft, 693, 694, 

695, 695n126

Supercritical Wing Technology, 113

Supermarine Swift, 28

Supersonic Aerospace International (SAI), 

249–50

Supersonic aircraft and flight. See also area 

rule: Air Force research activities, 182; 

ban of supersonic flight by FAA, 220–21, 

220n112, 224; challenges of, 262, 361–62; 

design of, 93, 96, 100–105, 440; engine fan 

casings, 769; future of, 260–62; materials 

and structures for, 569–73; ozone layer 

damage from, 789–90; passenger plane 

development, 181, 189–90; proposals for, 

245–46

Supersonic business jets (SSBJ): companies working on concepts, 249; concept 

for, 228–29; consortiums for exploring, 

249–50; economic conditions and, 260–

61; FAA sonic boom restrictions and, 249; 

future of, 260–62; sonic boom research, 

256–57

Supersonic combustion testing, 331–32

Supersonic Commercial Air Transport 

(SCAT) program: cranked arrow wing 

development, 69; deep stall, 67, 67n127; 

design of aircraft, 67, 108–9, 108n61; 

goals of, 108–9, 189–91; SCAT-4, 67, 108–

9, 191; SCAT-15, 67, 191; SCAT-15F, 67, 69, 

221; SCAT-15F model, 180; SCAT-16, 191; 

SCAT-17, 191; sonic boom research, 193

Supersonic Cruise Aircraft Research program. See Supersonic Cruise Research 

(SCR)/Supersonic Cruise Aircraft 

Research (SCAR)/Advanced Supersonic 

Technology (AST) program

Supersonic Cruise and Maneuver Prototype (SCAMP), 69, 807. See also General 

Dynamics F-16 Fighting Falcon

Supersonic Cruise Industry Alliance (SCIA), 

250

Supersonic Cruise Research (SCR)/Supersonic Cruise Aircraft Research (SCAR)/

Advanced Supersonic Technology (AST) 

program: AST Project Office, 220; cancellation of, 228; conferences on, 225–26, 

228–29; design of aircraft, 221; establishment of, 219; funding for, 221; goals of, 69, 

219, 220; name changes, 219; sonic boom 

research, 220, 224, 227

Supersonic Pressure Wind Tunnel, 208, 

208n66, 214 Supersonics Project, 254–

60, 261
NASA’s Contributions to Aeronautics

948

Supersonic Transport (SST): cancellation 

of, 68–69, 217, 218, 519, 789; Concorde 

Supersonic Transport (SST), 3, 45, 56, 

66; contract for, 211, 217–18; data to support, 519; design of, 217–18; flight locations, 218; flutter suppression research, 

388; funding for, 218, 219; opposition to, 

216–17; Presidential Advisory Committee 

(PAC), 192; request for proposals (RFP), 

210; research activities, 192; research 

activities, sonic booms and, 192–207, 210; 

research activities, Whitcomb’s work and, 

108–9, 108n61; research activities at Langley, 65–69; responsibility for, 191–92; XB70A testbeds, 191

Supersonic Transport Development Office, 

FAA, 199

Supersonic Transport Research Committee 

(STRC), 65–66, 67, 108, 108n61

Supersonic Vehicles Technology program, 

249

Supersonic wind tunnels, 280–81

Survivable Flight Control System program 

(F-4 SFCS program), 379–80, 384, 646–

48, 658n59, 660

Sutton, George, 286, 581

SV-5D (X-23) PRIME (Precision Recovery 

Including Maneuvering Entry), 27n51, 

278, 288, 304, 417, 418, 606–8

SV-5P (X-24A) lifting body, 27n51, 407, 607

Swallow aircraft, 60–61

Sweden: AFTI/F-16 aircraft, 679; Cooperative Test Organization/VECTOR, 706; 

Swedish Air Force, 679; wind energy, 839, 

848; X-31 aircraft, 706

Sweep theory, 12, 15, 15n27

Swept and delta wing aircraft: aeroelastic effects, 390; American research and 

designs, 7–19, 72–73, 93, 108; Blended 

Wing-Body (BWB), 2, 70–72; British aircraft, 27, 28; challenges of, 18–28, 27n51, 

50; conical camber, 51–53; delta wing 

development, 12–13, 13n20, 45–49, 51, 

53, 67; Dutch roll motions, 19, 50, 410–

12; early history of, 4–7, 414; flat-bottom 

delta aircraft, 55–57, 301, 302; flat-top 

delta aircraft, 55–56, 300–301; forwardswept wing (FSW), 21–22, 524, 527; German development of, 7, 8, 13, 13n20, 

14–15, 15n27, 72–73, 93; hypersonic aircraft, 53–57; iconic status of, 4; inertial 

coupling (roll coupling), 37–45, 40n74, 

44nn86–87, 393, 394–95, 407–9; Langley 

Memorial Aeronautical Laboratory testing, 13–15; Mach angle of wing, 15–16, 

18; Mach number of wing, 5, 12, 14, 18, 

52, 53; nose-rise problems, 414; pitch-up 

problems, 19, 20–28, 27n51, 32–33, 33n59, 

50, 53; sections, thickness of, 13; stall, 414; 

Supersonic Transport (SST), 65–69; test 

model, 14, 15n26; variable-geometry wing 

aircraft, 57–64, 61n115, 62n118; vortex lift 

research, 651n49

Sylvanus Albert Reed Award, 116

Synthetic aviation kerosene, 815–16

Synthetic fuel research, 815–17, 816n114

System Identification/Parameter Identification (PID), 504–5, 504n113; Autonomous 

Frequency Domain Identification (AUFREDI), 504, 505; MODE-ID program, 504

Systems Research Aircraft (SRA), 384–85, 

386, 676

System verification units (SVU), 839

Syvertson, Clarence A., 54–55, 56, 56n107, 

300–301

Szalai, Kenneth J., 654, 659, 663
Index 

949

T

T-2 aircraft, 691–92

T-2C Buckeye, 114–15

T-4 aircraft, 649

T-10 aircraft, 650, 650n46

T-34C Mentor, 481

T-38 aircraft, 32, 238, 239, 241

Ta 183 aircraft, 27n51, 36

Tactical Fighter Experiment (TFX), 61

Tactical Fighter Experimental (TFX), 61

Tactical Mobility Requirements Board 

(Howze Board), 139

Tail configuration. See also T-tail aircraft: 

adjustable horizontal tail, 362–63, 635, 

666n70; all-moving tail, 34–35, 35n65, 

73, 362–63, 363n4, 665–66; Dutch roll 

motions and, 411; pitch-up and, 32–37, 

33n60, 34nn61–62, 413; rolling tail, 278; 

stabilator tail, 34–35, 670

Tailless aircraft, 4–5, 13, 27, 27n51, 32–33, 

703–4. See also swept and delta wing 

aircraft

TALON (Technology for Advanced Low 

Nitrogen Oxide) combustor, 799, 817

Talos missile, 289

Tampa, 537

Tantalum, 575, 584

TAPS (Twin Annular Pre-mixing Swirler) 

combustor, 798–99, 817

TBM Avenger, 695n125

Technical Report 1007, 465–66

Technical Report server, 168

Technical Research Development Institute 

(TRDI), 692

Technology for Advanced Low Nitrogen 

Oxide (TALON) combustor, 799, 817

Technology transfer activities: Applications Engineering process, 488; Computer 

Software Management and Information 

Center (COSMIC), 477, 488–90, 490n62; 

at Dryden, 496; Industrial Application 

Centers, 488; NASA-developed software, 

490; Open Channel Foundation, 486, 490, 

490n62; Open Channel Software (OCS), 

489–90, 490n62; priority of, 487; publications, 487–88; Robert C. Byrd National 

Technology Transfer Center (NTTC), 490; 

Technology Utilization (TU) program for, 

487–88; Technology Utilization Officers 

(TUOs), 488

Technology Utilization (TU) program, 

487–88

Technology Utilization Officers (TUOs), 488

Teflon, 305, 582

Tel Aviv University, 512

Teledyne-Ryan Aeronautical, 688, 760

Teledyne-Ryan autopilots, 687–88

Teledyne-Ryan BQM-34E Firebee II RPV, 

245

Teledyne-Ryan BQM-34 Firebee II RPV, 236, 

242, 388–89

Teller, Edward, 439

Temperature effects. See aerothermodynamics; high-temperature entries; thermal entries

Tennessee Eastman Company, 481, 538, 539

Teraflop computers, 439

Terrain system with moving map, 675–76

Test Pilot School, USAF: Digital Flight Control for Tactical Aircraft (DIGITAC), 382–

83, 684–86; Have Coupling project, 685; 

sonic boom research, 248, 257

Tether, Anthony “Tony,” 337

Texas Instruments, 481, 539

TEXLESP-S, 503–4

TF30 turbofan engines, 754

TF34 turbofan engines, 154, 158, 162, 163

TF-104G aircraft, 688
NASA’s Contributions to Aeronautics

950

Theodorsen, Theodore, 15

Thermal Analyzer, NASTRAN, 501–2

Thermal barrier coatings (TBC), 610–13

Thermal Loads Laboratory (Dryden Flight 

Research Center), 553–55

Thermal loads project, 515, 516–19, 529

Thermal protection. See

 aerothermodynamics

Thermal Radiation Analysis System (TRASYS), 506

Thermodynamics, first law of, 432, 433

Theseus, 850

Thigpin, David J., 670

Thinking sessions, 98

Thiokol XLR99 engines, 294

Thomas, Arthur, 337–38

Thomas, Charles L., 223, 226

Thomas, Jim, 452

Thomas, Ronald, 827–28, 830

Thomas, Russell, 772

Thor-Able missiles/boosters, 288, 594

Thor-Delta missiles/boosters, 308, 309, 310, 

594

Thoria, 587–88

Thorium, 308

Thorium oxide, 595

Thor missiles/boosters, 308, 309, 310, 592, 

594

Three-axis flight control system, 633

Three-dimensional flow-field calculations, 

447–49

Thrust vectoring, 700, 701, 703–4, 705–7, 

716–17

Thrust-vectoring exhaust nozzles, 715–16, 

717

Thrust-vectoring paddles, 699, 700, 706

Tiernay Turbines, 539

Tilt rotor aircraft, civil transportation use 

of, 166–68

Tilt Rotor Research Aircraft (TRRA) program, 155, 156, 160–62, 165, 166. See also

Bell V-22 Osprey Tilt Rotor

Time-marching process, 443–44, 445–46

Tinker Air Force Base, 199

Titan III-C missile, 302

Titan III missile/booster, 400

Titan II missile/booster, 302, 328

Titan I missile, 301

Titan intercontinental ballistic missile 

(ICBM), 286

Titanium: aircraft built with, 570–71, 616, 

620; B-120 version, 570; Beta-21S, 617–

18, 621; bonding to, 609; characteristics 

of, 615–16; chemical incompatibilities, 

570–71; chemical milling, 569; experience working with, 312, 602; high-speed 

aircraft and, 569–71; hot structures, 312; 

hypersonic aircraft and flight, 293; rapidsolidification technique and, 279; sonic 

booms and, 210; Space Shuttle orbiter, 

608–9; SR-15 Blackbird construction, 312, 

570–71, 602; superalloy addition, 575, 583; 

superalloys compared to, 574; X-15 aircraft, 295

Titanium carbide, 307, 590

Titanium Metals, 617

Toll, Thomas, 291

Topp, L.J., 471

Tornado interceptor and strike fighter program, 62, 649

Total In-Flight Simulator (TIFS), 402n62

TRACOR, Inc., 205–6

Transatmospheric flight: aircraft for, 279; 

development of technology for, 277–

79; hypersonic flight and, 278–79; reentry challenges, 285–89, 291–92; stability 

issues, 292

Trans Atmospheric Vehicles (TAV), 337
Index 

951

Trans Atmospheric Vehicles (TAV) Program 

Office, 337, 339

TRansfer ANalysis Code to Interface Thermal Structural (TRANCITS), 546

Transonic Aerodynamic Branch, 94, 108, 

125

Transonic aerodynamic effects, 5

Transonic Aircraft Technology (TACT) program, 113, 386

Transonic Dynamics Tunnel (TDT), 149–

50, 161, 164

Transonic flow and airfoils, 444–46

Transonic Pressure Tunnel (TPT), 109–10, 

118, 119, 123, 127–28

Transonic regime: aircraft design, 93–94, 

362–63, 444–46; challenges of, 91–92; 

importance of operation in, 94; Mach limits, 92; research tools, 92–93; shock waves 

in, 94, 97, 98–99, 111, 445; transonic drag 

research, 94–100; transonic flow, visualization of, 98; wind tunnel research, 92–97

Transonic transport fuselage, 127

Transonic transport wing configuration, 52, 

52n100, 68

Transonic Tunnel (High-Speed Tunnel), 

95n16

Transpiration cooling, 579–80

Transportation, Department of: Climactic Impact Assessment Program (CIAP), 

789; NASTRAN use by, 481; ozone layer 

research, 789–90; SST program, 192; 

supersonic research, 219

Transport Canada, 251

Transport delay, 661

TRASYS Radiative Heat Transfer program, 

550

Tremaine, Stanley A., 337

Tridymite, 600

Trim tabs, electrically operated, 633, 634–35

Trippensee, Gary, 701, 705

Truckee, California, 850

Tsien, Hsue-shen, 12, 14, 571

T-tail aircraft: deep stall, 67n127, 413; F-101 

Voodoo, 33–34; longitudinal stability, 

36–37; performance and, 33, 413

Tullahoma, Tennessee, 280, 291, 787

Tungsten, 308, 575, 584, 587, 595

Tungsten carbide, 592

Tupolev Bear, 102

Tupolev Tu-22M Backfire, 62

Tupolev Tu-144 aircraft, 789

Tupolev Tu-144 Charger Supersonic Transport (SST), 66

Tupolev Tu-160 Blackjack, 62, 62n118

Turbine blades: environmental barrier coatings, 612–13; fabrication of, 613–15; materials for, 609–10; thermal barrier coatings, 

610–13

Turbofan engines: Digital Electronic Engine 

Control (DEEC) program, 709–10; materials for, 571; Quiet Engine program, 737, 

741; TF30 turbofan engines, 754; TF34 

turbofan, 154, 158; TF34 turbofan engines, 

162, 163

Turbojet engines: General Electric J79 

turbo jet, 104; invention of, 3, 90; noise 

from, 737; Pratt & Whitney J57 turbojet, 

100, 101, 102; Pratt & Whitney J75 turbojet, 103; Quiet Engine program, 737, 741; 

speed capabilities, 289, 571; transonic 

flight with, 92

Turbopropeller engines: Advanced Turboprop (ATP) project, 744, 747–50, 786, 788, 

789, 790–91, 799–803; Collier Trophy for, 

748, 800; noise from, 800, 802; operation 

of, 747–48; public acceptance of, 747

Turbulence, 340–42, 439, 450–51

Turner, M.J., 471
NASA’s Contributions to Aeronautics

952

Twin Annular Pre-mixing Swirler (TAPS) 

combustor, 798–99, 817

Twining, Nathan F., 101

Typhoon, 692

TZM molybdenum: ASSET vehicles, 594; 

X-20 Dyna-Soar, 584–87, 588, 594

U

U-2 spy plane, 297, 811, 812–13

Udvar-Hazy Center, National Air and Space 

Museum, 166, 638n15

UH-1 Huey, 147

UH-60 Blackhawk: Airloads Program, 170–

71; RASCAL configuration, 170; UH-60A 

Blackhawk, 510, 519–20, 529–30

Ulster Folk and Transport Museum, Flight 

Experience exhibit, 641

Ultra Efficient Engine Technology (UEET) 

program, 613, 762–66, 786, 790, 797–98

Ultra-High Bypass Turbofan engine, 750–51

Unconventional (decoupled) control modes, 

652, 669–70

Unducted Fan (UDF) engine, 749, 750, 

801–2

Union Carbide, 601

Unitary Wind Tunnel: at Ames, 214, 235, 

586; at Langley, 235, 585–86

Unit displacement method, 469, 470

United Aircraft Corporations, Vought-Sikorsky Aircraft Division, 8–9

United Airlines Flight 232 crash, 757, 766

United Nations International Panel on Climate Change (IPCC) report, 790, 790n22

United States Air Force. See Air Force, U.S. 

(USAF)

United States Air Force Museum, 202

United States Army. See Army, U.S. (USA)

United States Army Air Forces (USAAF). 

See Army Air Forces (AAF)

United States Ballistic Missile Defense 

Organization, 850

United States Capital, 11

United States Forest Service, 200

United States House of Representatives. See

Congress, U.S.

United States Marine Corps. See Marine 

Corps, U.S. (USMC)

United States National Park Service, 201

United States Navy (USN). See Navy, U.S. 

(USN)

United States Senate. See Congress, U.S.

United Technologies Corporation, 799, 839

Universal Analytics, 484, 486, 538

Universal Automatic Computer (UNIVAC) 

computers, 436, 469, 484

University of Akron, 768

University of California at Berkeley, School 

of Engineering, 162, 471

University of Chicago, 199, 490

University of Colorado, 244

University of Florida, 71

University of Georgia, 488, 542

University of Iowa, 549

University of Manchester, 94

University of Maryland at College Park, 169, 

452, 542

University of Michigan, 332, 406

University of Missouri, 542

University of Naples, 542

University of New Brunswick, 542

University of Pennsylvania, 436, 468–69

University of Southern California, 71

University of Texas at San Antonio, 498–99

University of Toronto, 225, 240

University of Virginia, 512, 538, 542

University of Wichita, 845

Unmanned aerial vehicles (UAVs): advantages of UAVs, 861–62; Boeing X-48B, 
Index 

953

2, 71–72; Centurion, 850, 856, 859, 862, 

864–66; control systems, 856; development of, 71, 823, 849–50, 852, 853, 870–

71; flight termination, 858, 858n48; fuel 

cells, 856, 863–64, 869, 871; hardware reliability, 859, 869; Helios Prototype, 850, 

853, 856, 862, 863–64, 866–72, 866n61; 

Mishap Investigation Board (MIB) report, 

870–71; Pathfinder, 850, 855–59, 861, 862, 

863; Pathfinder Plus, 850, 859–60, 862; 

record altitude flights, 859, 860; redundant systems, 855–56; rotary wing, 136; 

sensors and communication equipment, 

854; solar cells, 862–63, 868, 872; solarpowered, 850, 857, 858, 860, 872; weather 

and, 858–59

Unstable aircraft, 379–80, 383, 526, 631, 

632, 633, 651–52, 753

Unstarts, 336, 708–9, 754

Upwinding, 445

USSR. See Soviet Union (USSR)

V

V-2 ballistic missiles, 183–84, 280, 282, 571, 

635–36

V-22 Osprey Tilt Rotor, 136, 155, 161, 166

VAK-191B aircraft, 649

Vandenberg Air Force Base, 607

Vanderplaats, Garret, 493

Van Dyke, Milton O., 442

Vanguard satellite project, 500

Variable cycle engine (VCE), 220

Variable-geometry wing aircraft, 57–64, 

61n115, 62n118

Variable stability aircraft, 401–2

Vari-Eze, 120–21

Vectoring, Extremely Short Takeoff and 

Landing Control and Tailless Operation 

Research (VECTOR), 705–7

Vehicle Integration Branch, 220

Vehicle Systems program, 765

Vehicle Technology Directorate (VTD), 

162–63

Venona signals intelligence decryption program, 17n30

Vermont wind power program, 826, 828, 

831

Vertical Motion Simulator (VMS), 151, 167

Vertical/Short Take-Off and Landing (V/

STOL) aircraft, 25, 60, 146

Vertical Take-Off and Landing (VTOL) aircraft: Dornier Do-31 aircraft, 648–49; 

research activities, 148–49; Short S.C.1 

aircraft, 639–41; VAK-191B aircraft, 649; 

VJ-101fighter, 648

Vertical Take-Off and Landing (VTOL) 

Approach and Landing Technology (VALT) 

program, 151

Vertol Aircraft Corporation, 136

Vertol CH-47 Chinook, 151, 162, 163; CH47C Chinook, 152; CH-47D Chinook, 

509–10

Vertol Division, Boeing Company, 136

Vertol YHC-1A, 148

Vestas wind turbine, 822, 848

Vibration testing, 328–29

Vickers Swallow aircraft, 60–61

Vickers VC10K tanker, 643

Vickers VC10 tranport, 642–43, 677

Vickers/Vickers-Supermarine: F-86 intelligence collection, 24; variable-geometry 

wing research, 60–61

Vietnam war, 289, 632, 644–45, 644n32

View factors: RAVFAC program, 502; VIEW 

program, 502

Virgin Atlantic, 619

Virginia-class submarines, 723

Virginia Polytechnic Institute, 493
NASA’s Contributions to Aeronautics

954

Virginia Tech, 494

Visco-Elastic and Hyperelastic Structures 

(VISCEL), 503

Viterna, Larry A., 846

VJ-101fighter, 648

VLOADS launch loads and dynamics program, 513–14

Voice-Controlled Interactive Device (VCID)/

Voice Command System, 672–73, 674, 675

Voigt, Waldemar, 14, 48n94, 58–59

Völkenrode, 14

Volta Congress on High Speeds in Aviation, 

5–7, 15n27, 93

Volvo, 706

Volvo RM-1 engine, 706

Von Braun, Wernher, 286, 512, 636

Von Kármán, Josephine “Pipa,” 17n30

Von Kármán, Theodore, 6, 14, 17n30, 216, 

291, 502

Von Kármán Institute, 244

Von Neumann, John, 284

Vortex flow control (VFC), 698

Vortex generators, 65

Vortex lift, 651, 651n49

Vortex ring state (VRS), 168

Vought, 480

Vought F4U-5 Corsair, 401

Vought F-8A fighter, 112–13

Vought F8U Crusader series. See also Digital Fly-By-Wire F-8 program and aircraft: area rule and redesign of, 103–4, 

105; digital flight control system, 380–

81, 753; F-8C, 380–81, 654–55, 753; F-8 

Supercritical wing, 89; F-8 Super Cruise 

Wing project, 655; F8U-1, 32, 37, 43; F8U3, 194; XF8U-1, 37, 363

Vought Missiles and Space Company, 

576–77

Vought nose cap, 306

Vought-Sikorsky Aircraft Division, United 

Aircraft Corporation, 8–9

Vought XF8U-1 Crusader, 37, 363

Voyager aircraft, 121, 121n106, 576, 578

VS-300 helicopter, 9, 135, 136, 142

VTOL Approach and Landing Technology 

(VALT) program, 151

Vulcan bomber, 45, 639

W

Walkden, Frank, 187, 196

Walker, Harry, 695

Walker, Joseph A. “Joe,” 42, 294, 298, 412

Wall, Steven D., 475

Wallis, Barnes, 60–61, 61n115

Wallops Flight Center/Wallops Island Station: Aeronautical Research Radar Complex, 151; instrument flight rules (IFR) 

investigation, 151; noise research, 165; 

Rotor Systems Research Aircraft (RSRA) 

program, 154–55; sonic boom research, 

194, 198, 206, 239; transonic research, 92

Walton, William C., Jr., 508

Ward, G.N., 94

Washington, DC, 217, 427

Washington Dulles International Airport, 

166, 635n15

Waspalloy, 592

Wasp-waist fuselage design, 89, 99, 107

Wasserfall (Waterfall) missiles, 638n15

Waterman, Waldo, 5

Water-wall cooling, 307, 589–90

WC-135B Stratotanker, 204

Weapon systems, 22

Weber, Richard, 331–32

Wedge-shaped vertical tail, 283

Wegener, Steven, 851, 861

Weidman, Deene J., 536

Weinberger, Caspar, 336, 338
Index 

955

Weinstein, Leonard M., 239

Welch, George “Wheaties,” 41, 42, 393

Westenhoff and Novick, 538

Western USA Sonic Boom Survey, 240–41, 

242

Westinghouse, 538

Westinghouse Electric Corporation, 829, 

834, 842

Westland Pterodactyls, 5

West Virginia University, 719

Whitcomb, Kenneth F., 90

Whitcomb, Richard Travis “Dick”: aerodynamic refinement and efficiency 

improvements, 89–90, 128–29; area rule, 

announcement of, 106–7; area rule, benefits of, 89–90, 126; area rule, development 

of, 51, 52, 95n14, 97–100; area rule, models, 88, 99; area rule, recognition for, 107–

8; area rule, transonic transport fuselage, 

127; area rule, validation in flight, 100–

106; character of, 115–16, 125–26; Collier Trophy, 52, 93–94, 107; contributions 

and legacy, 125–29; cut-and-try research 

technique, 113–14; Distinguished Service Medal, 107; early life of, 90–91; education of, 91; Exceptional Service Medal, 

107; influences on, 91; interference studies, 53, 65; intuitive reasoning and imagination, 95n14, 98, 114, 126; Medal for 

Exceptional Scientific Achievement, 116; 

National Medal of Science, 116; reputation of, 94; resignation of, 125–26; SCAT4, 67; SST program, 65, 67; supercritical 

wing, aviation industry use of, 114–15; 

supercritical wing, benefits of, 89–90, 126, 

805–6, 817; supercritical wing, development of, 805; supercritical wing, invention of, 68, 109–12; supercritical wing, 

models of, 110–11; supercritical wing, 

naming of, 111; supercritical wing, patent for, 114; supercritical wing, publication of findings, 111, 113; supercritical 

wing, recognition for, 116; supercritical wing, testing of, 112–14; Supersonic 

Transport study, 108–9, 108n61; Sylvanus 

Albert Reed Award, 116; thinking sessions, 

98; Transonic Aerodynamic Branch, 108, 

125; transonic drag research, 94–100; transonic transport wing configuration, 52, 

52n100, 68; visualization capabilities, 98; 

winglets, announcement of, 120; winglets, aviation industry use of, 120–22, 

806; winglets, benefits of, 89–90, 119–20, 

123–25, 126, 806, 817; winglets, development of, 68, 117–19; winglets, flight-testing of, 122–23; winglets, models of, 118, 

119; winglets, placement and design of, 

119; winglets, publication of findings, 118; 

winglets, wind tunnel studies of, 118, 119; 

Wright Brothers Memorial Trophy, 116

hite, Robert, 294

hite, Thomas D., 285

hitehead, Robert, 340

hite House: Executive Order to support 

aeronautics research, 171–72; National 

Sonic Boom Evaluation Program, 203–4; 

Presidential Advisory Committee (PAC) 

on SST, 192, 210–11; Science and Technology, Office of, 203, 230–31; supersonic 

research, 219

hite Sands Missile Range, 200, 238–39

hitman, Gerald B., 185

hitten, James, 291

ierzbanowski, Theodore Ted,” 695

ieselsberger, Carl, 280

illiams, Robert “Bob,” 337, 338, 340

illiamsburg, Virginia, 233, 537

illiams International, 760, 762, 797

W

W

W

W

W

W

W

W

W

W

W

W
NASA’s Contributions to Aeronautics

956

Windcharger Corporation, 826

Wind energy: cost of turbines, 843–44; Federal Wind Energy Program, 823–25, 827, 

828–29, 831, 833, 840, 843, 846–49; history of wind turbines, 825–28; horizontal-axis wind turbine (HAWT) generators, 

823, 825; manufacturers of turbines, 848; 

Mod-0A wind turbine, 831, 833–35; Mod-0 

wind turbine testbed, 827, 830, 831–33, 

846; Mod-1A wind turbine, 836; Mod-1 

wind turbine, 831, 835–36, 845; Mod-2 

wind turbine, 830, 831, 836–39, 840, 843, 

847, 848; Mod-3 wind turbine, 840; Mod-4 

wind turbine, 840; Mod-5A wind turbine, 

840, 841; Mod-5B wind turbine, 840, 841–

44, 847; Mod-5 wind turbine, 831, 840; 

Mod-6 wind turbine, 840; Nibe A wind 

turbine, 848; objectives of program, 828–

29; RANN symposium, 827–28; research 

activities at Lewis/Glenn, 844–49; rotor 

blade research, 844–46, 847; system verification units (SVU), 839; Vestas wind turbine, 822, 848; wind-electric dynamo, 825; 

Wind Energy Project Office, 830; Wind 

Power Office, 829–30

Windows, 307, 590–91

Wind tunnel and shock tunnel research: 

aerodynamic predictions, 397–98; conical camber, 51; DM-1 glider research, 47; 

hingeless rotor concept, 150; pitch-up, 25; 

rotary wing aircraft and helicopters, 140, 

143, 144; SCAT-15F, 67; sonic booms, 208–

9, 208n66, 214, 222, 235; Space Shuttle 

studies, 323–30, 327; speed limits, 573; 

winglets, 118, 119; X-48B, 71

Wing aspect ratio, 119–20, 407

Wing box research, 618–19

Wing dropping, 21

Wing fences, 414

Wing flaperons, 694, 697

Wing flaps, 414, 526, 595

Wing-flow model, 92

Winglets: announcement of, 120; aviation industry use of, 120–22, 806; benefits of, 89–90, 119–20, 123–25, 126, 806, 

817; blended winglets, 124; concept of, 

117; development of, 65, 68, 117–19, 786; 

flight-testing, 122–23; models, 118, 119; 

placement and design of, 119; publication of findings, 118; wind tunnel studies, 118, 119

Wingspan-to-fuselage length ratio, 3–4, 37, 

40n74, 407

Wingtip stall, 414

Wingtip vortexes, 117, 118, 119

Wing twist, 414

Wing warping, 386, 633

Wisneski, Jim, 701

Wlezien, Richard W., 244, 250–51, 254

Wolfe, Alex V., 673

Wolverine Ventures, Inc., 761

Women mathematicians, 465, 494

Wood, Drury W., Jr., 649

Woods, Robert J., 59, 290

Worcester Polytechnic Institute, 91, 107–8

World War I, 17, 70, 387, 463, 633

World War II: aluminum use, 571; flight 

control systems, 631, 633–34, 633n1, 635–

37; Link Trainer, 295, 392; piston engines, 

problems with, 736; rotary wing aircraft 

use, 142; V-2 missile shock waves, 183–84

Wright, Orville, 107, 364, 385, 386, 632–33

Wright, Ray, 95

Wright, Wilbur, 364, 385, 386, 632–33

Wright Aeronautical ramjets, 289

Wright Air Development Center, Flight Control Laboratory, 639

Wright Brothers Memorial Trophy, 116
Index 

957

Wright Field Aircraft Laboratory, 9–10

Wright Flyer, 37, 364, 385, 391, 633, 752

Wright Laboratory, 541

Wright-Patterson Air Force Base: Aeronautical Research Laboratory, 186; Air 

Force Museum, 648, 679, 709; Flight 

Dynamics Laboratory, 592; Matrix Methods in Structural Mechanics conference, 

474–75; NASTRAN development, 474; 

NASTRAN Users’ Colloquia, 540, 541; 

Numerical Propulsion System Simulation (NPSS) program, 760

Wright R-1820 Cyclone engine, 736

Wright R-3350 Duplex Cyclone engine, 736

Wyle Laboratories, 240, 242, 246

Wyle Research, 226

X

X-1 aircraft: Collier Trophy, 93; development and manufacture of, 108, 290, 494; 

engines for, 294; flight control systems, 

362–63, 752; gliding landings, 402–3; 

lift-to-drag ratio (L/D), 403; Round One 

research aircraft, 55n106; supersonic 

capabilities, 94; tail configuration, 411; 

X-1A aircraft, 40–41, 40n74, 43, 292, 363, 

407; X-1B aircraft, 40n74, 298; X-1D aircraft, 40n74; Yeager’s flight, 40–41, 292, 

363, 407

X-2 aircraft: accident with, 44n87, 394–95, 

407, 410; control coupling, 410; design 

of, 34; development and manufacture of, 

17n30, 290; dihedral effects, 410; flight 

control systems, 298; inertial coupling, 

44n87; pressure suit for, 297; simulator, 

394–95; tail configuration, 411

X2 helicopter, 158

X-3 “Stiletto,” 3, 4, 37, 40, 41–42, 43, 440

X-4 aircraft, 27–28

X-5 aircraft, 17n30, 58, 59–60, 413

X-7 testbed, 278, 289, 290, 572

X-15 aircraft: ablation cooling research, 

417–18; accident with, 299, 373–74; centrifuge program, 399–400; contracts for, 293; 

contributions from program, 299; design 

confidence, 307, 308; development of, 

278, 283, 290–93; at Dryden, 294; Dutch 

roll motions, 411–12; emergencies, practice for, 296; flight control system, 298–

99, 372–74, 376, 379, 642; flight-testing, 

294–95; flight time, 299; funding for, 291; 

heating, structural, 414–15; heat-sink 

structures, 415–17; inertial coupling (roll 

coupling), 409; instrumentation development, 298; landing of, 405; landings of, 

403–4; lift-to-drag ratio (L/D), 403; limitcycle oscillation research, 368; materials for, 292, 293, 295, 571, 592; pressure 

suits, 276, 278, 297, 298; reaction controls, 

298; record hypersonic flight, 278, 293, 

299, 642; redundant systems, 365–66; roll 

damper, 412; Round Two research aircraft, 

54, 55, 290; side stick flight controller, 642, 

658n59; simulator, 295–97, 395–97, 403–

4, 405, 409, 411–12; sonic boom research, 

197, 226; speed of, 415, 642; Stability Augmentation Systems, 365–66, 399; stability 

issues, 292, 411–12; structural resonances, 

375–76; tail configuration, 278, 411; thermal barrier coatings, 611; thermal damage, 294–95, 416, 418; thermal data from, 

517–18; X-15A-2 aircraft, 64, 335, 406, 

417–18

X-17 missile, 287–88

X-20 Dyna-Soar: cancellation of, 304, 308, 

400, 592; centrifuge program, 400; contributions from program, 307–8; design confidence, 307–8; development of, 300–305; 
NASA’s Contributions to Aeronautics

958

flight control system, 372; flutter-testing 

panels, 585–86; hot structures, 305–8, 

418, 582–92; justification of, 302–5; landing skids, 592; materials for, 305–8; missiles to launch, 301–2; nose cap, 306, 308, 

587–89, 598; Round Three research aircraft, 55–56; Round Two research aircraft, 

55n106; schematic drawing, 583; simulator, 405–6; windows, 307

X-23 (SV-5D) PRIME (Precision Recovery 

Including Maneuvering Entry), 27n51, 

278, 288, 304, 417, 418, 606–8

X-24A (SV-5P) lifting body, 27n51, 407, 607

X-24B lifting body, 56, 404, 407

X-24C hypersonic lifting body proposal, 

299, 448–49

X-27 aircraft, 34n61

X-29 aircraft: aeroelastic effects, 496, 528–

29, 694; aeroelastic tailoring, 694, 695; 

angle of attack capabilities, 694, 695, 

696–98; benefits of, 697–98; contract for, 

692–93; control laws, 696; control surfaces, 694–95; difference between two 

X-29s, 693–94; exhibit showing, 698–99; 

finite element analysis, 460, 515, 526–29, 

530; flight-testing, 695–98; fly-by-wire 

system, 383, 692–98, 753; forward-swept 

wing design, 22, 527, 693–94; NASTRAN 

model, 460; redundant systems, 695; sensor selection thresholds, 698; strain gages, 

528; supercritical wing, 693, 694, 695, 

695n126; vortex flow control, 698; X-29A 

aircraft, 460

X-30 National Aero-Space Plane, 230–31, 

241, 279, 336–40, 342, 419, 615–21

X-31 aircraft: accident with, 704–5; air combat simulation, 703; angle of attack, 701–2, 

706–7; autothrottle system, 706; Enhanced 

Fighter Maneuverability, 699–705; exhibit 

showing, 707; flight-testing, 701–7; Flush 

Air Data System, 707; fly-by-wire system, 700–705; Kiel probe, 701, 704, 705, 

705n149; poststall maneuvering, 702–3, 

703n144; tail configuration, 703–4; thrust 

vectoring, 700, 701, 703–4, 705–7; thrust 

vectoring paddles, 699, 700, 706; VECTOR 

program, 705–7

X-31 Enhanced Fighter Maneuverability 

(EFM) aircraft, 699–707

X-33 Research Vehicle, 384

X-38 aircraft, 27n51

X-43 Hyper-X vehicle, 279, 341, 343–45, 

344n127, 568

X-48B Blended Wing-Body UAV, 2, 71–72

X-51A vehicle, 279, 345–46

XB-52 aircraft, 20

XB-70 Valkyrie: aeroelastic effects research, 

392; SCAT program and, 108; sonic boom 

research, 201–2, 201n49, 203, 204, 205, 

210; XB-70-1 testbed, 201, 202, 204; 

XB-70-2 testbed, 201–2, 204; XB-70A, 

56n107, 191

XCH-62 Heavy Lift Helicopter (HLH) program, 158

XF2Y-1 Sea Dart, 45, 103

XF4D-1 Skyray, 32–33

XF8U-1 Crusader, 37, 363

XF10F-1 Jaguar, 59, 60

XF-88 aircraft, 33

XF-92A aircraft, 45, 46n90, 49–51, 100

XF-103 interceptor fighter, 289, 290, 333

XF-104 Starfighter, 185

XH-51N helicopter, 150

XH-59A Advancing Blade Concept (ABC) 

helicopter, 152–53, 157–58

XP-86 Sabre, 18, 635

XP-92 aircraft, 46

XS-1 aircraft: design of, 3, 28, 34, 34n62, 36, 
Index 

959

184; falling body test, 38, 43; flight control 

systems, 635; inspiration for, 14n22; sonic 

boom from, 182, 182n3; supersonic flight 

by, 182, 183; testing of, 35; variable-geometry wing tests, 59; wingspan-to-fuselage 

length ratio, 37; XS-1 FSW planform, 21; 

Yeager’s flight, 182, 182n3, 283, 635

XS-2 aircraft, 34, 34n62

XV-1 Convertiplane, 134, 147

XV-15 Tilt-Rotor Research Aircraft (TRRA), 

155, 156, 160–62, 165, 166

X-Wing Stoppable Rotor, 157, 158

X-Wing Stoppable Rotor program, 157, 158, 

163–64, 165

Y

Yaw, 19, 407–8, 410

Yaw, adverse, 19–20, 409

Yaw damper, 637

YB-49 flying wing jet bomber, 19, 637

YC-14 transport, 114–15, 670–72

YC-15 aircraft, 114, 379–80

Yeager, Charles E. “Chuck”: Collier Trophy, 

93; flight control systems, 752; supersonic 

aircraft research, 94; X-1A flight, 40–41, 

292, 363, 407; XS-1 flight, 182, 182n3, 

283, 635

YF-4E Phantom Survivable Flight Control 

System program, 379–80, 384, 646–48, 

658n59, 660

YF-12 Blackbird: digital engine control 

system, 754; Loads Alleviation and Mode 

Suppression (LAMS) system, 392; sonic 

boom research, 202–3, 204, 210, 227, 

227n133; thermal loads project, 515, 516–

19, 529, 554; YF-12A, 202–23; YF-12C, 

516, 708–9; YF-12 Experiment Symposium, 519

YF-16 Lightweight Fighter (LWF), 371; 

angle of attack limiters, 665–66; flighttesting, 667–68; flutter suppression 

research, 388, 389; fly-by-wire system, 

379–80, 647, 664–69; pilot-induced oscillations, 370; roll stick force gradient, 667; 

side stick flight controller, 658, 664, 665, 

668; simulator studies, 666–67; strain 

gages, 385–86; structural resonance 

research, 378; success of, 668; unintentional first flight, 667

YF-17 Lightweight Fighter (LWF), 388, 389, 

664, 669

YF-22 aircraft, 402

YF-86D low-tail testbed, 36–37

YF-100A Super Sabre, 34, 363

YF-100 Super Sabre, 37, 40, 41, 42, 185

YF-102 Delta Dagger: area rule and, 51, 89, 

100–102, 103; handling qualities, 53; sonic 

booms, 185; YF-102A prototype, 89, 101–

2, 103

YG-2 autogiro, 141

YO-3A aircraft, 242–43

York, Herbert, 301

YR-4B helicopter, 142, 143–44

Yttria, 587

Yttria-stabilized zirconia, 610, 611–12, 613

Z

Zahm, Albert, 11

Zero CO2 project, 813

Zhu, Dong-ming, 613

Zirconia: magnesia-stabilized zirconia, 611; 

nose cones/caps, 587, 588, 589, 594, 595, 

598; reusable surface installation (RSI), 

601; yttria-stabilized zirconia, 610, 611–

12, 613

Zirconium, 574, 584, 588, 614

Zirconium Corporation of America, 588

Zirconium oxide, 306, 308
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