The NASA History Series

National Aeronautics and Space Administration

NASA History Division

Office of External Relations

Washington, D.C.

2007

Thomas A. Heppenheimer has been a freelance 

writer since 1978. He has written extensively on 

aerospace, business and government, and the 

history of technology. He has been a frequent 

contributor to American Heritage and its affiliated 

publications, and to Air & Space Smithsonian. 

He has also written for the National Academy of 

Sciences, and contributed regularly to Mosaic of the 

National Science Foundation. He has written some 

300 published articles for more than two dozen 

publications.

He has also written twelve hardcover books. 

Three of them–Colonies in Space (1977), Toward 

Distant Suns (1979) and The Man-Made Sun

(1984)-have been alternate selections of the 

Book-of-the-Month Club. His Turbulent Skies

(1995), a history of commercial aviation, is part 

of the Technology Book Series of the Alfred P. 

Sloan Foundation. It also has been produced as 

a four-part, four-hour Public Broadcasting System 

television series Chasing the Sun.

Under contract to NASA, Heppenheimer has written 

that agency’s authorized history of the Space 

Shuttle, in two volumes. Volume 1, The Space 

Shuttle Decision (1999), has been reissued in 

paperback by the Smithsonian Institution Press and 

has been selected as an Outstanding Academic 

Title. The present book reflects his longstanding 

activity in hypersonics, for which he has written 

three technical reviews for Pasha Publications.

He holds a Ph.D. in aerospace engineering from 

the University of Michigan, and is an associate 

fellow of the American Institute of Aeronautics and 

Astronautics. He has also held research fellowships 

in planetary science at the California Institute of 

Technology and at the Max Planck Institute in 

Heidelberg, Germany.

About the Cover: Hypersonic Plane by Leslie 

Bossinas. Artist’s concept of an aerospace plane 

showing aero-thermal heating effects caused by 

friction as the vehicle flies hypersonically through 

the atmosphere. The National Aero-Space Plane 

program provided technology for space launch 

vehicles and hypersonic cruise vehicles. This 

vehicle with advanced airbreathing engines would 

have the capability to take off horizontally from 

and land on conventional runways, accelerate to 

orbit, and cruise hypersonically in the atmosphere 

between Earth destinations. (NASA Art Program, 

Image 86-HC-217).

A History of Hypersonics

Facing the Heat Barrier:

T. A. Heppenheimer

NASA SP-2007-4232
T. A. Heppenheimer

Facing the Heat Barrier:

A History of Hypersonics
The NASA History Series

National Aeronautics and Space Administration

NASA History Division

Office of External Relations

Washington, DC

September 2007

NASA SP-2007-4232

Facing the Heat Barrier:

A History of Hypersonics

T. A. Heppenheimer
Library of Congress Cataloging-in-Publication Data

Heppenheimer, T. A., 1947-

 Facing the heat barrier : a history of hypersonics / T.A. Heppenheimer.

 p. cm.—NASA History Series (NASA SP-2007-4232)

 Includes bibliographical references.

 1. Aerodynamics, Hypersonic--Research--United States--History--20th century. 2. 

Research aircraft--United States--History--20th century. I. Title. III. NASA SP-2007-4232.

TL571.5.H47 2006

629.132’3060973--dc22

 2006030654
Table of Contents

Acknowledgements. . 

Introduction. .

Abbreviations and Acronyms. .

Chapter 1: First Steps in Hypersonic Research. 

German Work with High-Speed Flows. 

Eugen Sänger. 

NACA-Langley and John Becker. .

Chapter 2: Nose Cones and Re-entry. .

The Move Toward Missiles. .

Approaching the Nose Cone. .

Ablation. .

Flight Test. .

Chapter 3: The X-15. .

Origins of the X-15. .

The Air Force and High-Speed Flight. .

X-15: The Technology. .

X-15: Some Results. .

Chapter 4: First Thoughts of Hypersonic Propulsion. .

Ramjets As Military Engines. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 91

Origins of the Scramjet. .

Combined-Cycle Propulsion Systems. . 

Aerospaceplane. . 

Chapter 5: Widening Prospects for Re-entry. . 

Winged Spacecraft and Dyna-Soar. . 

The Technology of Dyna-Soar. . 

Heat Shields for Mercury and Corona. . 

Gemini and Apollo. . 

Chapter 6: Hypersonics and the Space Shuttle. . 

Preludes: Asset and Lifting Bodies. . 

Asset Flight Tests. . 

Reusable Surface Insulation. . 

Designing the Shuttle. . 

The Loss of Columbia. . 

Chapter 7: The Fading, the Comeback. . 

Scramjets Pass Their Peak. . 

Scramjets at NASA-Langley. . 

The Advent of NASP. . 

The Decline of NASP. . 
Acknowledgements

It is a pleasure to note the numerous people who helped me with this book. My 

personal involvement in hypersonics dates to 1982. I wrote a number of free-lance 

articles, along with three book-length reviews, before beginning work on the present 

book in 1999. During these past two decades, several dozen people kindly granted 

interviews in the course of these assignments. This book draws on discussions with 

the following:

J. Leland Atwood, Robert Barthelemy, George Baum, Fred Billig, Richard 

Booton, Peter Bradshaw, William Cabot, Robert Cooper, Scott Crossfield, 

Paul Czysz, William Dannevik, Anthony duPont, James Eastham, John 

Erdos, Maxime Faget, George Gleghorn, Edward Hall, Lawrence Huebner, 

Antony Jameson, Robert Jones, Arthur Kantrowitz, James Keller, George 

Keyworth, William “Pete” Knight, John Lumley, Luigi Martinelli, Robert 

Mercier, Parviz Moin, Gerhard Neumann, Louis Nucci, Philip Parrish, 

John Pike, Heinz Pitsch, Jerry Rising, Anatol Roshko, Paul Rubbert, Ron 

Samborsky, Robert Sanator, George Schairer, David Scott, Christian Stemmer, Arthur Thomas, Steven Weinberg, and Robert Williams.

In the NASA History Division, NASA Chief Historian Steven Dick served effectively as my editor. NASA-Langley has an excellent library, where I received particular help from Sue Miller and Garland Gouger. In addition, Dill Hunley, the 

historian at NASA-Dryden, hosted me for a week-long visit. The archivist, Archie 

DiFante, gave similar strong support during my visits to Maxwell Air Force Base. 

The Science and Technology Corporation, administered my work under subcontract, for which I give thanks to Andrea Carden, Carla Coombs, Sue Crotts, Marion 

Kidwell, and George Wood.

Dennis Jenkins provided me with documents and answered a number of questions. The artists Don Dixon and Chris Butler, who helped me on previous book 

projects, provided valuable assistance on this one as well. In addition, as for previous 

books, Phyllis LaVietes served as my secretary.

This book reflects my interest in the National Aerospace Plane effort, which I 

covered as a writer beginning in 1985. It is a pleasure to recall my ongoing friendships with Robert Williams, who gave me access to sources; Fred Billig, who helped 

me learn the trade of hypersonics; and Arthur Kantrowitz, who was present at the 

beginning. These three stand out among the dozens of people with whom it has 

been my privilege to conduct interviews and discussions.

vii

Chapter 8: Why NASP Fell Short. . 

Aerodynamics. . 

Propulsion. . 

Materials. . 

Chapter 9: Hypersonics After NASP. . 

The X-33 and X-34. . 

Scramjets Take Flight. . 

Recent Advances in Fluid Mechanics. . 

Hypersonics and the Aviation Frontier. . 

Bibliography. . 

NASA History Series. . 

Index. . 

ix

Introduction

As an approach to the concept of hypersonic flight, one may begin by thinking of 

a sequence of high-performing aircraft that have flown at successively higher speeds. 

At Mach 2, twice the speed of sound, typical examples included the F-104 fighter 

and the Concorde commercial airliner. Though dramatically rakish in appearance, 

they were built of aluminum, the most familiar of materials, and used afterburning 

turbojets for propulsion.1

At Mach 3 and higher, there was the Lockheed SR-71 that cruised at 85,000 

feet. The atmosphere at such altitudes, three times higher than Mount Everest, has 

a pressure only one-fiftieth of that at sea level. Even so, this airplane experienced 

aerodynamic heating that brought temperatures above 500°F over most of its surface. In turn, this heating brought requirements that dominated the problems of 

engineering design. Aluminum was out as a structural material; it lost strength at 

that high temperature. Titanium had to be used instead. Temperature-resistant fuels 

and lubricants also became necessary. Even so, this aircraft continued to rely on 

afterburning turbojets for propulsion.2

At Mach 4, the heating became still more severe and the difficulties of design 

were more daunting. No version of the turbojet has served at such speeds; it has 

been necessary to use a ramjet or rocket. The X-7, a ramjet testbed craft of the 

1950s, was built of steel and had better temperature resistance than the SR-71. Still, 

when it flew past Mach 4.3 in 1958, the heating became so severe that it produced 

structural failure and a breakup of the vehicle in flight.3

Yet Mach 4 still counts as merely supersonic flight, not as hypersonic. For more 

than half a century analysts have defined hypersonic speeds as Mach 5 and higher.4

Only rocket-powered craft have flown so fast—and Mach 5 defines only the lower 

bound of the hypersonic regime. An important range of hypersonic speeds extends 

from Mach 20 to 25 and includes the velocities of long-range ballistic missiles and 

of satellites re-entering from orbit. Moreover, flight above Mach 35 was a matter 

of national concern during the Apollo program, for its piloted Command Module 

entered the atmosphere at such speeds when returning from the Moon.

Specifically, the hypersonic regime is defined as the realm of speed wherein the 

physics of flows is dominated by aerodynamic heating. This heating is far more 

intense than at speeds that are merely supersonic, even though these lesser velocities 

have defined the performance of the SR-71 and X-7.

Hypersonics nevertheless was a matter of practical military application before 

the term entered use. Germany’s wartime V-2 rocket flew above Mach 5,5

 but steel 

proved suitable for its construction and aerodynamic heating played only a limited 
xi

Introduction

Corona program in strategic reconnaissance, Apollo, and the space shuttle. These 

activities deterred nuclear war, gained accurate estimates of the Soviet threat, sent 

astronauts to the Moon and brought them home, and flew to and from space in a 

reusable launch vehicle. This list covers many of the main activities of the postwar 

missile and space industry, and supports Hallion’s viewpoint.

But in pursuing technical revolution, engineers succeed in actually solving their 

problems, as when the Apollo program sent men to the Moon. These people do not 

merely display brilliant ingenuity while falling short of success. Unfortunately, the 

latter has been the case in the important area of hypersonic propulsion.

The focus has involved the scramjet as a new engine. It has taken form as a prime 

mover in its own right, capable of standing alongside such engines as the turboprop 

and ramjet. Still, far more so than the other engines, the scramjet has remained in 

the realm of experiment. Turboprops powered the Lockheed Electra airliner, P-3 

antisubmarine aircraft, and C-130 transport. Ramjets provided propulsion for the 

successful Bomarc and Talos antiaircraft missiles. But the scramjet has powered only 

such small experimental airplanes as the X-43A.

Why? From the outset, the scramjet has faced overwhelming competition from 

a successful alternative: the rocket. This has strongly inhibited funding and has 

delayed its development to a point at which it could be considered seriously. On 

paper, scramjets offer superior performance. They therefore drew attention in the 

mid-1980s, during the heyday of NASP, at a time when Air Force officials had 

become disenchanted with the space shuttle but faced huge prospective demand for 

access to space in President Reagan’s Strategic Defense Initiative. For once, then, 

scramjets gained funding that served to push their development—and their performance fell well short of people’s hopes.

Within this book, Chapter 1 covers the immediate postwar years, when America 

still had much to learn from the Europeans. It focuses on two individuals: Eugen 

Sänger, who gave the first proposal for a hypersonic bomber, and John Becker, who 

built America’s first hypersonic wind tunnel.

Chapter 2 covers the first important area of hypersonic research and development, which supported the advent of strategic missiles during the 1950s. The focus 

was on solving the re-entry problem, and this chapter follows the story through 

flight tests of complete nose cones.

Chapter 3 deals with the X-15, which took shape at a time when virtually the 

whole of America’s capability in hypersonics research was contained within Becker’s 

11-inch instrument. Today it is hard to believe that so bold and so successful a step 

in aviation research could stand on so slender a foundation. This chapter shows how 

it happened.

Chapter 4 introduces hypersonic propulsion and emphasizes the work of Antonio Ferri, an Italian aerodynamicist who was the first to give a credible concept for a 

scramjet engine. This chapter also surveys Aerospaceplane, a little-known program of 

role in its overall design.6

 The Germans used wind-tunnel tests to ensure that this 

missile would remain stable in flight, but they did not view its speed regime as 

meriting a name of its own. Hsue-shen Tsien, an aerodynamicist at the California 

Institute of Technology, coined the term in 1946.7

 Since then, it has involved three 

significant areas of application.

The first was the re-entry problem, which came to the forefront during the mid1950s. The Air Force by then was committed to developing the Atlas ICBM, which 

was to carry a nuclear warhead to Moscow. Left to itself, this warhead would have 

heated up like a meteor when it fell back into the atmosphere. It would not have 

burned up—it was too massive—but it certainly would have been rendered useless. Hence, it was necessary to devise a heat shield to protect it against this intense 

aerodynamic heating.

The successful solution to this problem opened the door to a host of other initiatives. The return of film-carrying capsules from orbit became routine, and turned 

strategic reconnaissance of the Soviet Union into an important element of national 

defense. Piloted space flight also became feasible, for astronauts now could hope to 

come back safely. Then, as the engineering methods for thermal protection were 

further improved, thoughts of a space shuttle began to flourish. They took shape as 

a reusable launch vehicle, the first of its kind.

Hypersonic technologies also became important as policy makers looked ahead to 

an era in which the speed and performance of fighters and bombers might increase 

without limit. This expectation led to the X-15. Though designed during the 1950s, 

this rocket-powered research airplane set speed and altitude marks that were not surpassed until the advent of the shuttle. Aerodynamic heating again defined its design 

requirements, and it was built of the nickel alloy Inconel X. It routinely withstood 

temperatures of 1200°F as it flew to Mach 6,8

 and reached altitudes high enough for 

some of its pilots to qualify as astronauts.

Only rocket engines could propel a vehicle at such speeds, but hypersonic propulsion has represented a third important area of application. Here the hope has 

persisted that innovative airbreathing engines—scramjets—might cope with intense 

aerodynamic heating while offering fuel economy far surpassing that of a rocket. 

Other work has emphasized airbreathing rockets, which could give improved performance by eliminating the need to carry liquid oxygen in a tank. These concepts have 

held their own importance. They lay behind the National Aerospace Plane (NASP) 

program of 1985-1995, which sought to lay groundwork for single-stage vehicles 

that were to use both types of engine and were to fly from a runway to orbit.

The Air Force historian Richard Hallion has written of a “hypersonic revolution,” as if to place the pertinent technologies on par with the turbojet and liquidpropellant rocket.9

 The present book takes a more measured view. Work in hypersonics had indeed brought full success in the area of re-entry. Consequences have 

included strategic missiles, the Soviet and American man-in-space programs, the 
xiii

Introduction

1 F-104: Gunston, Fighters, pp. 120-126. Concorde: Heppenheimer, Turbulent, pp. 202-203, 

208.

2 Crickmore, SR-71, pp. 89-91, 95-99, 194.

3 Ritchie, “Evaluation.”Steel: Miller, X-Planes, p. 119.

4 See, for example, Anderson, History, pp. 438-439.

5 Top speed of the V-2 is given as 1,600 meters per second (Dornberger, V-2, p. xix) and as 1,700 

meters per second (Naval Research Laboratory, Upper, cited in Ley, Rockets, pp. 596-597); the 

speed of sound at the pertinent altitudes is 295 meters per second (Kuethe and Chow, Foundations, p. 518).

6 Ley, Rockets, p. 243; Neufeld, Rocket, pp. 85-94.

7 Tsien, “Similarity.”

8 1200oF: NASA SP-2000-4518, diagram, p. 25.

9 Hallion, Hypersonic.

paper studies that investigated the feasibility of flight to orbit using such engines.

The next two chapters cover important developments in re-entry that followed 

the ICBM. Chapter 5, “Widening Prospects for Re-Entry,” shows how work in this 

area supported the manned space program while failing to offer a rationale for a 

winged spacecraft, Dyna-Soar. Chapter 6, “Hypersonics and the Shuttle,” begins by 

outlining developments during the mid-1960s that made it plausible that NASA’s 

reusable space transporter would be designed as a lifting body and built using hot 

structures. In fact, the shuttle orbiter came forth as a conventional airplane with 

delta wings, and was built with aluminum structure covered with thermal-protecting tiles. This discussion indicates how those things happened.

Chapter 7, “The Fading, the Comeback,” shows how work with scramjets did 

not share the priority afforded to the topic of re-entry. Instead it faded, and by the 

late 1960s only NASA-Langley was still pursuing studies in this area. This ongoing 

effort nevertheless gave important background to the National Aerospace Plane—

but it was not technical success that won approval for NASP. As noted, it was the 

Strategic Defense Initiative. Within the Strategic Defense Initiative, the scramjet 

amounted to a rabbit being pulled from a hat, to satisfy Air Force needs. NASP was 

not well-founded at the outset; it was more of a leap of faith.

Chapter 8, “Why NASP Fell Short,” explains what happened. In summary, the 

estimated performance of its scramjet engine fell well below initial hopes, while the 

drag was higher than expected. Computational aerodynamics failed to give accurate 

estimates in critical technical areas. The ejector ramjet, a key element of the propulsion system, proved to lack the desired performance. In the area of materials, metallurgists scored an impressive success with a new type of titanium called Beta-21S. It 

had only half the density of the superalloys that had been slated for Dyna-Soar, but 

even greater weight savings would have been needed for NASP.

Finally, Chapter 9 discusses “Hypersonics After NASP.” Recent developments 

include the X-33 and X-34 launch vehicles, which represent continuing attempts 

to build the next launch vehicle. Scramjets have lately taken flight, not only as 

NASA’s X-43A but also in Russia and in Australia. In addition, the new topic of 

Large Eddy Simulation, in computational fluid mechanics, raises the prospect that 

analysts indeed may learn, at least on paper, just how good a scramjet may be.

What, in the end, can we conclude? During the past half-century, the field 

of hypersonics has seen three major initiatives: missile nose cones, the X-15, and 

NASP. Of these, only one—the X-15—reflected ongoing progress in aeronautics. 

The other two stemmed from advances in nuclear weaponry: the hydrogen bomb, 

which gave rise to the ICBM, and the prospect of an x-ray laser, which lay behind 

the Strategic Defense Initiative and therefore behind NASP.

This suggests that if hypersonics is to flourish anew, it will do so because of 

developments in the apparently unrelated field of nuclear technology.
Abbreviations and

Acronyms

AAF Army Air Forces

AAS American Astronautical Society

ACES Air Collection and Enrichment System

AEC Atomic Energy Commission

AEDC Arnold Engineering Development Center

AEV Aerothermodynamic Elastic Vehicle (ASSET)

AFB Air Force Base

AFSC Air Force Systems Command

AGARD Advisory Group for Aeronautical Research and Development

AIAA American Institute for Aeronautics and Astronautics

AIM Aerothermodynamic Integration Model (HRE)

APL Applied Physics Laboratory (Johns Hopkins University)

APU auxiliary power unit

ARDC Air Research and Development Command (USAF)

ARS American Rocket Society

ASD Aeronautical Systems Division (USAF)

ASSET Aerothermodynamic/elastic Structural Systems Environmental Tests

ASV Aerothermodynamic Structural Vehicle (ASSET)

BMW Bayerische Motoren-Werke

BTU British Thermal Unit

CASI Center for Aerospace Information

CD coefficient of drag

CDE Concept Demonstrator Engine (NASP)

CFD computational fluid dynamics

CFHT Continuous Flow Hypersonic Tunnel (NACA-Langley)

CIA Central Intelligence Agency

CIAM Central Institute of Aviation Motors (Moscow)

CL coefficient of lift

CO2 carbon dioxide

DARPA Defense Advanced Research Projects Agency

DM Deutschmark

DNS Direct Numerical Simulation (CFD)

DOD Department of Defense

DSB Defense Science Board

DTIC Defense Technical Information Center

DVL Deutsche Versuchsanstalt für Luftfahrtforschung

ELV expendable launch vehicle

oF degrees Fahrenheit

FAI Federation Aeronautique Internationale

FDL Flight Dynamics Laboratory (USAF)

FY Fiscal Year

g force of gravity
GAO General Accounting Office (United States Congress)

GASL General Applied Science Laboratories, Inc.

GE General Electric

HGS homogeneous gas sample (shock tubes)

HRE Hypersonic Research Engine

HXEM Hyper-X Engine Module

HXFE Hyper-X Flight Engine

HYWARDS Hypersonic Weapons Research and Development Supporting System

IBM International Business Machines

ICBM intercontinental ballistic missile

IFTV Incremental Flight Test Vehicle

ILRV Integrated Launch and Re-entry Vehicle

ISABE International Society for Air Breathing Propulsion I

sp specific impulse

ISTAR Integrated System Test of an Airbreathing Rocket K degrees Kelvin

LACE Liquid Air Cycle Engine

L/D lift-to-drag ratio

LES Large-Eddy Simulation (CFD)

LH

2 liquid hydrogen

LSCIR Low Speed Component Integration Rig (Pratt & Whitney)

LSS Low Speed System (NASP)

MBB Messerschmitt-Boelkow-Blohm

MIT Massachusetts Institute of Technology

MOL Manned Orbiting Laboratory

MOU Memorandum of Understanding

MSC Manned Spacecraft Center (now Johnson Space Center)

NACA National Advisory Committee for Aeronautics

NASA National Aeronautics and Space Administration

NASP National Aerospace Plane

NIFTA Non-Integral Fuselage Tank Article (NASP)

NMASAP NASP Materials and Structures Augmentation Program

OAL Ordnance Astrophysics Laboratory

OMSF Office of Manned Space Flight (NASA)

P & W Pratt & Whitney

POBATO propellants on board at takeoff

PRIME Precision Recovery Including Maneuvering Entry

PROFAC propulsive fluid accumulator

PSi pounds per square inch

RCC Reusable Carbon-Carbon

RENE Rocket Engine Nozzle Ejector

ROLS Recoverable Orbital Launch System

RSI reusable surface insulation

SAB Scientific Advisory Board (USAF)

SAGE Semi-Automatic Ground Environment

SAM Structures Assembly Model (HRE)

SAMPE Society for Advancement of Materials and Process Engineering

SCRAM Supersonic Combustion Ramjet Missile (APL)

SDI Strategic Defense Initiative

SSME Space Shuttle Main Engine

SSTO single stage to orbit

SXPE Subscale Parametric Engine (NASP)

TAV Trans-Atmospheric Vehicle

TPS thermal protection system

TZM titanium-zirconium-molybdenum alloy

USAF United States Air Force

VKF Von Karman Facility (AEDC)

WADC Wright Air Development Center (USAF)

Today’s world of high-speed flight is international, with important contributions having recently been made in Japan, Australia, and Russia as well as in the 

United States. This was even truer during World War II, when Adolf Hitler sponsored development programs that included early jet fighters and the V-2 missile. 

America had its own research center at NACA’s Langley Memorial Aeronautical 

Laboratory, but in important respects America was little more than an apt pupil of 

the wartime Germans. After the Nazis surrendered, the U.S. Army brought Wernher von Braun and his rocket team to this country, and other leading researchers 

found themselves welcome as well.

1

First Steps in

Hypersonic Research

1

Liftoff of a V-2 rocket. (U.S. Army)
2

Facing the Heat Barrier: A History of Hypersonics

3

First Steps in Hypersonic Research

After that, though, work in supersonics went forward with new emphasis. Jakob 

Ackeret, a colleague of Prandtl, took the lead in building supersonic wind tunnels. 

He was Swiss, and he built one at the famous Eidgenossische Technische Hochschule in Zurich. This attracted attention in nearby Italy, where the dictator Benito 

Mussolini was giving strong support to aviation. Ackeret became a consultant to the 

Italian Air Force and built a second wind tunnel in Guidonia, near Rome. It reached 

speeds approaching 2,500 miles per hour (mph), which far exceeded those that were 

available anywhere else in the world.2

These facilities were of the continuous-flow type. Like their subsonic counterparts, they ran at substantial power levels and could operate all day. At the Technische Hochschule in Aachen, the aerodynamicist Carl Wiesenberger took a different approach in 1934 by building an intermittent-flow facility that needed much 

less power. This “blowdown” installation relied on an evacuated sphere, which 

sucked outside air through a nozzle at speeds that reached Mach 3.3.

This wind tunnel was small, having a test-section diameter of only four inches. 

But it set the pace for the mainstream of Germany’s wartime supersonic research. 

Wieselberger’s assistant, Rudolf Hermann, went to Peenemunde, the center of that 

country’s rocket development, where in 1937 he became head of its new Aerodynamics Institute. There he built a pair of large supersonic tunnels, with 16-inch test 

sections, that followed Aachen’s blowdown principle. They reached Mach 4.4, but 

not immediately. A wind tunnel’s performance depends on its nozzle, and it took 

time to develop proper designs. Early in 1941 the highest working speed was Mach 

2.5; a nozzle for Mach 3.1 was still in development. The Mach 4.4 nozzles were not 

ready until 1942 or 1943.3

The Germans never developed a true capability in hypersonics, but they came 

close. The Mach 4.4 tunnels introduced equipment and methods of investigation 

that carried over to this higher-speed regime. The Peenemunde vacuum sphere was 

constructed of riveted steel and had a diameter of 40 feet. Its capacity of a thousand 

cubic meters gave run times of 20 seconds.4

 Humidity was a problem; at Aachen, 

Hermann had learned that moisture in the air could condense when the air cooled 

as it expanded through a supersonic nozzle, producing unwanted shock waves that 

altered the anticipated Mach number while introducing nonuniformities in the 

direction and velocity of flow. At Peenemunde he installed an air dryer that used 

silica gel to absorb the moisture in the air that was about to enter his supersonic 

tunnels.5

Configuration development was at the top of his agenda. To the modern mind 

the V-2 resembles a classic spaceship, complete with fins. It is more appropriate to 

say that spaceship designs resemble the V-2, for that missile was very much in the 

forefront during the postwar years, when science fiction was in its heyday.6

 The V-2 

needed fins to compensate for the limited effectiveness of its guidance, and their 

Some of their best work had supported the V-2, using a pair of tunnels that operated at Mach 4.4. This was just short of hypersonic, but these facilities made a key 

contribution by introducing equipment and research methods that soon found use 

in studying true hypersonic flows. At Peenemunde, one set of experiments introduced a wind-tunnel nozzle of specialized design and reached Mach 8.8, becoming 

the first to achieve such a speed. Other German work included the design of a 

76,000-horsepower installation that might have reached Mach 10.

The technical literature also contained an introductory discussion of a possible 

application. It appeared within a wartime report by Austria’s Eugen Sänger, who had 

proposed to build a hypersonic bomber that would extend its range by repeatedly 

skipping off the top of the atmosphere like a stone skipping over water. This concept 

did not enter the mainstream of postwar weapons development, which gave pride 

of place to the long-range ballistic missile. Still, Sänger’s report introduced skipping 

entry as a new mode of high-speed flight, and gave a novel suggestion as to how 

wings could increase the range of a rocket-powered vehicle.

Within Langley, ongoing research treated flows that were merely supersonic. 

However, the scientist John Becker wanted to go further and conduct studies of 

hypersonic flows. He already had spent several years at Langley, thereby learning 

his trade as an aerodynamicist. At the same time he still was relatively young, which 

meant that much of his career lay ahead of him. In 1947 he achieved a major 

advance in hypersonics by building its first important research instrument, an 11-

inch wind tunnel that operated at Mach 6.9.

German Work with High-Speed Flows

At the Technische Hochschule in Hannover, early in the twentieth century, the 

physicist Ludwig Prandtl founded the science of aerodynamics. Extending earlier 

work by Italy’s Tullio Levi-Civita, he introduced the concept of the boundary layer. 

He described it as a thin layer of air, adjacent to a wing or other surface, that clings 

to this surface and does not follow the free-stream flow. Drag, aerodynamic friction, 

and heat transfer all arise within this layer. Because the boundary layer is thin, the 

equations of fluid flow simplified considerably, and important aerodynamic complexities became mathematically tractable.1

As early as 1907, at a time when the Wright Brothers had not yet flown in public, 

Prandtl launched the study of supersonic flows by publishing investigations of a 

steam jet at Mach 1.5. He now was at Göttingen University, where he built a small 

supersonic wind tunnel. In 1911 the German government founded the Kaiser-Wilhelm-Gesellschaft, an umbrella organization that went on to sponsor a broad range 

of institutes in many areas of science and engineering. Prandtl proposed to set up 

a center at Göttingen for research in aerodynamics and hydrodynamics, but World 

War I intervened, and it was not until 1925 that this laboratory took shape.
4

Facing the Heat Barrier: A History of Hypersonics

5

First Steps in Hypersonic Research

colleague of Prandtl, had proposed the use 

of such wings in supersonic flight.9 Walter 

Dornberger, director of V-2 development, 

describes witnessing a wind-tunnel test of a 

model’s stability.

The model had “two knifelike, very 

thin, swept-back wings.” Mounted at its 

center of gravity, it “rotated at the slightest touch.” When the test began, a technician opened a valve to start the airflow. In 

Dornberger’s words, 

“The model moved abruptly, turning 

its nose into the oncoming airstream. 

After a few quickly damping oscillations 

of slight amplitude, it lay quiet and 

stable in the air that hissed past it at 4.4 

times the speed of sound. At the nose, 

and at the edges of the wing supports 

and guide mechanism, the shock waves 

could be clearly seen as they traveled 

diagonally backward at a sharp angle.

As the speed of the airflow fell off and the test ended, the model was 

no longer lying in a stable position. It made a few turns around its center 

of gravity, and then it came to a standstill with the nose pointing downward. The experiment Dr. Hermann had wished to show me had succeeded 

perfectly. This projectile, shaped like an airplane, had remained absolutely 

stable at a supersonic speed range of almost 3,500 mph.”10

Work on the A-9 languished for much of the war, for the V-2 offered problems 

aplenty and had far higher priority. But in 1944, as the Allies pushed the Germans 

out of France and the Russians closed in from the east, Dornberger and Von Braun 

faced insistent demands that they pull a rabbit from a hat and increase the V-2’s 

range. The rabbit was the A-9, with its wings promising a range of 465 miles, some 

three times that of the standard V-2.11

Peenemunde’s Ludwig Roth proceeded to build two prototypes. The V-2 was 

known to its builders as the A-4, and Roth’s A-9 now became the A-4b, a designation 

that allowed it to share in the high priority of that mainstream program. The A-4b 

took shape as a V-2 with swept wings and with a standard set of fins that included 

slightly enlarged air vanes for better control. Certainly the A-4b needed all the help it 

could get, for the addition of wings had made it highly sensitive to winds.

design was trickier than it looked. They could not be too wide, or the V-2 would be 

unable to pass through railroad tunnels. Nor could they extend too far below the 

body of the missile, or the rocket exhaust, expanding at high altitude, would burn 

them off.

The historian Michael Neufeld notes that during the 1930s, “no one knew how 

to design fins for supersonic flight.” The A-3, a test missile that preceded the V-2, 

had proven to be too stable; it tended merely to rise vertically, and its guidance 

system lacked the authority to make it tilt. Its fins had been studied in the Aachen 

supersonic tunnel, but this problem showed up only in flight test, and for a time 

it was unclear how to go further. Hermann Kurzweg, Rudolf Hermann’s assistant, 

investigated low-speed stability building a model and throwing it off the roof of his 

home. When that proved unsatisfactory, he mounted it on a wire, attached it to his 

car, and drove down an autobahn at 60 mph.

The V-2 was to fly at Mach 5, but for a time there was concern that it might 

not top Mach 1. The sound barrier loomed as potentially a real barrier, difficult to 

pierce, and at that time people did not know how to build a transonic wind tunnel 

that would give reliable results. Investigators studied this problem by building heavy 

iron models of this missile and dropping them from a Heinkel He-111 bomber. 

Observers watched from the ground; in one experiment, Von Braun himself piloted 

a plane and dove after the model to observe it from the air. The design indeed 

proved to be marginally unstable in the transonic region, but the V-2 had the thrust 

to power past Mach 1 with ease.

A second test missile, the A-5, also contributed to work on fin design. It supported development of the guidance system, but it too needed fins, and it served as 

a testbed for further flight studies. Additional flight tests used models with length 

of five feet that were powered with rocket engines that flew with hydrogen peroxide 

as the propellant.

These tests showed that an initial fin design given by Kurzweg had the best 

subsonic stability characteristics. Subsequently, extensive wind-tunnel work both 

at Peenemunde and at a Zeppelin facility in Stuttgart covered the V-2’s complete 

Mach range and refined the design. In this fashion, the V-2’s fins were designed with 

only minimal support from Peenemunde’s big supersonic wind tunnels.7

 But these 

tunnels came into their own later in the war, when investigators began to consider 

how to stretch this missile’s range by adding wings and thereby turning it into a 

supersonic glider.

Once the Germans came up with a good configuration for the V-2, they stuck 

with it. They proposed to use it anew in a two-stage missile that again sported fins 

that look excessively large to the modern eye, and that was to cross the Atlantic to 

strike New York.8

 But there was no avoiding the need for a new round of windtunnel tests in studying the second stage of this intercontinental missile, the A-9, 

which was to fly with swept wings. As early as 1935 Adolf Busemann, another 

 A-4b missile ready for launch. (U.S. Army)
6

Facing the Heat Barrier: A History of Hypersonics

7

First Steps in Hypersonic Research

The first A-4b launch took place late in December 1944. It went out of control 

and crashed as the guidance system failed to cope with its demands. Roth’s rocketeers tried again a month later, and General Dornberger describes how this flight 

went much better:

“The rocket, climbing vertically, reached a peak altitude of nearly 50 miles at a 

maximum speed of 2,700 mph. [It] broke the sound barrier without trouble. It flew 

with stability and steered automatically at both subsonic and supersonic speeds. On 

the descending part of the trajectory, soon after the rocket leveled out at the upper 

limit of the atmosphere and began to glide, a wing broke. This structural failure 

resulted from excessive aerodynamic loads.”12

This shot indeed achieved its research goals, for it was to demonstrate successful launch and acceleration through the sound barrier, overcoming drag from the 

wings, and it did these things. Gliding flight was not on the agenda, for while windtunnel tests could demonstrate stability in a supersonic glide, they could not guard 

against atmosphere entry in an improper attitude, with the A-4b tumbling out of 

control.13

Yet while the Germans still had lessons to learn about loads on a supersonic 

aircraft in flight, they certainly had shown that they knew their high-speed aerodynamics. One places their achievement in perspective by recalling that all through the 

1950s a far wealthier and more technically capable United States pursued a vigorous 

program in rocket-powered aviation without coming close to the A-4b’s performance. The best American flight, of an X-2 in 1956, approached 2,100 mph—and 

essentially duplicated the German failure as it went out of control, killing the pilot 

and crashing. No American rocket plane topped the 2,700 mph of the A-4b until 

the X-15 in 1961.14

Hence, without operating in the hypersonic regime, the Peenemunde wind tunnels laid important groundwork as they complemented such alternative research 

techniques as dropping models from a bomber and flying scale models under rocket 

power. Moreover, the Peenemunde aerodynamicist Siegfried Erdmann used his center’s facilities to conduct the world’s first experiments with a hypersonic flow.

In standard operation, at speeds up to Mach 4.4, the Peenemunde tunnels had 

been fed with air from the outside world, at atmospheric pressure. Erdmann knew 

that a hypersonic flow needed more, so he arranged to feed his tunnel with compressed air. He also fabricated a specialized nozzle and aimed at Mach 8.8, twice the 

standard value. His colleague Peter Wegener describes what happened:

“Everything was set for the first-ever hypersonic flow experiment. The highest 

possible pressure ratio across the test section was achieved by evacuating the sphere 

to the limit the remaining pump could achieve. The supply of the nozzle—in contrast to that at lower Mach numbers—was now provided by air at a pressure of 

about 90 atmospheres.... The experiment was initiated by opening the fast-acting 

valve. The flow of brief duration looked perfect as viewed via the optical system. 

Beautiful photographs of the flow about wedge-shaped models, cylinders, spheres, 

and other simple shapes were taken, photographs that looked just as one would 

expect from gas dynamics theory.”15

These tests addressed the most fundamental of issues: How, concretely, does one 

operate a hypersonic wind tunnel? Supersonic tunnels had been bedeviled by condensation of water vapor, which had necessitated the use of silica gel to dry the air. 

A hypersonic facility demanded far greater expansion of the flow, with consequent 

temperatures that were lower still. Indeed, such flow speeds brought the prospect of 

condensation of the air itself.

Conventional handbooks give the liquefaction temperatures of nitrogen and 

oxygen, the main constituents of air, respectively as 77 K and 90 K. These refer 

to conditions at atmospheric pressure; at the greatly rarefied pressures of flow in 

a hypersonic wind tunnel, the pertinent temperatures are far lower.16 In addition, 

Erdmann hoped that his air would “supersaturate,” maintaining its gaseous state 

because of the rapidity of the expansion and hence of the cooling.

This did not happen. In Wegener’s words, “Looking at the flow through the glass 

walls, one could see a dense fog. We know now that under the conditions of this 

particular experiment, the air had indeed partly condensed. The fog was made up of 

air droplets or solid air particles forming a cloud, much like the water clouds we see 

in the sky.”17 To prevent such condensation, it proved necessary not only to feed a 

hypersonic wind tunnel with compressed air, but to heat this air strongly.

One thus is entitled to wonder whether the Germans would have obtained useful 

results from their most ambitious wind-tunnel project, a continuous-flow system 

that was designed to achieve Mach 7, with a possible extension to Mach 10. Its 

power ratings pointed to the advantage of blowdown facilities, such as those of 

Peenemunde. The Mach 4.4 Peenemunde installations used a common vacuum 

sphere, evacuation of which relied on pumps with a total power of 1,100 horsepower. Similar power levels were required to dry the silica gel by heating it, after it 

became moist. But the big hypersonic facility was to have a one-meter test section 

and demanded 76,000 horsepower, or 57 megawatts.18

Such power requirements went beyond what could be provided in straightforward fashion, and plans for this wind tunnel called for it to use Germany’s largest 

hydroelectric plant. Near Kochel in Bavaria, two lakes—the Kochelsee and Walchensee—are separated in elevation by 660 feet. They stand close together, providing an ideal site for generating hydropower, and a hydro plant at that location had 

gone into operation in 1925, generating 120 megawatts. Since the new wind tunnel 

would use half of this power entirely by itself, the power plant was to be enlarged, 

with additional water being provided to the upper lake by a tunnel through the 

mountains to connect to another lake.19

In formulating these plans, as with the A-4b, Germany’s reach exceeded its grasp. 

Moreover, while the big hypersonic facility was to have generous provision for 
8

Facing the Heat Barrier: A History of Hypersonics

9

First Steps in Hypersonic Research

drying its air, there was nothing to prevent the air from condensing, which would 

have thrown the data wildly off.20 Still, even though they might have had to learn 

their lessons in the hard school of experience, Germany was well on its way toward 

developing a true capability in hypersonics by the end of World War II. And among 

the more intriguing concepts that might have drawn on this capability was one by 

the Austrian rocket specialist Eugen Sänger.

Eugen Sänger

Born in 1905, he was of the generation that came of age as ideas of space flight 

were beginning to germinate. Sänger’s own thoughts began to take shape while he 

was still in grammar school. His physics teacher gave him, as a Christmas present, a 

copy of a science-fiction novel, Auf Zwei Planeten (“On Two Planets”). “I was about 

16 years old,” Sänger later recalled. “Naturally I read this novel avidly, and thereafter 

dreamed of doing something like this in my own lifetime.” He soon broadened his 

readings with the classic work of Hermann Oberth. “I had to pass my examination 

in mechanics,” he continued, “and had, therefore, made a particular study of this 

and related subjects. Then I also started to check and recalculate in detail everything 

in Oberth’s book, and I became convinced that here was something that one could 

take seriously.”

He then attended the Technische Hochschule in Vienna, where he tried to win 

a doctoral degree in 1928 by submitting a dissertation on the subject of rocketpowered aircraft. He did not get very far, later recalling that his professor told him, 

“If you try, today, to take your doctor degree in spaceflight, you will most probably 

be an old man with a long beard before you have succeeded in obtaining it.” He 

turned his attention to a more conventional topic, the structural design of wings 

for aircraft, and won his degree a year later. But his initial attempt at a dissertation 

had introduced him to the line of study that he pursued during the next decade and 

then during the war.

In 1933 he turned this dissertation into a book, Raketenflugtechnik. It was the 

first text in this new field. He wrote of a rocket plane burning liquid oxygen and 

petrol, which was to reach Mach 10 along with altitudes of 60 to 70 miles. This concept was significant at the time, for the turbojet engine had not yet been invented, 

and futurists, such as Aldous Huxley who wrote Brave New World, envisioned rockets as the key to high-speed flight in centuries to come.21

Sänger’s altitudes became those of the X-15, a generation later. The speed of his 

concept was markedly higher. He included a three-view drawing. Its wings were 

substantially larger than those of eventual high-performance aircraft, although these 

wings gave his plane plenty of lift at low speed, during takeoff and landing. Its tail 

surfaces also were far smaller than those of the X-15, for he did not know about the Rocket aircraft of Eugen Sänger. Top, the Silbervogel. Bottom, the Amerika-Bomber that was to 

use a skipping entry. Note that both were low-wing monoplanes. (Courtesy of Willy Ley)
10

Facing the Heat Barrier: A History of Hypersonics

11

First Steps in Hypersonic Research

stability problems that loomed in supersonic flight. Still, he clearly had a concept 

that he could modify through further study.

In 1934, writing in the magazine Flug (“Flight”), he used an exhaust velocity of 

3,700 meters per second and gave a velocity at a cutoff of Mach 13. His Silbervogel, 

Silver Bird, now was a boost-glide vehicle, entering a steady glide at Mach 3.5 and 

covering 5,000 kilometers downrange while descending from 60 to 40 kilometers 

in altitude.

He stayed on at the Hochschule and conducted rocket research. Then in 1935, 

amid the Depression, he lost his job. He was in debt to the tune of DM 2,000, 

which he had incurred for the purpose of publishing his book, but he remained 

defiant as he wrote, “Nevertheless, my silver birds will fly!” Fortunately for him, 

at that time Hitler’s Luftwaffe was taking shape, and was beginning to support a 

research establishment. Sänger joined the DVL, the German Experimental Institute 

for Aeronautics, where he worked as technical director of rocket research. He did 

not go to Peenemunde and did not deal with the V-2, which was in the hands of 

the Wehrmacht, not the Luftwaffe. But once again he was employed, and he soon 

was out of debt.

He also began collaborating with the mathematician Irene Bredt, whom he later 

married. His Silbervogel remained on his mind as he conducted performance studies 

with help from Bredt, hoping that this rocket plane might evolve into an AmerikaBomber. He was aware that when transitioning from an initial ballistic trajectory 

into a glide, the craft was to re-enter the atmosphere at a shallow angle. He then 

wondered what would happen if the angle was too steep.

He and Bredt found that rather than enter a glide, the vehicle might develop so 

much lift that it would fly back to space on a new ballistic arc, as if bouncing off 

the atmosphere. Stones skipping over water typically make several such skips, and 

Sänger found that his winged craft would do this as well. With a peak speed of 3.73 

miles per second, compared with 4.9 miles per second as the Earth’s orbital velocity, 

it could fly halfway around the world and land in Japan, Germany’s wartime ally. 

At 4.4 miles per second, the craft could fly completely around the world and land 

in Germany.22

Sänger wrote up their findings in a document of several hundred pages, with 

the title (in English) of “On a Rocket Propulsion for Long Distance Bombers.” In 

December 1941 he submitted it for publication—and won a flat rejection the following March. This launched him into a long struggle with the Nazi bureaucracy, 

as he sought to get his thoughts into print.

His rocket craft continued to show a clear resemblance to his Silbervogel of the 

previous decade, for he kept the basic twin-tailed layout even as he widened the 

fuselage and reduced the size of the wings. Its bottom was flat to produce more lift, 

and his colleagues called it the Platteisen, the Flatiron. But its design proved to be 

patentable, and in June 1942 he received a piece of bright news as the government 

awarded him a Reichspatent concerning “Gliding Bodies for Flight Velocities Above 

Mach 5.” As he continued to seek publication, he won support from an influential 

professor, Walter Georgii. He cut the length of his manuscript in half. Finally, in 

September 1944 he learned that his document would be published as a Secret Command Report.

The print run came to fewer than a hundred copies, but they went to the people 

who counted. These included the atomic-energy specialist Werner Heisenberg, the 

planebuilder Willy Messerschmitt, the chief designer Kurt Tank at Focke-Wulf, 

Ernst Heinkel of Heinkel Aircraft, Ludwig Prandtl who still was active, as well as 

Wernher von Braun and his boss, General Dornberger. Some copies reached the 

Allies after the Nazi surrender, with three of them being taken to Moscow. There 

their content drew attention from the dictator Josef Stalin, who ordered a full translation. He subsequently decided that Sänger and Bredt were to be kidnapped and 

brought to Moscow.

At that time they were in Paris, working as consultants for the French air force. 

Stalin sent two agents after them, accompanied by his own son. They nevertheless 

remained safe; the Soviets never found them. French intelligence agents learned 

about the plot and protected them, and in any case, the Soviets may not have been 

looking very hard. One of them, Grigory Tokaty-Tokayev, was the chief rocket scientist in the Soviet air force. He defected to England, where he wrote his memoirs 

for the Daily Express and then added a book, Stalin Means War.

Sänger, for his part, remained actively involved with his rocket airplane. He succeeded in publishing some of the material from his initial report that he had had 

to delete. He also won professional recognition, being chosen in 1951 as the first 

president of the new International Astronautical Federation. He died in 1964, not 

yet 60. But by then the X-15 was flying, while showing more than a casual resemblance to his Silbervogel of 30 years earlier. His Silver Bird indeed had flown, even 

though the X-15 grew out of ongoing American work with rocket-powered aircraft 

and did not reflect his influence. Still, in January of that year—mere weeks before 

he died—the trade journal Astronautics & Aeronautics published a set of articles that 

presented new concepts for flight to orbit. These showed that the winged-rocket 

approach was alive and well.23

What did he contribute? He was not the first to write of rocket airplanes; that 

palm probably belongs to his fellow Austrian Max Valier, who in 1927 discussed 

how a trimotor monoplane of the day, the Junkers G-23, might evolve into a rocket 

ship. This was to happen by successively replacing the piston motors with rocket 

engines and reducing the wing area.24 In addition, World War II saw several military 

rocket-plane programs, all of which were piloted. These included Germany’s Me163 and Natter antiaircraft weapons as well as Japan’s Ohka suicide weapon, the 
12

Facing the Heat Barrier: A History of Hypersonics

13

First Steps in Hypersonic Research

Cherry Blossom, which Americans called Baka, “Fool.” The rocket-powered Bell 

X-1, with which Chuck Yeager first broke the sound barrier, also was under development well before war’s end.25

Nor did Sänger’s 1944 concept hold military value. It was to be boosted by a 

supersonic rocket sled, which would have been both difficult to build and vulnerable to attack. Even then, and with help from its skipping entry, it would have been 

a single-stage craft attaining near-orbital velocity. No one then, 60 years ago, knew 

how to build such a thing. Its rocket engine lay well beyond the state of the art. 

Sänger projected a mass-ratio, or ratio of fueled to empty weight, of 10—with the 

empty weight including that of the wings, crew compartment, landing gear, and 

bomb load. Structural specialists did not like that. They also did not like the severe 

loads that skipping entry would impose. And after all this Sturm und drang, the 

bomb load of 660 pounds would have been militarily useless.26

But Sänger gave a specific design concept for his rocket craft, presenting it in sufficient detail that other engineers could critique it. Most importantly, his skipping 

entry represented a new method by which wings might increase the effectiveness 

of a rocket engine. This contribution did not go away. The train of thought that 

led to the Air Force’s Dyna-Soar program, around 1960, clearly reflected Sänger’s 

influence. In addition, during the 1980s the German firm of Messerschmitt-Boelkow-Blohm conducted studies of a reusable wing craft that was to fly to orbit as 

a prospective replacement for America’s space shuttle. The name of this two-stage 

vehicle was Sänger.27

NACA-Langley and John Becker

During the war the Germans failed to match the Allies in production of airplanes, but they were well ahead in technical design. This was particularly true in 

the important area of jet propulsion. They fielded an operational jet fighter, the 

Me-262, and while the Yankees were well along in developing the Lockheed P-80 as 

a riposte, the war ended before any of those jets could see combat. Nor was the Me262 a last-minute work of desperation. It was a true air weapon that showed better 

speed and acceleration than the improved P-80A in flight test, while demonstrating an equal rate of climb.28 Albert Speer, Hitler’s minister of armaments, asserted 

in his autobiographical Inside the Third Reich (1970) that by emphasizing production of such fighters and by deploying the Wasserfall antiaircraft missile that was in 

development, the Nazis “would have beaten back the Western Allies’ air offensive 

against our industry from the spring of 1944 on.”29 The Germans thus might have 

prolonged the war until the advent of nuclear weapons.

Wartime America never built anything resembling the big Mach 4.4 wind tunnels 

at Peenemunde, but its researchers at least constructed facilities that could compare 

with the one at Aachen. The American installations did not achieve speeds to match 

Aachen’s Mach 3.3, but they had larger test sections. Arthur Kantrowitz, a young 

physicist from Columbia University who was working at Langley, built a nine-inch 

tunnel that reached Mach 2.5 when it entered operation in 1942. (Aachen’s had 

been four inches.) Across the country, at NACA’s Ames Aeronautical Laboratory, 

two other wind tunnels entered service during 1945. Their test sections measured 

one by three feet, and their flow speeds reached Mach 2.2.30

The Navy also was active. It provided $4.5 million for the nation’s first really 

large supersonic tunnel, with a test section six feet square. Built at NACA-Ames, 

operating at Mach 1.3 to 1.8, this installation used 60,000 horsepower and entered 

service soon after the war.31 The Navy also set up its Ordnance Aerophysics Laboratory in Daingerfield, Texas, adjacent to the Lone Star Steel Company, which had 

air compressors that this firm made available. The supersonic tunnel that resulted 

covered a range of Mach 1.25 to 2.75, with a test section of 19 by 27.5 inches. It 

became operational in June 1946, alongside a similar installation that served for 

high-speed engine tests.32

Theorists complemented the wind-tunnel builders. In April 1947 Theodore von 

Karman, a professor at Caltech who was widely viewed as the dean of American 

aerodynamicists, gave a review and survey of supersonic flow theory in an address 

to the Institute of Aeronautical Sciences. His lecture, published three months later 

in the Journal of the Aeronautical Sciences, emphasized that supersonic flow theory 

now was mature and ready for general use. Von Karman pointed to a plethora of 

available methods and solutions that not only gave means to attack a number of 

important design problems but also gave independent approaches that could permit 

cross-checks on proposed solutions.

John Stack, a leading Langley aerodynamicist, noted that Prandtl had given a 

similarly broad overview of subsonic aerodynamics a quarter-century earlier. Stack 

declared, “Just as Prandtl’s famous paper outlined the direction for the engineer in 

the development of subsonic aircraft, Dr. von Karman’s lecture outlines the direction for the engineer in the development of supersonic aircraft.”33

Yet the United States had no facility, and certainly no large one, that could reach 

Mach 4.4. As a stopgap, the nation got what it wanted by seizing German wind tunnels. A Mach 4.4 tunnel was shipped to the Naval Ordnance Laboratory in White 

Oak, Maryland. Its investigators had fabricated a Mach 5.18 nozzle and had conducted initial tests in January 1945. In 1948, in Maryland, this capability became 

routine.34 Still, if the U.S. was to advance beyond the Germans and develop the true 

hypersonic capability that Germany had failed to achieve, the nation would have to 

rely on independent research.

The man who pursued this research, and who built America’s first hypersonic 

tunnel, was Langley’s John Becker. He had been at that center since 1936; during 
14

Facing the Heat Barrier: A History of Hypersonics

15

First Steps in Hypersonic Research

the latter part of the war he was assistant chief of Stack’s Compressibility Research 

Division. He specifically was in charge of Langley’s 16-Foot High-Speed Tunnel, 

which had fought its war by investigating cooling problems in aircraft motors as 

well as the design of propellers. This facility contributed particularly to tests of 

the B-50 bomber and to the aerodynamic shapes of the first atomic bombs. It also 

assisted development of the Pratt & Whitney R-2800 Double Wasp, a widely used 

piston engine that powered several important wartime fighter planes, along with the 

DC-6 airliner and the C-69 transport, the military version of Lockheed’s Constellation.35

It was quite a jump from piston-powered warbirds to hypersonics, but Becker 

willingly made the leap. The V-2, flying at Mach 5, gave him his justification. In 

a memo to Langley’s chief of research, dated 3 August 1945, Becker noted that 

planned facilities were to reach no higher than Mach 3. He declared that this was 

inadequate: “When it is considered that all of these tunnels will be used, to a large 

extent, to develop supersonic missiles and projectiles of types which have already 

been operated at Mach numbers as high as 5.0, it appears that there is a definite 

need for equipment capable of higher test Mach numbers.”

Within this memo, he outlined a design concept for “a supersonic tunnel having 

a test section four-foot square and a maximum test Mach number of 7.0.” It was to 

achieve continuous flow, being operated by a commercially-available compressor of 

2,400 horsepower. To start the flow, the facility was to hold air within a tank that 

was compressed to seven atmospheres. This air was to pass through the wind tunnel 

before exhausting into a vacuum tank. With pressure upstream pushing the flow 

and with the evacuated tank pulling it, airspeeds within the test section would be 

high indeed. Once the flow was started, the compressor would maintain it.

A preliminary estimate indicated that this facility would cost $350,000. This was 

no mean sum, and Becker’s memo proposed to lay groundwork by first building a 

model of the big tunnel, with a test section only one foot square. He recommended 

that this subscale facility should “be constructed and tested before proceeding with 

a four-foot-square tunnel.” He gave an itemized cost estimate that came to $39,550, 

including $10,000 for installation and $6,000 for contingency.

Becker’s memo ended in formal fashion: “Approval is requested to proceed with 

the design and construction of a model supersonic tunnel having a one-foot-square 

test section at Mach number 7.0. If successful, this model tunnel would not only 

provide data for the design of economical high Mach number supersonic wind tunnels, but would itself be a very useful research tool.”36

On 6 August, three days after Becker wrote this memo, the potential usefulness of this tool increased enormously. On that day, an atomic bomb destroyed 

Hiroshima. With this, it now took only modest imagination to envision nucleartipped V-2s as weapons of the future. The standard V-2 had carried only a one-ton 

conventional warhead and lacked both range and accuracy. It nevertheless had been 

technically impressive, particularly since there was no way to shoot it down. But an 

advanced version with an atomic warhead would be far more formidable.

John Stack strongly supported Becker’s proposal, which soon reached the desk of 

George Lewis, NACA’s Director of Aeronautical Research. Lewis worked at NACA’s 

Washington Headquarters but made frequent visits to Langley. Stack discussed the 

proposal with Lewis in the course of such a visit, and Lewis said, “Let’s do it.”

Just then, though, there was little money for new projects. NACA faced a postwar budget cut, which took its total appropriation from $40.9 million in FY 1945 

to $24 million in FY 1946. Lewis therefore said to Stack, “John, you know I’m a 

sucker for a new idea, but don’t call it a wind tunnel because I’ll be in trouble with 

having to raise money in a formal way. That will necessitate Congressional review 

and approval. Call it a research project.” Lewis designated it as Project 506 and 

obtained approval from NACA’s Washington office on 18 December.37

A month later, in January 1946, Becker raised new issues in a memo to Stack. 

He was quite concerned that the high Mach would lead to so low a temperature that 

air in the flow would liquefy. To prevent this, he called for heating the air, declaring that “a temperature of 600oF in the pressure tank is essential.” He expected to 

achieve this by using “a small electrical heater.”

The pressure in that tank was to be considerably higher than in his plans of 

August. The tank would hold a pressure of 100 atmospheres. Instead of merely 

starting the flow, with a powered compressor sustaining in continuous operation, 

this pressure tank now was to hold enough air for operating times of 40 seconds. 

This would resolve uncertainties in the technical requirements for continuous operation. Continuous flows were still on the agenda but not for the immediate future. 

Instead, this wind tunnel was to operate as a blowdown facility.

Here, in outline, was a description of the installation as finally built. Its test section was 11 inches square. Its pressure tank held 50 atmospheres. It never received 

a compressor system for continuous flow, operating throughout its life entirely as a 

blowdown wind tunnel. But by heating its air, it indeed operated routinely at speeds 

close to Mach 7.38

Taking the name of 11-Inch Hypersonic Tunnel, it operated successfully for the 

first time on 26 November 1947. It did not heat its compressed air directly within 

the pressure tank, relying instead on an electric resistance heater as a separate component. This heater raised the air to temperatures as high as 900oF, eliminating air 

liquefaction in the test section with enough margin for Mach 8. Specialized experiments showed clearly that condensation took place when the initial temperature was 

not high enough to prevent it. Small particles promoted condensation by serving as 

nuclei for the formation of droplets. Becker suggested that such particles could have 

formed through the freezing of CO2

, which is naturally present in air. Subsequent 

research confirmed this conjecture.39
16

Facing the Heat Barrier: A History of Hypersonics

17

First Steps in Hypersonic Research

The facility showed initial early problems as well as a long-term problem. The 

early difficulties centered on the air heater, which showed poor internal heat conduction, requiring as much as five hours to reach a suitably uniform temperature 

distribution. In addition, copper tubes within the heater produced minute particles of copper oxide, due to oxidation of this metal at high temperature. These 

particles, blown within the hypersonic airstream, damaged test models and instruments. Becker attacked the problem of slow warmup by circulating hot air through 

the heater. To eliminate the problem of oxidation, he filled the heater with nitrogen 

while it was warming up.40

A more recalcitrant difficulty arose because the hot airflow, entering the nozzle, 

heated it and caused it to undergo thermal expansion. The change in its dimensions 

was not large, but the nozzle design was highly sensitive to small changes, with this 

expansion causing the dynamic pressure in the airflow to vary by up to 13 percent 

in the course of a run. Run times were as long as 90 seconds, and because of this, 

data taken at the beginning of a test did not agree with similar data recorded a 

minute later. Becker addressed this by fixing the angle of attack of each test model. 

He did not permit the angle to vary during a run, even though variation of this angle 

would have yielded more data. He also made measurements at a fixed time during 

each run.41

The wind tunnel itself represented an important object for research. No similar 

facility had ever been built in America, and it was necessary to learn how to use it 

most effectively. Nozzle design represented an early topic for experimental study. At 

Mach 7, according to standard tables, the nozzle had to expand by a ratio of 104.1 

to 1. This nozzle resembled that of a rocket engine. With an axisymmetric design, 

a throat of one-inch diameter would have opened into a channel having a diameter 

slightly greater than 10 inches. However, nozzles for Becker’s facility proved difficult 

to develop.

Conventional practice, carried over from supersonic wind tunnels, called for a 

two-dimensional nozzle. It featured a throat in the form of a narrow slit, having the 

full width of the main channel and opening onto that channel. However, for flow at 

Mach 7, this slit was to be only about 0.1 inch high. Hence, there was considerable 

interest in nozzles that might be less sensitive to small errors in fabrication.42

Initial work focused on a two-step nozzle. The first step was flat and constant in 

height, allowing the flow to expand to 10 inches wide in the horizontal plane and to 

reach Mach 4.36. The second step maintained this width while allowing the flow to 

expand to 10.5 inches in height, thus achieving Mach 7. But this nozzle performed 

poorly, with investigators describing its flow as “entirely unsatisfactory for use in a 

wind tunnel.” The Mach number reached 6.5, but the flow in the test section was 

“not sufficiently uniform for quantitative wind-tunnel test purposes.” This was due 

to “a thick boundary layer which developed in the first step” along the flat parallel 

walls set closely together at the top and bottom.43

A two-dimensional, single-step nozzle gave much better results. Its narrow slitlike throat indeed proved sensitive; this was the nozzle that gave the variation with 

time of the dynamic pressure. Still, except for this thermal-expansion effect, this 

nozzle proved “far superior in all respects” when compared with the two-step nozzle. 

In turn, the thermal expansion in time proved amenable to correction. This expansion occurred because the nozzle was made of steel. The commercially available alloy 

Invar had a far lower coefficient of thermal expansion. A new nozzle, fabricated 

from this material, entered service in 1954 and greatly reduced problems due to 

expansion of the nozzle throat.44

Another topic of research addressed the usefulness of the optical techniques used 

for flow visualization. The test gas, after all, was simply air. Even when it formed 

shock waves near a model under test, the shocks could not be seen with the unaided 

eye. Therefore, investigators were accustomed to using optical instruments when 

studying a flow. Three methods were in use: interferometry, schlieren, and shadowgraph. These respectively observed changes in air density, density gradient, and the 

rate of change of the gradient.

Such instruments had been in use for decades. Ernst Mach, of the eponymous 

Mach number, had used a shadowgraph as early as 1887 to photograph shock waves 

John Becker’s 11-inch hypersonic wind tunnel. (NASA)
18

Facing the Heat Barrier: A History of Hypersonics

19

First Steps in Hypersonic Research

produced by a speeding bullet. Theodor Meyer, a student of Prandtl, used schlieren to visualize supersonic flow in a nozzle in 1908. Interferometry gave the most 

detailed photos and the most information, but an interferometer was costly and difficult to operate. Shadowgraphs gave the least information but were the least costly 

and easiest to use. Schlieren apparatus was intermediate in both respects and was 

employed often.45

Still, all these techniques depended on the flow having a minimum density. One 

could not visualize shock waves in a vacuum because they did not exist. Highly 

rarefied flows gave similar difficulties, and hypersonic flows indeed were rarefied. 

At Mach 7, a flow of air fell in pressure to less than one part in 4000 of its initial 

value, reducing an initial pressure of 40 atmospheres to less than one-hundredth 

of an atmosphere.46 Higher test-section pressures would have required correspondingly higher pressures in the tank and upstream of the nozzle. But low test-section 

pressures were desirable because they were physically realistic. They corresponded to 

conditions in the upper atmosphere, where hypersonic missiles were to fly.

Becker reported in 1950 that the limit of usefulness of the schlieren method “is 

reached at a pressure of about 1 mm of mercury for slender test models at M = 7.0.”47

This corresponded to the pressure in the atmosphere at 150,000 feet, and there was 

interest in reaching the equivalent of higher altitudes still. A consultant, Joseph 

Kaplan, recommended using nitrogen as a test gas and making use of an afterglow 

that persists momentarily within this gas when it has been excited by an electrical 

discharge. With the nitrogen literally glowing in the dark, it became much easier to 

see shock waves and other features of the flow field at very low pressures.

“The nitrogen afterglow appears to be usable at static pressures as low as 100 

microns and perhaps lower,” Becker wrote.48 This corresponded to pressures of 

barely a ten-thousandth of an atmosphere, which exist near 230,000 feet. It also 

corresponded to the pressure in the test section of a blowdown wind tunnel with air 

in the tank at 50 atmospheres and the flow at Mach 13.8.49 Clearly, flow visualization would not be a problem.

Condensation, nozzle design, and flow visualization were important topics in 

their own right. Nor were they merely preliminaries. They addressed an important 

reason for building this tunnel: to learn how to design and use subsequent hypersonic facilities. In addition, although this 11-inch tunnel was small, there was much 

interest in using it for studies in hypersonic aerodynamics.

This early work had a somewhat elementary character, like the hypersonic experiments of Erdmann at Peenemunde. When university students take initial courses 

in aerodynamics, their textbooks and lab exercises deal with simple cases such as 

flow over a flat plate. The same was true of the first aerodynamic experiments with 

the 11-inch tunnel. The literature held a variety of theories for calculating lift, drag, 

and pressure distributions at hypersonic speeds. The experiments produced data 

that permitted comparison with theory—to check their accuracy and to determine 

circumstances under which they would fail to hold.

One set of tests dealt with cone-cylinder configurations at Mach 6.86. These 

amounted to small and simplified representations of a missile and its nose cone. 

The test models included cones, cylinders with flat ends, and cones with cylindrical afterbodies, studied at various angles of attack. For flow over a cone, the British 

researchers Geoffrey I. Taylor and J. W. Maccoll published a treatment in 1933. 

This quantitative discussion was a cornerstone of supersonic theory and showed its 

merits anew at this high Mach number. An investigation showed that it held “with 

a high degree of accuracy.”

The method of characteristics, devised by Prandtl and Busemann in 1929, was 

a standard analytical method for designing surfaces for supersonic flow, including 

wings and nozzles. It was simple enough to lend itself to hand computation, and it 

gave useful results at lower supersonic speeds. Tests in the 11-inch facility showed 

that it continued to give good accuracy in hypersonic flow. For flow with angle of 

attack, a theory put forth by Antonio Ferri, a leading Italian aerodynamicist, produced “very good results.” Still, not all preexisting theories proved to be accurate. 

One treatment gave good results for drag but overestimated some pressures and 

values of lift.50

Boundary-layer effects proved to be important, particularly in dealing with 

hypersonic wings. Tests examined a triangular delta wing and a square wing, the 

latter having several airfoil sections. Existing theories gave good results for lift and 

drag at modest angles of attack. However, predicted pressure distributions were often 

in error. This resulted from flow separation at high angles of attack—and from the 

presence of thick laminar boundary layers, even at zero angle of attack. These finds 

held high significance, for the very purpose of a hypersonic wing was to generate a 

pressure distribution that would produce lift, without making the vehicle unstable 

and prone to go out of control while in flight.

The aerodynamicist Charles McLellan, who had worked with Becker in designing the 11-inch tunnel and who had become its director, summarized the work 

within the Journal of the Aeronautical Sciences. He concluded that near Mach 7, the 

aerodynamic characteristics of wings and bodies “can be predicted by available theoretical methods with the same order of accuracy usually obtainable at lower speeds, 

at least for cases in which the boundary layer is laminar.”51

At hypersonic speeds, boundary layers become thick because they sustain large 

temperature changes between the wall and the free stream. Mitchel Bertram, a colleague of McLellan, gave an approximate theory for the laminar hypersonic boundary 

layer on a flat plate. Using the 11-inch tunnel, he showed good agreement between 

his theory and experiment in several significant cases. He noted that boundarylayer effects could increase drag coefficients at least threefold, when compared with 
20

Facing the Heat Barrier: A History of Hypersonics

21

First Steps in Hypersonic Research

values using theories that include only free-stream flow and ignore the boundary 

layer. This emphasized anew the importance of the boundary layer in producing 

hypersonic skin friction.52

These results were fundamental, both for aerodynamics and for wind-tunnel 

design. With them, the 11-inch tunnel entered into a brilliant career. It had been 

built as a pilot facility, to lay groundwork for a much larger hypersonic tunnel that 

could sustain continuous flows. This installation, the Continuous Flow Hypersonic 

Tunnel (CFHT), indeed was built. Entering service in 1962, it had a 31-inch test 

section and produced flows at Mach 10.53

Still, it took a long time for this big tunnel to come on line, and all through the 

1950s the 11-inch facility continued to grow in importance. At its peak, in 1961, 

it conducted more than 2,500 test runs, for an average of 10 per working day. It 

remained in use until 1972.54 It set the pace with its use of the blowdown principle, 

which eliminated the need for costly continuous-flow compressors. Its run times 

proved to be adequate, and the CFHT found itself hard-pressed to offer much 

that was new. It had been built for continuous operation but found itself used in a 

blowdown mode most of the time. Becker wrote that his 11-inch installation “far 

exceeded” the CFHT “in both the importance and quality of its research output.” 

He described it as “the only ‘pilot tunnel’ in NACA history to become a major 

research facility in its own right.”55

Yet while the work of this wind tunnel was fundamental to the development of 

hypersonics, in 1950 the field of hypersonics was not fundamental to anything in 

particular. Plenty of people expected that America in time would build missiles and 

aircraft for flight at such speeds, but in that year no one was doing so. This soon 

changed, and the key year was 1954. In that year the Air Force embraced the X-15, 

a hypersonic airplane for which studies in the 11-inch tunnel proved to be essential. 

Also in that year, advances in the apparently unrelated field of nuclear weaponry 

brought swift and emphatic approval for the development of the ICBM. With this, 

hypersonics vaulted to the forefront of national priority.

1 Anderson, History, pp. 251-255.

2 Wegener, Peenemunde, pp. 23-24, 167; Von Karman and Edson, Wind, p. 221.

3 Wegener, Peenemunde, pp. 22-23; Neufeld, Rocket, pp. 87-88.

4 Neufeld, Rocket, p. 87.

5 Wegener, Peenemunde, pp. 24-25; Shapiro, Compressible, pp. 203-04.

6 See, for example, Miller and Durant, Worlds, pp. 9, 17, 23.

7 Neufeld, Rocket, pp. 86, 88-91 (quote, p. 89). Zeppelin: Wattendorf, German, p. 19.

8 Wegener, Peenemunde, photos following p. 84.

9 Von Karman and Edson, Wind, pp. 218-19.

10 Dornberger, V-2, pp. 122-23, 127-28.

11 Hallion, Hypersonic, pp. xvi, xviii; Neufeld, Rocket, pp. 248-50, 283.

12 Neufeld, Rocket, pp. 250-51; Dornberger, V-2, p. 268.

13 Neufeld, Rocket, p. 250.

14 NASA SP-4303, pp. 77, 316, 330.

15 Wegener, Peenemunde, p. 70.

16 Lukasiewicz, Experimental, pp. 71-76.

17 Wegener, Peenemunde, pp. 70-71.

18 Neufeld, Rocket, p. 87; Wegener, Peenemunde, p. 24; Wattendorf, German, p. 4.

19 Wegener, Peenemunde, pp. 32, 75, photos following p. 84.

20 Wattendorf, German, p. 4.

21 Spaceflight, May 1973, pp. 166-71 (quotes, pp. 168, 170); Huxley, Brave, pp. 58, 59, 61.

22 Spaceflight, May 1973, pp. 166, 171-72 (quote, p. 166); Ley, Rockets, pp. 533-537.

23 Spaceflight, May 1973, pp. 171-72, 175-76; Ley, Rockets, pp. 533-34, 535; Ordway and Sharpe, 

Rocket Team, pp. 327-28. 

24 Spaceflight, May 1973, pp. 168-69.

25 Ley, Rockets, pp. 514-19, 524; Allen and Polmar, Downfall, pp. 103, 226.

26 Ordway and Sharpe, Rocket Team, p. 329; Jenkins, Space Shuttle, p. 2.

27 “Sänger.” MBB brochure, August 1986.

28 Boyne, Arrow, p. 139.

29 Speer, Inside, pp. 364-66.

30 Anderson, History, p. 435; NASA: SP-440, pp. 51-52; SP-4305, p. 467.

31 NASA: SP-440, p. 52; SP-4302, pp. 63-64.

32 AIAA Paper 79-0219, pp. 3-4.

33 Journal of the Aeronautical Sciences, July 1947. pp. 373-409 (Stack quote, p. 406). 

34 Hermann, “Supersonic,” p. 439; Anspacher et al., Legacy, pp. 209-10.

35 Becker: Professional resume; NASA SP-4305, p. 54. R-2800 engine: “Dependable Engines” 

(Pratt & Whitney).
22

Facing the Heat Barrier: A History of Hypersonics

36 Memo, Becker to Chief of Research, 3 August 1945 (includes quotes); see also NASA SP-4305, 

pp. 344-346.

37 John Becker interview by Walter Bonney, March 1973 (quotes, p. 4). NACA budget: NASA 

SP-4305, p. 428. Project approval noted in memo, Becker to Stack, 16 January 1946.

38 Memo, Becker to Stack, 16 January 1946 (includes quotes).

39 Journal of Applied Physics, July 1950, pp. 619-21; NACA TN 3302. Air heater: NASA SP-4305, 

p. 471.

40 Becker, memo for record, 23 January 1948.

41 AIAA Paper 88-0230, p. 6.

42 NACA TN 2171, p. 3.

43 Ibid., pp. 6, 21 (quotes, pp. 1, 19).

44 NACA TN 2223 (quote, p. 11); AIAA Paper 88-0230, pp. 6-7.

45 Shapiro, Compressible, pp. 59-68. Photos by Mach and Meyer: Anderson, History, pp. 376, 

382.

46 Shapiro, Compressible, table, p. 620.

47 Journal of Applied Physics, July 1950, pp. 619-28 (quote, p. 621).

48 Ibid. (quote, p. 622).

49 230,000 feet: Shapiro, Compressible, table, pp. 612-13. Mach 13.8: calculated from Shapiro, 

Compressible, eq. 4.14b, p. 83.

50 NACA RM L51J09 (quotes, p. 19).

51 NACA RM L51D17; Journal of the Aeronautical Sciences, October 1951, pp. 641-48 (quote, p. 

648).

52 NACA TN 2773.

53 AIAA Paper 88-0230, pp. 8-9; NACA: SP-440, pp. 94-95; TM X-1130, p. 27.

54 AIAA Paper 88-0230, p. 7.

55 Becker, handwritten notes, January 1989 (includes quotes); NASA: SP-4305, p. 471; RP-1132, 

p. 256.
The ICBM concept of the early 1950s, called Atlas, was intended to carry an 

atomic bomb as a warhead, and there were two things wrong with this missile. It was 

unacceptably large and unwieldy, even with a warhead of reduced weight. In addition, to compensate for this limited yield, Atlas demanded unattainable accuracy in 

aim. But the advent of the hydrogen bomb solved both problems. The weight issue 

went away because projected H-bombs were much lighter, which meant that Atlas 

could be substantially smaller. The accuracy issue also disappeared. Atlas now could 

miss its target by several miles and still destroy it, by the simple method of blowing 

away everything that lay between the aim point and the impact point.

Studies by specialists, complemented by direct tests of early H-bombs, brought a 

dramatic turnaround during 1954 as Atlas vaulted to priority. At a stroke, its designers faced the re-entry problem. They needed a lightweight nose cone that could 

protect the warhead against the heat of atmosphere entry, and nothing suitable was 

in sight. The Army was well along in research on this problem, but its missiles did 

not face the severe re-entry environment of Atlas and its re-entry studies were not 

directly applicable.

The Air Force approached this problem systematically. It began by working 

with the aerodynamicist Arthur Kantrowitz, who introduced the shock tube as an 

instrument that could momentarily reproduce flow conditions that were pertinent. 

Tests with rockets, notably the pilotless X-17, complemented laboratory experiments. The solution to the problem of nose-cone design came from George Sutton, 

a young physicist who introduced the principle of ablation. Test nose cones soon 

were in flight, followed by prototypes of operational versions.

The Move Toward Missiles

In August 1945 it took little imagination to envision that the weapon of the 

future would be an advanced V-2, carrying an atomic bomb as the warhead and able 

to cross oceans. It took rather more imagination, along with technical knowledge, to 

see that this concept was so far beyond the state of the art as not to be worth pursu23

Nose Cones and Re-entry

2
24

Facing the Heat Barrier: A History of Hypersonics

25

Nose Cones and Re-entry

Laboratory had a Nonrotating Engine Branch that was funding development of 

both ramjets and rocket engines. Its director, Weldon Worth, dealt specifically with 

ramjets.5

 A modification of the MX-770 design added two ramjet engines, mounting them singly at the tips of the vertical fins.6

 The missile also received a new name: 

Navaho. This reflected a penchant at North American for names beginning with 

“NA.”7

Then, within a few months during 1949 and 1950, the prospect of world war 

emerged. In 1949 the Soviets exploded their first atomic bomb. At nearly the same 

time, China’s Mao Zedong defeated the Nationalists of Chiang Kai-shek and proclaimed the People’s Republic of China. The Soviets had already shown aggressiveness by subverting the democratic government of Czechoslovakia and by blockading 

Berlin. These new developments raised the prospect of a unified communist empire 

armed with the industry that had defeated the Nazis, wielding atomic weapons, and 

deploying the limitless manpower of China.

President Truman responded both publicly and with actions that were classified. In January 1950 he announced a stepped-up nuclear program, directing “the 

Atomic Energy Commission to continue its work on all forms of atomic weapons, 

including the so-called hydrogen or super bomb.” In April he gave his approval to a 

secret policy document, NSC-68. It stated that the United States would resist communist expansion anywhere in the world and would devote up to twenty percent of 

the gross national product to national defense.8

 Then in June, in China’s back yard, 

North Korea invaded the South, and America again was at war.

These events had consequences for the missile program, as the design and mission of Navaho changed dramatically during 1950. Bollay’s specialists, working with 

Air Force counterparts, showed that they could anticipate increases in its range to 

as much as 5,500 nautical miles. Conferences among Air Force officials, held at 

the Pentagon in August, set this intercontinental range as a long-term goal. A letter 

from Major General Donald Putt, Director of Research and Development within 

the Air Materiel Command, became the directive instructing North American to 

pursue this objective. An interim version, Navaho II, with range of 2,500 nautical 

miles, appeared technically feasible. The full-range Navaho III represented a longterm project that was slated to go forward as a parallel effort.

The thousand-mile Navaho of 1948 had taken approaches based on the V-2 to 

their limit. Navaho II, the initial focus of effort, took shape as a two-stage missile 

with a rocket-powered booster. The booster was to use two such engines, each with 

thrust of 120,000 pounds. A ramjet-powered second stage was to ride it during 

initial ascent, accelerating to the supersonic speed at which the ramjet engines could 

produce their rated thrust. This second stage was then to fly onward as a cruise missile, at a planned flight speed of Mach 2.75.9

A rival to Navaho soon emerged. At Convair, structural analyst Karel Bossart 

held a strong interest in building an ICBM. As a prelude, he had built three rockets 

ing. Thus, in December Vannevar Bush, wartime head of the Office of Scientific 

Research and Development, gave his views in congressional testimony:

“There has been a great deal said about a 3,000 miles high-angle rocket. 

In my opinion, such a thing is impossible for many years. The people have 

been writing these things that annoy me, have been talking about a 3,000 

mile high-angle rocket shot from one continent to another, carrying an 

atomic bomb and so directed as to be a precise weapon which would land 

exactly on a certain target, such as a city. I say, technically, I don’t think 

anyone in the world knows how to do such a thing, and I feel confident 

that it will not be done for a very long period of time to come. I think we 

can leave that out of our thinking.”1

Propulsion and re-entry were major problems, but guidance was worse. For 

intercontinental range, the Air Force set the permitted miss distance at 5,000 feet 

and then at 1,500 feet. The latter equaled the error of experienced bombardiers 

who were using radar bombsights to strike at night from 25,000 feet. The view at 

the Pentagon was that an ICBM would have to do as well when flying all the way 

to Moscow. This accuracy corresponded to hitting a golf ball a mile and having it 

make a hole in one. Moreover, each ICBM was to do this entirely through automatic control.2

The Air Force therefore emphasized bombers during the early postwar years, 

paying little attention to missiles. Its main program, such as it was, called for a missile that was neither ballistic nor intercontinental. It was a cruise missile, which was 

to solve its guidance problem by steering continually. The first thoughts dated to 

November 1945. At North American Aviation, chief engineer Raymond Rice and 

chief scientist William Bollay proposed to “essentially add wings to the V-2 and 

design a missile fundamentally the same as the A-9.”

Like the supersonic wind tunnel at the Naval Ordnance Laboratory, here was 

another concept that was to carry a German project to completion. The initial 

design had a specified range of 500 miles,3

 which soon increased. Like the A-9, 

this missile—designated MX-770—was to follow a boost-glide trajectory and then 

extend its range with a supersonic glide. But by 1948 the U.S. Air Force had won 

its independence from the Army and had received authority over missile programs 

with ranges of 1,000 miles and more. Shorter-range missiles remained the concern of the Army. Accordingly, late in February, Air Force officials instructed North 

American to stretch the range of the MX-770 to a thousand miles.

A boost-glide trajectory was not well suited for a doubled range. At Wright Field, 

the Air Force development center, Colonel M. S. Roth proposed to increase the 

range by adding ramjets.4

 This drew on work at Wright, where the Power Plant 
26

Facing the Heat Barrier: A History of Hypersonics

27

Nose Cones and Re-entry

This concept continued to call for an atomic bomb as the warhead, but by then 

the hydrogen bomb was in the picture. The first test version, named Mike, detonated at Eniwetok Atoll in the Pacific on 1 November 1952. Its fireball spread so far 

and fast as to terrify distant observers, expanding until it was more than three miles 

across. “The thing was enormous,” one man said. “It looked as if it blotted out the 

whole horizon, and I was standing 30 miles away.” The weapons designer Theodore 

Taylor described it as “so huge, so brutal—as if things had gone too far. When the 

heat reached the observers, it stayed and stayed and stayed, not for seconds but for 

minutes.” Mike yielded 10.4 megatons, nearly a thousand times greater than the 13 

kilotons of the Hiroshima bomb of 1945.

Mike weighed 82 tons.14 It was not a weapon; it was a physics experiment. Still, 

its success raised the prospect that warheads of the future might be smaller and 

yet might increase sharply in explosive power. Theodore von Karman, chairman of 

the Air Force Scientific Advisory Board, sought estimates from the Atomic Energy 

Commission of the size and weight of future bombs. The AEC refused to release this 

information. Lieutenant General James Doolittle, Special Assistant to the Air Force 

Chief of Staff, recommended creating a special panel on nuclear weapons within the 

SAB. This took form in March 1953, with the mathematician John von Neumann 

as its chairman. Its specialists included Hans Bethe, who later won the Nobel Prize, 

and Norris Bradbury who headed the nation’s nuclear laboratory at Los Alamos, 

New Mexico.

In June this group reported that a thermonuclear warhead with the 3,000-pound 

Atlas weight could have a yield of half a megaton. This was substantially higher than 

that of the pure-fission weapons considered previously. It gave renewed strength to 

the prospect of a less stringent aim requirement, for Atlas now might miss by far 

more than 1,500 feet and still destroy its target.

Three months later the Air Force Special Weapons Center issued its own estimate, anticipating that a hydrogen bomb of half-megaton yield could weigh as little 

as 1,500 pounds. This immediately opened the prospect of a further reduction in 

the size of Atlas, which might fall in weight from 440,000 pounds to as little as 

240,000. Such a missile also would need fewer engines.15

Also during September, Bruno Augenstein of the Rand Corporation launched 

a study that sought ways to accelerate the development of an ICBM. In Washington, Trevor Gardner was Special Assistant for Research and Development, reporting to the Air Force Secretary. In October he set up his own review committee. He 

recruited von Neumann to serve anew as its chairman and then added a dazzling 

array of talent from Caltech, Bell Labs, MIT, and Hughes Aircraft. In Gardner’s 

words, “The aim was to create a document so hot and of such eminence that no one 

could pooh-pooh it.”16

He called his group the Teapot Committee. He wanted particularly to see it call 

for less stringent aim, for he believed that a 1,500-foot miss distance was preposin the shape of a subscale V-2 and had demonstrated his ideas for lightweight structure in flight test. The Rand Corporation, an influential Air Force think tank, had 

been keeping an eye on this work and on the burgeoning technology of missiles. In 

December 1950 it issued a report stating that long-range ballistic missiles now were 

in reach. A month later the Air Force responded by giving Bossart, and Convair, a 

new study contract. In August 1951 he christened this missile Atlas, after Convair’s 

parent company, the Atlas Corporation.

The initial concept was a behemoth. Carrying an 8,000-pound warhead, it was 

to weigh 670,000 pounds, stand 160 feet tall by 12 feet in diameter, and use seven 

of Bollay’s new 120,000-pound engines. It was thoroughly unwieldy and represented a basis for further studies rather than a concept for a practical weapon. Still, 

it stood as a milestone. For the first time, the Air Force had a concept for an ICBM 

that it could pursue using engines that were already in development.10

For the ICBM to compete with Navaho, it had to shrink considerably. Within 

the Air Force’s Air Research and Development Command, Brigadier General John 

Sessums, a strong advocate of long-range missiles, proposed that this could be done 

by shrinking the warhead. The size and weight of Atlas were to scale in proportion 

with the weight of its atomic weapon, and Sessums asserted that new developments 

in warhead design indeed would give high yield while cutting the weight.

He carried his argument to the Air Staff, which amounted to the Air Force’s 

board of directors. This brought further studies, which indeed led to a welcome 

reduction in the size of Atlas. The concept of 1953 called for a length of 110 feet 

and a loaded weight of 440,000 pounds, with the warhead tipping the scale at only 

3,000 pounds. The number of engines went down from seven to five.11

There also was encouraging news in the area of guidance. Radio guidance was out 

of the question for an operational missile; it might be jammed or the ground-based 

guidance center might be destroyed in an attack. Instead, missile guidance was to 

be entirely self-contained. All concepts called for the use of sensitive accelerometers 

along with an onboard computer, to determine velocity and location. Navaho was 

to add star trackers, which were to null out errors by tracking stars even in daylight. 

In addition, Charles Stark Draper of MIT was pursuing inertial guidance, which 

was to use no external references of any sort. His 1949 system was not truly inertial, 

for it included a magnetic compass and a Sun-seeker. But when flight-tested aboard 

a B-29, over distances as great at 1,737 nautical miles, it showed a mean error of 

only 5 nautical miles.12

For Atlas, though, the permitted miss distance remained at 1,500 feet, with the 

range being 5500 nautical miles. The program plan of October 1953 called for a 

leisurely advance over the ensuing decade, with research and development being 

completed only “sometime after 1964,” and operational readiness being achieved in 

1965. The program was to emphasize work on the major components: propulsion, 

guidance, nose cone, lightweight structure. In addition, it was to conduct extensive 

ground tests before proceeding toward flight.13
28

Facing the Heat Barrier: A History of Hypersonics

29

Nose Cones and Re-entry

terous. The Teapot Committee drew on findings by Augenstein’s group at Rand, 

which endorsed a 1,500-pound warhead and a three-mile miss distance. The formal 

Teapot report, issued in February 1954, declared “the military requirement” on miss 

distance “should be relaxed from the present 1,500 feet to at least two, and probably three, nautical miles.” Moreover, “the warhead weight might be reduced as far 

as 1,500 pounds, the precise figure to be determined after the Castle tests and by 

missile systems optimization.”17

The latter recommendation invoked Operation Castle, a series of H-bomb tests 

that began a few weeks later. The Mike shot of 1952 had used liquid deuterium, 

a form of liquid hydrogen. It existed at temperatures close to absolute zero and 

demanded much care in handling. But the Castle series was to test devices that used 

lithium deuteride, a dry powder that resembled salt. The Mike approach had been 

chosen because it simplified the weapons physics, but a dry bomb using lithium 

promised to be far more practical.

The first such bomb was detonated on 1 March as Castle Bravo. It produced 15 

megatons, as its fireball expanded to almost four miles in diameter. Other Castle 

H-bombs performed similarly, as Castle Romeo went to 11 megatons and Castle 

Yankee, a variant of Romeo, reached 13.5 megatons. “I was on a ship that was 30 

miles away,” the physicist Marshall Rosenbluth recalls about Bravo, “and we had this 

horrible white stuff raining out on us.” It was radioactive fallout that had condensed 

from vaporized coral. “It was pretty frightening. There was a huge fireball with 

these turbulent rolls going in and out. The thing was glowing. It looked to me like a 

diseased brain.” Clearly, though, bombs of the lithium type could be as powerful as 

anyone wished—and these test bombs were readily weaponizable.18

The Castle results, strongly complementing the Rand and Teapot reports, 

cleared the way for action. Within the Pentagon, Gardner took the lead in pushing 

for Atlas. On 11 March he met with Air Force Secretary Harold Talbott and with 

the Chief of Staff, General Nathan Twining. He proposed a sped-up program that 

would nearly double the Fiscal Year (FY) 1955 Atlas budget and would have the first 

missiles ready to launch as early as 1958. General Thomas White, the Vice Chief of 

Staff, weighed in with his own endorsement later that week, and Talbott responded 

by directing Twining to accelerate Atlas immediately.

White carried the ball to the Air Staff, which held responsibility for recommending approval of new programs. He told its members that “ballistic missiles were 

here to stay, and the Air Staff had better realize this fact and get on with it.” Then 

on 14 May, having secured concurrence from the Secretary of Defense, White gave 

Atlas the highest Air Force development priority and directed its acceleration “to 

the maximum extent that technology would allow.” Gardner declared that White’s 

order meant “the maximum effort possible with no limitation as to funding.”19

This was a remarkable turnaround for a program that at the moment lacked 

even a proper design. Many weapon concepts have gone as far as the prototype 

stage without winning approval, but Atlas gained its priority at a time when the 

accepted configuration still was the 440,000-pound, five-engine concept of 1953. 

Air Force officials still had to establish a formal liaison with the AEC to win access 

to information on projected warhead designs. Within the AEC, lightweight bombs 

still were well in the future. A specialized device, tested in the recent series as Castle 

Nectar, delivered 1.69 megatons but weighed 6,520 pounds. This was four times 

the warhead weight proposed for Atlas.

But in October the AEC agreed that it could develop warheads weighing 1,500 

to 1,700 pounds, with a yield of one megaton. This opened the door to a new Atlas 

design having only three engines. It measured 75 feet long and 10 feet in diameter, 

with a weight of 240,000 pounds—and its miss distance could be as great as five miles. 

This took note of the increased yield of the warhead and further eased the problem of 

guidance. The new configuration won Air Force approval in December.20

Approaching the Nose Cone

An important attribute of a nose cone was its shape, and engineers were reducing drag to a minimum by crafting high-speed airplanes that displayed the ultimate 

in needle-nose streamlining. The X-3 research aircraft, designed for Mach 2, had a 

long and slender nose that resembled a church steeple. Atlas went even further, with 

an early concept having a front that resembled a flagpole. This faired into a long and 

slender cone that could accommodate the warhead.21

This intuitive approach fell by the wayside in 1953, as the NACA-Ames aerodynamicists H. Julian Allen and Alfred Eggers carried through an elegant analysis 

of the motion and heating of a re-entering nose cone. This work showed that they 

were masters of the simplifying assumption. To make such assumptions successfully 

represents a high art, for the resulting solutions must capture the most essential 

aspects of the pertinent physics while preserving mathematical tractability. Their 

paper stands to this day as a landmark. Quite probably, it is the single most important paper ever written in the field of hypersonics.

They calculated total heat input to a re-entry vehicle, seeking shapes that would 

minimize this. That part of the analysis enabled them to critique the assertion that 

a slender and sharply-pointed shape was best. For a lightweight nose cone, which 

would slow significantly in the atmosphere due to drag, they found a surprising 

result: the best shape, minimizing the total heat input, was blunt rather than sharp.

The next issue involved the maximum rate of heat transfer when averaged over 

an entire vehicle. To reduce this peak heating rate to a minimum, a nose cone of 

realistic weight might be either very sharp or very blunt. Missiles of intermediate 

slenderness gave considerably higher peak heating rates and “were definitely to be 

avoided.”
30

Facing the Heat Barrier: A History of Hypersonics

31

Nose Cones and Re-entry

This result applied to the entire vehicle, but heat-transfer rates were highest at 

the nose-cone tip. It was particularly important to minimize the heating at the tip, 

and again their analysis showed that a blunt nose cone would be best. As Allen and 

Eggers put it, “not only should pointed bodies be avoided, but the rounded nose 

should have as large a radius as possible.”22

How could this be? The blunt body set up a very strong shock wave, which produced intense heating of the airflow. However, most of this heat was carried away 

in the flow. The boundary layer served to insulate the vehicle, and relatively little of 

this heat reached its surface. By contrast, a sharp and slender nose cone produced 

a shock that stood very close to this surface. At the tip, the boundary layer was too 

thin to offer protection. In addition, skin friction produced still more heating, for 

the boundary layer now received energy from shock-heated air flowing close to the 

vehicle surface.23

This paper was published initially as a classified document, but it took time to 

achieve its full effect. The Air Force did not adopt its principle for nose-cone design 

until 1956.24 Still, this analysis outlined the shape of things to come. Blunt heat 

shields became standard on the Mercury, Gemini, and Apollo capsules. The space 

shuttle used its entire undersurface as a heat shield that was particularly blunt, raising its nose during re-entry to present this undersurface to the flow.

Yet while analysis could indicate the general shape for a nose cone, only experiment could demonstrate the validity of a design. At a stroke, Becker’s Mach 7 facility, which had been far in the forefront only recently, suddenly became inadequate. 

An ICBM nose cone was to re-enter the atmosphere at speeds above Mach 20. Its 

kinetic energy would vaporize five times its weight of iron. Temperatures behind the 

bow shock would reach 9000 K, hotter than the surface of the Sun. Research scientist Peter Rose wrote 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.”25

Though hot, the 9000 K air actually would be cool, considering its situation, 

because its energy would go into dissociating molecules of gas. However, the ions 

and dissociated atoms were only too likely to recombine at the surface of the nose 

cone, thereby delivering additional heat. Such chemical effects also might trip the 

boundary layer from laminar to turbulent flow, with the rate of heat transfer increasing substantially as a result. In the words of Rose:

“The presence of free-atoms, electrons, and molecules in excited states 

can be expected to complicate heat transfer through the boundary layer 

by additional modes of energy transport, such as atom diffusion, carrying 

the energy of dissociation. Radiation by transition from excited energy 

states may contribute materially to radiative heat transfer. There is also a 

possibility of heat transfer by electrons and ions. The existence of large 

amounts of energy in any of these forms will undoubtedly influence the 

familiar flow phenomena.”26

Within the Air Force, the Aircraft Panel of the Scientific Advisory Board (SAB) 

issued a report in October 1954 that looked ahead to the coming decade:

“In the aerodynamics field, it seems to us pretty clear that over the next 10 

years the most important and vital subject for research and development 

is the field of hypersonic flows; and in particular, hypersonic flows with 

[temperatures at a nose-cone tip] which may run up to the order of 

thousands of degrees. This is one of the fields in which an ingenious 

and clever application of the existing laws of mechanics is probably not 

adequate. It is one in which much of the necessary physical knowledge 

still remains unknown at present and must be developed before we arrive 

at a true understanding and competence. The reason for this is that the 

temperatures which are associated with these velocities are higher than 

temperatures which have been produced on the globe, except in connection 

with the nuclear developments of the past 10 or 15 years and that there 

are problems of dissociation, relaxation times, etc., about which the basic 

physics is still unknown.”27

The Atlas program needed a new experimental technique, one that could overcome the fact that conventional wind tunnels produced low temperatures due to 

their use of expanding gases, and hence the pertinent physics and chemistry associated with the heat of re-entry were not replicated. Its officials found what they 

wanted at a cocktail party.

This social gathering took place at Cornell University around Thanksgiving of 

1954. The guests included university trustees along with a number of deans and 

senior professors. One trustee, Victor Emanuel, was chairman of Avco Corporation, which already was closely involved in work on the ICBM. He had been in 

Washington and had met with Air Force Secretary Harold Talbott, who told him of 

his concern about problems of re-entry. Emanuel raised this topic at the party while 

talking with the dean of engineering, who said, “I believe we have someone right 

here who can help you.”28

That man was Arthur Kantrowitz, a former researcher at NACA-Langley who 

had taken a faculty position at Cornell following the war. While at Langley during 

the late 1930s, he had used a $5,000 budget to try to invent controlled thermonuclear fusion. He did not get very far. Indeed, he failed to gain results that were 

sufficient even to enable him to write a paper, leaving subsequent pioneers in con-
32

Facing the Heat Barrier: A History of Hypersonics

33

Nose Cones and Re-entry

trolled fusion to start again from scratch. Still, as he recalls, “I continued my interest in high temperatures with the hope that someday I could find something that I 

could use to do fusion.”29

In 1947 this led him to the shock tube. This instrument produced very strong 

shocks in a laboratory, overcoming the limits of wind tunnels. It used a driving gas 

at high pressure in a separate chamber. This gas burst through a thin diaphragm to 

generate the shock, which traveled down a long tube that was filled with a test gas. 

High-speed instruments could observe this shock. They also could study a small 

model immersed within the hot flow at high Mach that streamed immediately 

behind the shock.30

When Kantrowitz came to the shock tube, it already was half a century old. The 

French chemist Paul Vieille built the first such devices prior to 1900, using them to 

demonstrate that a shock wave travels faster than the speed of sound. He proposed 

that his apparatus could prove useful in studying mine explosions, which took place 

in shafts that resembled his long tubes.31

The next important shock-tube researcher, Britain’s William Payman, worked 

prior to World War II. He used diaphragm-bursting pressures as high as 1100 

pounds per square inch and introduced high-speed photography to observe the 

shocked flows. He and his colleagues used the shock tube for experimental verification of equations in gasdynamics that govern the motion of shock waves.32

At Princeton University during that war, the physicist Walter Bleakney went further. He used shock tubes as precision instruments, writing, “It has been found that 

successive ‘shots’ in the tube taken with the same initial conditions reproduce one 

another to a surprising degree. The velocity of the incident shock can be reproduced 

to 0.1 percent.” He praised the versatility of the device, noting its usefulness “for 

studying a great variety of problems in fluid dynamics.” In addition to observations 

of shocks themselves, the instrument could address “problems of detonation and 

allied phenomena. The tube may be used as a wind tunnel with a Mach number 

variable over an enormous range.” This was the role it took during the ICBM program.33

At Cornell, Kantrowitz initiated a reach for high temperatures. This demanded 

particularly high pressure in the upstream chamber. Payman had simply used compressed air from a thick-walled tank, but Kantrowitz filled his upstream chamber 

with a highly combustible mix of hydrogen and oxygen. Seeking the highest temperatures, he avoided choosing air as a test gas, for its diatomic molecules absorbed 

energy when they dissociated or broke apart, which limited the temperature rise. 

He turned instead to argon, a monatomic gas that could not dissociate, and reached 

18,000 K.

He was a professor at Cornell, with graduate students. One of them, Edwin 

Resler, wrote a dissertation in 1951, “High Temperature Gases Produced by Strong 

Shock Waves.” In Kantrowitz’s hands, the versatility of this instrument appeared 

anew. With argon as the test gas, it served for studies of thermal ionization, a physical effect separate from dissociation in which hot atoms lost electrons and became 

electrically charged. Using nitrogen or air, the shock tube examined dissociation 

as well, which increased with the higher temperatures of stronger shocks. Higher 

Mach values also lay within reach. As early as 1952, Kantrowitz wrote that “it is 

possible to obtain shock Mach numbers in the neighborhood of 25 with reasonable 

pressures and shock tube sizes.”34

Other investigators also worked with these devices. Raymond Seeger, chief of 

aerodynamics at the Naval Ordnance Laboratory, built one. R. N. Hollyer conducted experiments at the University of Michigan. At NACA-Langley, the first shock 

tube entered service in 1951. The Air Force also was interested. The 1954 report of 

the SAB pointed to “shock tubes and other devices for producing extremely strong 

shocks” as an “experimental technique” that could give new insights into fundamental problems of hypersonics.35

Thus, when Emanuel met Kantrowitz at that cocktail party, this academic physicist indeed was in a position to help the Atlas effort. He had already gained handson experience by conducting shock-tube experiments at temperatures and shock 

velocities that were pertinent to re-entry of an ICBM. Emanuel then staked him 

to a new shock-tube center, Avco Research Laboratory, which opened for business 

early in 1955.

Kantrowitz wanted the highest shock velocities, which he obtained by using 

lightweight helium as the driver gas. He heated the helium strongly by adding a 

mixture of gaseous hydrogen and oxygen. Too little helium led to violent burning 

with unpredictable detonations, but use of 70 percent helium by weight gave a controlled burn that was free of detonations. The sudden heating of this driver gas also 

ruptured the diaphragm.

Standard optical instruments, commonly used in wind-tunnel work, were available for use with shock tubes as well. These included the shadowgraph, schlieren 

apparatus, and Mach-Zehnder interferometer. To measure the speed of the shock, 

it proved useful to install ionization-sensitive pickups that responded to changes in 

electrical resistance as shock waves passed. Several such pickups, spaced along the 

length of the tube, gave good results at speeds up to Mach 16.

Within the tube, the shock raced ahead of the turbulent mix of driver gases. 

Between the shock and the driver gases lay a “homogeneous gas sample” (HGS), a 

cylindrical slug of test gas moving nearly with the speed of the shock. The measured 

speed of the shock, together with standard laws of gasdynamics, permitted a complete calculation of the pressure, temperature, and internal energy of the HGS. Even 

when the HGS experienced energy absorption due to dissociation of its constituent 

molecules, it was possible to account for this through a separate calculation.36
34

Facing the Heat Barrier: A History of Hypersonics

35

Nose Cones and Re-entry

The HGS swept over a small model of a nose cone placed within the stress. The 

time for passage was of the order of 100 microseconds, with the shock tube thus 

operating as a “wind tunnel” having this duration for a test. This nevertheless was 

long enough for photography. In addition, specialized instruments permitted study 

of heat transfer. These included thin-gauge resistance thermometers for temperature 

measurements and thicker-gauge calorimeters to determine heat transfer rates.

Metals increase their electrical resistance in response to a temperature rise. Both 

the thermometers and the calorimeters relied on this effect. To follow the sudden 

temperature increase behind the shock, the thermometer needed a metal film that 

was thin indeed, and Avco researchers achieved a thickness of 0.3 microns. They did 

this by using a commercial product, Liquid Bright Platinum No. 05, from Hanovia 

Chemical and Manufacturing Company. This was a mix of organic compounds of 

platinum and gold, dissolved in oils. Used as a paint, it was applied with a brush 

and dried in an oven.

The calorimeters used bulk platinum foil that was a hundred times thicker, at 

0.03 millimeters. This thickness diminished their temperature rise and allowed the 

observed temperature increase to be interpreted as a rate of heat transfer. Both the 

thermometers and calorimeters were mounted to the surface of nose-cone models, 

which typically had the shape of a hemisphere that faired smoothly into a cylinder 

at the rear. The models were made of Pyrex, a commercial glass that did not readily 

crack. In addition, it was a good insulator.37

The investigator Shao-Chi Lin also used a shock tube to study thermal ionization, which made the HGS electrically conductive. To measure this conductivity, 

Shao used a nonconducting shock tube made of glass and produced a magnetic field 

within its interior. The flow of the conducting HGS displaced the magnetic lines 

of force, which he observed. He calibrated the system by shooting a slug of metal 

having known conductivity through the field at a known speed. Measured HGS 

conductivities showed good agreement with values calculated from theory, over a 

range from Mach 10 to Mach 17.5. At this highest flow speed, the conductivity of 

air was an order of magnitude greater than that of seawater.38

With shock tubes generating new data, there was a clear need to complement 

the data with new solutions in aerodynamics and heat transfer. The original AllenEggers paper had given a fine set of estimates, but they left out such realistic effects 

as dissociation, recombination, ionization, and changes in the ratio of specific heats. 

Again, it was necessary to make simplifying assumptions. Still, the first computers 

were at hand, which meant that solutions did not have to be in closed form. They 

might be equations that were solvable electronically.

Recombination of ions and of dissociated diatomic molecules—oxygen and 

nitrogen—was particularly important at high Mach, for this chemical process could 

deliver additional heat within the boundary layer. Two simplified cases stood out. In 

“equilibrium flow,” the recombination took place instantly, responding immediately 

to the changing temperature and pressure within the boundary layer. The extent of 

ionization and dissociation then were simple point functions of the temperature and 

pressure at any location, and they could be calculated directly.

The other limiting case was “frozen flow.” One hesitates to describe a 9000 K 

airstream as “frozen,” but here it meant that the chemical state of the boundary layer 

retained its condition within the free stream behind the bow shock. Essentially this 

means that recombination proceeded so slowly that the changing conditions within 

the boundary layer had no effect on the degrees of dissociation and ionization. These 

again could be calculated directly, although this time as a consequence of conditions 

behind the shock rather than in the boundary layer. Frozen flow occurred when the 

air was rarefied.

These approximations avoided the need to deal with the chemistry of finite reaction rates, wherein recombination would not instantly respond to the rapidly varying flow conditions across the thickness of a boundary layer but would lag behind 

the changes. In 1956 the aerodynamicist Lester Lees proposed a heat-transfer theory 

that specifically covered those two limiting cases.39 Then in 1957, Kantrowitz’s colleagues at Avco Research Laboratory went considerably further.

The Avco lab had access to the talent of nearby MIT. James Fay, a professor 

of mechanical engineering, joined with Avco’s Frederick Riddell to treat anew the 

problem of heat transfer in dissociated air. Finite reaction-rate chemistry was at the 

heart of their agenda, and again they needed a simplifying assumption: that the 

airflow velocity was zero. However, this condition was nearly true at the forward tip 

of a nose cone, where the heating was most severe.

Starting with a set of partial differential equations, they showed that these equations reduced to a set of nonlinear ordinary differential equations. Using an IBM 

650 computer, they found that a numerical solution of these nonlinear equations was 

reasonably straightforward. In dealing with finite-rate chemistry, they introduced 

a “reaction rate parameter” that attempted to capture the resulting effects. They 

showed that a re-entering nose cone could fall through 100,000 feet while transitioning from the frozen to the equilibrium regime. Within this transition region, 

the boundary layer could be expected to be partly frozen, near the free stream, and 

partly in equilibrium, near the wall.

The Fay-Riddell theory appeared in the February 1958 Journal of the Aeronautical Sciences. That same issue presented experimental results, also from Avco, that 

tested the merits of this treatment. The researchers obtained shock-tube data with 

shock Mach numbers as high as 17.5. At this Mach, the corresponding speed of 

17,500 feet per second approached the velocity of a satellite in orbit. Pressures 

within the shock-tube test gas simulated altitudes of 20,000, 70,000, and 120,000 

feet, with equilibrium flow occurring in the models’ boundary layers even at the 

highest equivalent height above the ground.
36

Facing the Heat Barrier: A History of Hypersonics

37

Nose Cones and Re-entry

Most data were taken with calorimeters, although data points from thin-gauge 

thermometers gave good agreement. The measurements showed scatter but fit neatly 

on curves calculated from the Fay-Riddell theory. The Lees theory underpredicted 

heat-transfer rates at the nose-cone tip, calling for rates up to 30 percent lower than 

those observed. Here, within a single issue of that journal, two papers from Avco 

gave good reason to believe that theoretical and experimental tools were at hand 

to learn the conditions that a re-entering ICBM nose cone would face during its 

moments of crisis.40

Still, this was not the same as actually building a nose cone that could survive 

this crisis. This problem called for a separate set of insights. These came from the 

U.S. Army and were also developed independently by an individual: George Sutton 

of General Electric.

Ablation

In 1953, on the eve of the Atlas go-ahead, investigators were prepared to consider several methods for thermal protection of its nose cone. The simplest was 

the heat sink, with a heat shield of thick copper absorbing the heat of re-entry. 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 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.41

Still, each of these approaches held difficulties. Though potentially valuable, 

transpiration cooling was poorly understood as a topic for design. The hot-structure 

concept raised questions of suitably refractory metals along with the prospect of 

losing the entire nose cone if a shingle came off. The heat-sink approach was likely 

to lead to high weight. Even so, it seemed to be the most feasible way to proceed, 

and early Atlas designs specified use of a heat-sink nose cone.42

The Army had its own activities. Its missile program was separate from that of 

the Air Force and was centered in Huntsville, Alabama, with the redoubtable Wernher von Braun as its chief. He and his colleagues came to Huntsville in 1950 and 

developed the Redstone missile as an uprated V-2. It did not need thermal protection, but the next missile would have longer range and would certainly need it.43

Von Braun was an engineer. He did not set up a counterpart of Avco Research 

Laboratory, but his colleagues nevertheless proceeded to invent their way toward a 

nose cone. Their concern lay at the tip of a rocket, but their point of departure came 

at the other end. They were accustomed to steering their missiles by using jet vanes, 

large tabs of heat-resistant material that dipped into the exhaust. These vanes then 

deflected the exhaust, changing the direction of flight. Von Braun’s associates thus 

had long experience in testing materials by placing them within the blast of a rocket 

engine. This practice carried over to their early nose-cone work.44

The V-2 had used vanes of graphite. In November 1952, these experimenters 

began testing new materials, including ceramics. They began working with nosecone models late in 1953. In July 1954 they tested their first material of a new type: 

a reinforced plastic, initially a hard melamine resin strengthened with glass fiber. 

New test facilities entered service in June 1955, including a rocket engine with 

thrust of 20,000 pounds and a jet diameter of 14.5 inches.45

The pace accelerated after November of that year, as Von Braun won approval 

from Defense Secretary Charles Wilson to proceed with development of his next 

missile. This was Jupiter, with a range of 1,500 nautical miles.46 It thus was markedly less demanding than Atlas in its thermal-protection requirements, for it was to 

re-enter the atmosphere at Mach 15 rather than Mach 20 and higher. Even so, the 

Huntsville group stepped up its work by introducing new facilities. These included 

a rocket engine of 135,000 pounds of thrust for use in nose-cone studies.

The effort covered a full range of thermal-protection possibilities. Transpiration cooling, for one, raised unpleasant new issues. Convair fabricated test nose 

cones with water tanks that had porous front walls. The pressure in a tank could be 

adjusted to deliver the largest flow of steam when the heat flux was greatest. But this 

technique led to hot spots, where inadequate flow brought excessive temperatures. 

Transpiration thus fell by the wayside.

Heat sink drew attention, with graphite holding promise for a time. It was light 

in weight and could withstand high temperatures. But it also was a good heat conductor, which raised problems in attaching it to a substructure. Blocks of graphite 

also contained voids and other defects, which made them unusable.

By contrast, hot structures held promise. Researchers crafted lightweight shingles of tungsten and molybdenum backed by layers of polished corrugated steel 

and aluminum, to provide thermal insulation along with structural support. When 

the shingles topped 3,250oF, the innermost layer stayed cool and remained below 

200oF. Clearly, hot structures had a future.

The initial work with a reinforced plastic, in 1954, led to many more tests of 

similar materials. Engineers tested such resins as silicones, phenolics, melamines, 

Teflon, epoxies, polyesters, and synthetic rubbers. Filler materials included soft 

glass, fibers of silicon dioxide and aluminum silicate, mica, quartz, asbestos, nylon, 

graphite, beryllium, beryllium oxide, and cotton.
38

Facing the Heat Barrier: A History of Hypersonics

39

Nose Cones and Re-entry

Fiber-reinforced polymers proved to hold particular merit. The studies focused on 

plastics reinforced with glass fiber, with a commercially-available material, Micarta 

259-2, demonstrating noteworthy promise. The Army stayed with this choice as it 

moved toward flight test of subscale nose cones in 1957. The first one used Micarta 

259-2 for the plastic, with a glass cloth as the filler.47

In this fashion the Army ran well ahead of the Air Force. Yet the Huntsville 

work did not influence the Atlas effort, and the reasons ran deeper than interservice rivalry. The relevance of that work was open to question because Atlas faced a 

far more demanding re-entry environment. In addition, Jupiter faced competition 

from Thor, an Air Force missile of similar range. It was highly likely that only one 

would enter production, so Air Force designers could not merely become apt pupils 

of the Army. They had to do their own work, seeking independent approaches and 

trying to do better than Von Braun.

Amid this independence, George Sutton came to the re-entry problem. He had 

received his Ph.D. at Caltech in 1955 at age 27, jointly in mechanical engineering 

and physics. His only experience within the aerospace industry had been a summer 

job at the Jet Propulsion Laboratory, but he jumped into re-entry with both feet 

after taking his degree. He joined Lockheed and became closely involved in studying materials suitable for thermal protection. Then he was recruited by General 

Electric, leaving sunny California and arriving in snowy Schenectady, New York, 

early in 1956.

Heat sinks for Atlas were ascendant at that time, with Lester Lees’s heat-transfer 

theory appearing to give an adequate account of the thermal environment. Sutton 

was aware of the issues and wrote a paper on heat-sink nose cones, but his work 

soon led him in a different direction. There was interest in storing data within a 

small capsule that would ride with a test nose cone and that might survive re-entry 

if the main cone were to be lost. This capsule needed its own thermal protection, 

and it was important to achieve light weight. Hence it could not use a heat sink. 

Sutton’s management gave him a budget of $75,000 to try to find something more 

suitable.48

This led him to re-examine the candidate materials that he had studied at Lockheed. He also learned that other GE engineers were working on a related problem. 

They had built liquid propellant rocket engines for the Army’s Hermes program, 

with these missiles being steered by jet vanes in the fashion of the V-2 and Redstone. 

The vanes were made from alternating layers of glass cloth and thermosetting resins. 

They had become standard equipment on the Hermes A-3, but some of them failed 

due to delamination. Sutton considered how to avoid this:

“I theorized that heating would char the resin into a carbonaceous 

mass of relatively low strength. The role of the fibers should be to hold 

the carbonaceous char to virgin, unheated substrate. Here, low thermal 

conductivity was essential to minimize the distance from the hot, exposed 

surface to the cool substrate, to minimize the mass of material that had to 

be held by the fibers as well as the degradation of the fibers. The char itself 

would eventually either be vaporized or be oxidized either by boundary 

layer oxygen or by CO2

 in the boundary layer. The fibers would either melt 

or also vaporize. The question was how to fabricate the material so that the 

fibers interlocked the resin, which was the opposite design philosophy to 

existing laminates in which the resin interlocks the fibers. I believed that a 

solution might be the use of short fibers, randomly oriented in a soup of 

resin, which was then molded into the desired shape. I then began to plan 

the experiments to test this hypothesis.”49

Sutton had no pipeline to Huntsville, but his plan echoed that of Von Braun. 

He proceeded to fabricate small model nose cones from candidate fiber-reinforced 

plastics, planning to test them by immersion in the exhaust of a rocket engine. GE 

was developing an engine for the first stage of the Vanguard program; prototypes 

were at hand, along with test stands. Sutton arranged for an engine to produce an 

exhaust that contained free oxygen to achieve oxidation of the carbon-rich char.

He used two resins along with five types of fiber reinforcement. The best performance came with the use of Refrasil reinforcement, a silicon-dioxide fiber. Both 

resins yielded composites with a heat capacity of 6,300 BTU per pound or greater. 

This was astonishing. The materials had a density of 1.6 times that of water. Yet they 

absorbed more than six times as much heat, pound for pound, as boiling water!50

Jupiter missile with ablative nose cone. (U.S. Army)
40

Facing the Heat Barrier: A History of Hypersonics

41

Nose Cones and Re-entry

Here was a new form of thermal protection: ablation. An ablative 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 flux. Ablation 

promised all the advantages of transpiration cooling, within a system that could be 

considerably lighter and yet more capable.51

Sutton presented his experimental results in June 1957 at a technical conference 

held at the firm of Ramo-Wooldridge in Los Angeles. This company was providing 

technical support to the Air Force’s Atlas program management. Following this talk, 

George Solomon, one of that firm’s leading scientists, rose to his feet and stated that 

ablation was the solution to the problem of thermal protection.

The Army thought so too. It had invented ablation on its own, considerably earlier and amid far deeper investigation. Indeed, at the moment when Sutton gave his 

talk, Von Braun was only two months away from a successful flight test of a subscale 

nose cone. People might argue whether the Soviets were ahead of the United States 

in missiles, but there was no doubt that the Army was ahead of the Air Force in nose 

cones. Jupiter was already slated for an ablative cone, but Thor was to use heat sink, 

as was the intercontinental Atlas.

Already, though, new information was available concerning transition from laminar to turbulent flow over a nose cone. Turbulent heating would be far more severe, 

and these findings showed that copper, the best heat-sink material, was inadequate 

for an ICBM. Materials testing now came to the forefront, and this work needed 

new facilities. A rocket-engine exhaust could reproduce the rate of heat transfer, 

but in Kantrowitz’s words, “a rocket is not hot enough.”52 It could not duplicate the 

temperatures of re-entry.

A shock tube indeed gave a suitably hot flow, but its duration of less than a 

millisecond was hopelessly inadequate for testing ablative materials. Investigators 

needed a new type of wind tunnel that could produce a continuous flow, but at 

temperatures far greater than were available. Fortunately, such an installation did 

not have to reproduce the hypersonic Mach numbers of re-entry; it sufficed to 

duplicate the necessary temperatures within the flow. The instrument that did this 

was the arc tunnel.

It heated the air with an electric arc, which amounted to a man-made stroke of 

lightning. Such arcs were in routine use in welding; Avco’s Thomas Brogan noted 

that they reached 6500 K, “a temperature which would exist at the [tip] of a blunt 

body flying at 22,000 feet per second.” In seeking to develop an arc-heated wind 

tunnel, a point of departure lay in West Germany, where researchers had built a 

“plasma jet.”53

This device swirled water around a long carbon rod that served as the cathode. 

The motion of the water helped to keep the arc focused on the anode, which was 

also of carbon and which held a small nozzle. The arc produced its plasma as a mix 

of very hot steam and carbon vapor, which was ejected through the nozzle. This 

invention achieved pressures of 50 atmospheres, with the plasma temperature at the 

nozzle exit being measured at 8000 K. The carbon cathode eroded relatively slowly, 

while the water supply was easily refilled. The plasma jet therefore could operate for 

fairly long times.54

At NACA-Langley, an experimental arc tunnel went into operation in May 1957. 

It differed from the German plasma jet by using an electric arc to heat a flow of air, 

nitrogen, or helium. With a test section measuring only seven millimeters square, it 

was a proof-of-principle instrument rather than a working facility. Still, its plasma 

temperatures ranged from 5800 to 7000 K, which was well beyond the reach of a 

conventional hypersonic wind tunnel.55

At Avco, Kantrowitz paid attention when he heard the word “plasma.” He 

had been studying such ionized gases ever since he had tried to invent controlled 

fusion. His first arc tunnel was rated only at 130 kilowatts, a limited power level 

that restricted the simulated altitude to between 165,000 and 210,000 feet. Its hot 

plasma flowed from its nozzle at Mach 3.4, but when this flow came to a stop when 

impinging on samples of quartz, the temperature corresponded to flight velocities 

as high as 21,000 feet per second. Tests showed good agreement between theory 

and experiment, with measured surface temperatures of 2700 K falling within three 

percent of calculated values. The investigators concluded that opaque quartz “will 

effectively absorb about 4000 BTU per pound for ICBM and [intermediate-range] 

trajectories.”56

In Huntsville, Von Braun’s colleagues found their way as well to the arc tunnel. 

They also learned of the initial work in Germany. In addition, the small California 

firm of Plasmadyne acquired such a device and then performed experiments under 

contract to the Army. In 1958 Rolf Buhler, a company scientist, discovered that 

when he placed a blunt rod of graphite in the flow, the rod became pointed. Other 

investigators attributed this result to the presence of a cool core in the arc-heated jet, 

but Sutton succeeded in deriving this observed shape from theory.

This immediately raised the prospect of nose cones that after all might be sharply 

pointed rather than blunt. Such re-entry bodies would not slow down in the upper 

atmosphere, perhaps making themselves tempting targets for antiballistic missiles, 

but would continue to fall rapidly. Graphite still had the inconvenient features noted 

previously, but a new material, pyrolytic graphite, promised to ease the problem of 

its high thermal conductivity.

Pyrolytic graphite was made by chemical vapor deposition. One placed a temperature-resistant form in an atmosphere of gaseous hydrocarbons. The hot surface 

broke up the gas molecules, a process known as pyrolysis, and left carbon on the surface. The thermal conductivity then was considerably lower in a direction normal 

to the surface than when parallel to it. The low value of this conductivity, in the 

normal direction, made such graphite attractive.57
42

Facing the Heat Barrier: A History of Hypersonics

43

Nose Cones and Re-entry

Having whetted their appetites with the 130-kilowatt facility, Avco went on to 

build one that was two orders of magnitude more powerful. It used a 15-megawatt 

power supply and obtained this from a bank of 2,000 twelve-volt truck batteries, 

with motor-generators to charge them. They provided direct current for run times 

of up to a minute and could be recharged in an hour.58

With this, Avco added the high-power arc tunnel to the existing array of hypersonic flow facilities. These included aerodynamic wind tunnels such as Becker’s, 

along with plasma jets and shock tubes. And while the array of ground installations 

proliferated, the ICBM program was moving toward a different kind of test: fullscale flight.

Flight Test

The first important step in this direction came in January 1955, when the Air 

Force issued a letter contract to Lockheed that authorized them to proceed with the 

X-17. It took shape as a three-stage missile, with all three stages using solid-propellant rocket motors from Thiokol. It was to reach Mach 15, and it used a new flight 

mode called “over the top.”

The X-17 was not to fire all three stages to achieve a very high ballistic trajectory. Instead it started with only its first stage, climbing to an altitude of 65 to 100 

miles. Descent from such an altitude imposed no serious thermal problems. As it 

re-entered the atmosphere, large fins took hold and pointed it downward. Below 

100,000 feet, the two upper stages fired, again while pointing downward. These 

stages accelerated a test nose cone to maximum speed, deep within the atmosphere. 

This technique prevented the nose cone from decelerating at high altitude, which 

would have happened with a very high ballistic flight path. Over-the-top also gave 

good control of both the peak Mach and of its altitude of attainment.

The accompanying table summarizes the results. Following a succession of subscale and developmental flights that ran from 1955 into 1956, the program conducted two dozen test firings in only eight months. The start was somewhat shaky 

as no more than two of the first six X-17s gained full success, but the program 

soon settled down to routine achievement. The simplicity of solid-propellant rocketry enabled the flights to proceed with turnaround times of as little as four days. 

Launches required no more than 40 active personnel, with as many as five such 

flights taking place within the single month of October 1956. All of them flew from 

a single facility: Pad 3 at Cape Canaveral.59

X-17 FLIGHT TESTS

Date Nose-Cone Shape Results

17 Jul 1956 Hemisphere Mach 12.4 at 40,000 feet.

27 Jul 1956 Cubic Paraboloid Third stage failed to ignite.

18 Aug 1956 Hemisphere Missile exploded 18 sec. after launch.

23 Aug 1956 Blunt Mach 12.4 at 38,500 feet.

28 Aug 1956 Blunt Telemetry lost prior to apogee.

8 Sep 1956 Cubic Paraboloid Upper stages ignited while ascending.

1 Oct 1956 Hemisphere Mach 12.1 at 36,500 feet.

5 Oct 1956 Hemisphere Mach 13.7 at 54,000 feet.

13 Oct 1956 Cubic Paraboloid Mach 13.8 at 58,500 feet.

18 Oct 1956 Hemisphere Mach 12.6 at 37,000 feet.

25 Oct 1956 Blunt Mach 14.2 at 59,000 feet.

5 Nov 1956 Blunt (Avco) Mach 12.6 at 41,100 feet.

16 Nov 1956 Blunt (Avco) Mach 13.8 at 57,000 feet.

23 Nov 1956 Blunt (Avco) Mach 11.3 at 34,100 feet.

3 Dec 1956 Blunt (Avco) Mach 13.8 at 47,700 feet.

11 Dec 1956 Blunt Cone (GE) Mach 11.4 at 34,000 feet.

8 Jan 1957 Blunt Cone (GE) Mach 11.5 at 34,600 feet.

15 Jan 1957 Blunt Cone (GE) Upper stages failed to ignite.

29 Jan 1957 Blunt Cone (GE) Missile destroyed by Range Safety.

7 Feb 1957 Blunt Cone (GE) Mach 14.4 at 57,000 feet.

14 Feb 1957 Hemisphere Mach 12.1 at 35,000 feet.

1 Mar 1957 Blunt Cone (GE) Mach 11.4 at 35,600 feet.

11 Mar 1957 Blunt (Avco) Mach 11.3 at 35,500 feet.

21 Mar 1957 Blunt (Avco) Mach 13.2 at 54,500 feet.

Source: “Re-Entry Test Vehicle X-17,” pp. 30, 32.

Many nose cones approached or topped Mach 12 at altitudes below 40,000 feet. 

This was half the speed of a satellite, at altitudes where airliners fly today. One places 

this in perspective by noting that the SR-71 cruised above Mach 3, one-fourth this 

speed, and at 85,000 feet, which was more than twice as high. Thermal problems 

dominated its design, with this spy plane being built as a titanium hot structure. 

The X-15 reached Mach 6.7 in 1967, half the speed of an X-17 nose cone, and at 
44

Facing the Heat Barrier: A History of Hypersonics

45

Nose Cones and Re-entry

102,000 feet. Its structure was Inconel X heat sink, and it had further protection 

from a spray-on ablative. Yet it sustained significant physical damage due to high 

temperatures and never again approached that mark.60

Another noteworthy flight involved a five-stage NACA rocket that was to accomplish its own over-the-top mission. It was climbing gently at 96,000 feet when the 

third stage ignited. Telemetry continued for an additional 8.2 seconds and then 

suddenly cut off, with the fifth stage still having half a second to burn. The speed 

was Mach 15.5 at 98,500 feet. The temperature on the inner surface of the skin was 

2,500oF, close to the melting point, with this temperature rising at nearly 5,300oF 

per second.61

How then did X-17 nose cones survive flight at nearly this speed, but at little 

more than one-third the altitude? They did not. They burned up in the atmosphere. 

They lacked thermal protection, whether heat sink or ablative (which the Air Force, 

the X-17’s sponsor, had not invented yet), and no attempt was made to recover 

them. The second and third stages ignited and burned to depletion in only 3.7 

seconds, with the thrust of these stages being 102,000 and 36,000 pounds, respectively.62 Acceleration therefore was extremely rapid; exposure to conditions of very 

high Mach was correspondingly brief. The X-17 thus amounted to a flying shock 

tube. Its nose cones lived only long enough to return data; then they vanished into 

thin air.

Yet these data were priceless. They included measurements of boundary-layer 

transition, heat transfer, and pressure distributions, covering a broad range of peak 

Mach values, altitudes, and nose-cone shapes. The information from this program 

complemented the data from Avco Research Laboratory, contributing materially to 

Air Force decisions that selected ablation for Atlas (and for Titan, a second ICBM), 

while retaining heat sink for Thor.63

As the X-17 went forward during 1956 and 1957, the Army weighed in with its 

own flight-test effort. Here were no over-the-top heroics, no ultrashort moments 

at high Mach with nose cones built to do their duty and die. The Army wanted 

nothing less than complete tests of true ablating nose cones, initially at subscale and 

later at full scale, along realistic ballistic trajectories. The nose cones were to survive 

re-entry. If possible, they were to be recovered from the sea.

The launch vehicle was the Jupiter-C, another product of Von Braun. It was 

based on the liquid-fueled Redstone missile, which was fitted with longer propellant 

tanks to extend the burning time. Atop that missile rode two additional stages, both 

of which were built as clusters of small solid-fuel rockets.

The first flight took place from Cape Canaveral in September 1956. It carried 

no nose cone; this launch had the purpose of verifying the three-stage design, particularly its methods for stage separation and ignition. A dummy solid rocket rode 

atop this stack as a payload. All three stages fired successfully, and the flight broke all 

performance records. The payload reached a peak altitude of 682 miles and attained 

an estimated range of 3,335 miles.64

Nose-cone tests followed during 1957. Each cone largely duplicated that of the 

Jupiter missile but was less than one-third the size, having a length of 29 inches 

and maximum diameter of 20 inches. The weight was 314 pounds, of which 83 

Thor missile with heat-sink nose cone. (U.S. Air Force)
46

Facing the Heat Barrier: A History of Hypersonics

47

Nose Cones and Re-entry

pounds constituted the mix of glass cloth and Micarta plastic that formed the ablative material. To aid in recovery in the ocean, each nose cone came equipped with a 

balloon for flotation, two small bombs to indicate position for sonar, a dye marker, 

a beacon light, a radio transmitter—and shark repellant, to protect the balloon from 

attack.65

The first nose-cone flight took place in May. Telemetry showed that the re-entry 

vehicle came through the atmosphere successfully and that the ablative thermal protection indeed had worked. However, a faulty trajectory caused this nose cone to fall 

480 miles short of the planned impact point, and this payload was not recovered.

Full success came with the next launch, in August. All three stages again fired, 

pushing the nose cone to a range of 1,343 statute miles. This was shorter than the 

planned range of Jupiter, 1,725 miles, but still this payload experienced 95 percent 

of the total heat transfer that it would have received at the tip for a full-range flight. 

The nose cone also was recovered, giving scientists their first close look at one that 

had actually survived.66

In November President Eisenhower personally displayed it to the nation. The 

Soviets had stirred considerable concern by placing two Sputnik satellites in orbit, 

thus showing that they already had an ICBM. Speaking on nationwide radio and 

television, Ike sought to reassure the public. He spoke of American long-range 

bombers and then presented his jewel: “One difficult obstacle on the way to producing a useful long-range weapon is that of bringing a missile back from outer 

space without its burning up like a meteor. This object here in my office is the nose 

cone of an experimental missile. It has been hundreds of miles into outer space and 

back. Here it is, completely intact.”67

Jupiter then was in flight test and became the first missile to carry a full-size nose 

cone to full range.68 But the range of Jupiter was far shorter than that of Atlas. The 

Army had taken an initial lead in nose-cone testing by taking advantage of its early 

start, but by the time of that flight—May 1958—all eyes were on the Air Force and 

on flight to intercontinental range.

Atlas also was in flight test during 1958, extending its range in small steps, but it 

still was far from ready to serve as a test vehicle for nose cones. To attain 5,000-mile 

range, Air Force officials added an upper stage to the Thor. The resulting rocket, the 

Thor-Able, indeed had the job of testing nose cones. An early model, from General 

Electric, weighed more than 600 pounds and carried 700 pounds of instruments.69

Two successful flights, both to full range, took place during July 1958. The 

first one reached a peak altitude of 1,400 miles and flew 5,500 miles to the South 

Atlantic. Telemetered data showed that its re-entry vehicle survived the fiery passage through the atmosphere, while withstanding four times the heat load of a Thor 

heat-sink nose cone. This flight carried a passenger, a mouse named Laska in honor 

of what soon became the 49th state. Little Laska lived through decelerations during 

re-entry that reached 60 g, due to the steepness of the trajectory, but the nose cone 

was not recovered and sank into the sea. Much the same happened two weeks later, 

with the mouse being named Wickie. Again the reentry vehicle came through the 

atmosphere successfully, but Wickie died for his country as well, for this nose cone 

also sank without being recovered.70

A new series of tests went forward during 1959, as General Electric introduced 

the RVX-1 vehicle. Weighing 645 pounds, 67 inches long with a diameter at the 

base of 28 inches, it was a cylinder with a very blunt nose and a conical afterbody for 

stability.71 A flight in March used phenolic nylon as the ablator. This was a phenolic 

resin containing randomly oriented one-inch-square pieces of nylon cloth. Light 

weight was its strong suit; with a density as low as 72 pounds per cubic foot, it was 

only slightly denser than water. It also was highly effective as insulation. Following 

flight to full range, telemetered data showed that a layer only a quarter-inch thick 

could limit the temperature rise on the aft body, which was strongly heated, to less 

than 200oF. This was well within the permissible range for aluminum, the most 

familiar of aerospace materials. For the nose cap, where the heating was strongest, 

GE installed a thick coating of molded phenolic nylon.72

Within this new series of flights, new guidance promised enhanced accuracy and 

a better chance of retrieval. Still, that March flight was not recovered, with another 

shot also flying successfully but again sinking beneath the waves. When the first 

recovery indeed took place, it resulted largely from luck.

Early in April an RVX-1 made a flawless flight, soaring to 764 miles in altitude 

and sailing downrange to 4,944 miles. Peak speed during re-entry was Mach 20, 

Jupiter nose cone. (U.S. Army)
48

Facing the Heat Barrier: A History of Hypersonics

49

Nose Cones and Re-entry

or 21,400 feet per second. Peak heating occurred at Mach 16, or 15,000 feet per 

second, and at 60,000 feet. The nose cone took this in stride, but searchers failed to 

detect its radio signals. An Avco man in one of the search planes saved the situation 

by spotting its dye marker. Aircraft then orbited the position for three hours until a 

recovery vessel arrived and picked it up.73

It was the first vehicle to fly to intercontinental range and return for inspection. 

Avco had specified its design, using an ablative heat shield of fused opaque quartz. 

Inspection of the ablated surface permitted comparison with theory, and the results 

were described as giving “excellent agreement.” The observed value of maximum 

ablated thickness was 9 percent higher than the theoretical value. The weight loss 

of ablated material agreed within 20 percent, while the fraction of ablated material 

that vaporized during re-entry was only 3 percent higher than the theoretical value. 

Most of the differences could be explained by the effect of impurities on the viscosity of opaque quartz.74

A second complete success was achieved six weeks later, again with a range 

of 5,000 miles. Observers aboard a C-54 search aircraft witnessed the re-entry, 

acquired the radio beacon, and then guided a recovery ship to the site.75 This time 

the nose-cone design came from GE. That company’s project engineer, Walter Schafer, wanted to try several materials and to instrument them with breakwire sensors. 

These were wires, buried at various depths within the ablative material, that would 

break as it eroded away and thus disclose the rate of ablation. GE followed a suggestion from George Sutton and installed each material as a 60-degree segment around 

the cylinder and afterbody, with the same material being repeated every 180 degrees 

for symmetry.76

Within the fast-paced world of nose-cone studies, each year had brought at least 

one new flight vehicle. The X-17 had flown during 1956. For the Jupiter-C, success 

had come in 1957. The year 1958 brought both Jupiter and the Thor-Able. Now, 

in 1959, the nose-cone program was to gain final success by flying full-size re-entry 

vehicles to full range aboard Atlas.

The program had laid important groundwork in November 1958, when this 

missile first flew to intercontinental distance. The test conductor, with the hopeful 

name of Bob Shotwell, pushed the button and the rocket leaped into the night. It 

traced an arc above the Moon as it flew across the starry sky. It dropped its twin 

booster engines; then, continuing to accelerate, the brilliant light of its main engine 

faded. Now it seemed to hang in the darkness like a new star, just below Orion. 

Shotwell and his crew contained their enthusiasm for a full seven minutes; then they 

erupted in shouts. They had it; the missile was soaring at 16,000 miles per hour, 

bound for a spot near the island of St. Helena in the South Atlantic, a full 6,300 

miles from the Cape. In Shotwell’s words, “We knew we had done it. It was going 

like a bullet; nothing could stop it.”77

Atlas could carry far heavier 

loads than Thor-Able, and its 

first nose cone reflected this. It 

was the RVX-2, again from General Electric, which had the shape 

of a long cone with a round tip. 

With a length of 147 inches and 

a width at the base of 64 inches, 

it weighed some 2,500 pounds. 

Once more, phenolic nylon was 

used for thermal protection. It 

flew to a range of 5,047 miles in 

July 1959 and was recovered. It 

thereby became the largest object 

to have been brought back following re-entry.78

Attention now turned to 

developmental tests of a nose 

cone for the operational Atlas. 

This was the Mark 3, also from 

GE. Its design returned to the 

basic RVX-1 configuration, 

again with a blunt nose at the 

front of a cylinder but with 

a longer conical afterbody. It was slightly smaller than the RVX-2, with a length 

of 115 inches, diameter at the cylinder of 21 inches, and diameter at the base of 

36 inches. Phenolic nylon was specified throughout for thermal protection, being 

molded under high pressure for the nose cap and tape-wound on the cylinder and 

afterbody. The Mark 3 weighed 2,140 pounds, making it somewhat lighter than 

the RVX-2. The low density of phenolic nylon showed itself anew, for of this total 

weight, only 308 pounds constituted ablative material.79

The Mark 3 launches began in October 1959 and ran for several months, with 

this nose cone entering operational service the following April.80 The flights again 

were full -range, with one of them flying 5,000 miles to Ascension Island and another 

going 6,300 miles. Re-entry speeds went as high as 22,500 feet per second. Peak 

heat transfer occurred near Mach 14 and 40,000 feet in altitude, approximating 

the conditions of the X-17 tests. The air at that height was too thin to breathe, but 

the nose cone set up a shock wave that compressed the incoming flow, producing a 

wind resistance with dynamic pressure of more than 30 atmospheres. Temperatures 

at the nose reached 6,500oF.81

Nose cones used in flight test. Top, RVX-1; bottom, RVX2. (U.S. Air Force)
50

Facing the Heat Barrier: A History of Hypersonics

51

Nose Cones and Re-entry

Each re-entry vehicle was extensively instrumented, mounting nearly two dozen 

breakwire ablation sensors along with pressure and temperature sensors. The latter 

were resistance thermometers employing 0.0003-inch tungsten wire, reporting temperatures to 2000oF with an accuracy of 25 to 50°F. The phenolic nylon showed 

anew that it had the right stuff, for it absorbed heat at the rate of 3,000 BTU per 

pound, making it three times as effective as boiling water. A report from GE noted, 

“all temperature sensors located on the cylindrical section were at locations too far 

below the initial surface to register a temperature rise.”82

With this, the main effort in re-entry reached completion, and its solution—

ablation—had proved to be relatively simple. The process resembled the charring of 

wood. Indeed, Kantrowitz recalls Von Braun suggesting that it was possible to build 

a nose cone of lightweight balsa soaked in water and frozen. In Kantrowitz’s words, 

“That might be a very reasonable ablator.”83

Experience with ablation also contrasted in welcome fashion with a strong tendency of advanced technologies to rely on highly specialized materials. Nuclear 

energy used uranium-235, which called for the enormous difficulty of isotope 

separation, along with plutonium, which had to be produced in a nuclear reactor and then be extracted from highly radioactive spent fuel. Solid-state electronics 

depended on silicon or germanium, but while silicon was common, either element 

demanded refinement to exquisite levels of purity.

Ablation was different. Although wood proved inappropriate, once the basic 

concept was in hand the problem became one of choosing the best candidate from 

a surprisingly wide variety of possibilities. These generally were commercial plastics that served as binders, with the main heat resistance being provided by glass 

or silica. Quartz also worked well, particularly after being rendered opaque, while 

pyrolytic graphite exemplified a new material with novel properties.

The physicist Steven Weinberg, winner of a Nobel Prize, stated that a researcher 

never knows how difficult a problem is until the solution is in hand. In 1956 Theodore von Karman had described re-entry as “perhaps one of the most difficult problems one can imagine. It is certainly a problem that constitutes a challenge to the 

best brains working in these domains of modern aerophysics.”84 Yet in the end, amid 

all the ingenuity of shock tubes and arc tunnels, the fundamental insights derived 

from nothing deeper than testing an assortment of candidate materials in the blast 

of rocket engines.

1 Vannevar Bush, testimony before the Senate Special Committee on Atomic Energy, December 

1945, quoted in Clarke, Profiles, p. 9 and Time, 22 May 1964, p. 25.

2 MacKenzie, Inventing, pp. 113-14. 1,500 feet: Neufeld, Ballistic, pp. 69-102 passim.

3 North American Aviation Report AL-1347, pp. 1-6 (quote, p. 4); Fahrney, History, p. 1291, 

footnote 3.

4 North American Aviation Report AL-1347, pp. 38-39; Colonel Roth: Letter, Colonel M. S. 

Roth to Power Plant Lab, Air Materiel Command, USAF, 11 February 1948.

5 Author interview, Colonel Edward Hall 29 August 1996, Folder 18649, NASA Historical 

Reference Collection, NASA History Division, Washington, D.C. 20546.

6 North American Aviation Report AL-1347, p. 39.

7 Other North American Aviation names of that era included the NAKA, NALAR, and NASTY 

battlefield missiles, the Navion private plane, and the NATIV research rocket; see “Thirty Years 

of Rocketdyne,” photo, “Armament Rockets,” 1962; Murray, Atwood, pp. 32, 44.

8 Rhodes, Dark Sun, p. 407 (includes quote); Manchester, Caesar, p. 642.

9 “Standard Missile Characteristics: XSM-64 Navaho”; Report AL-1347, op. cit., p. 88; Development of the Navaho, pp. 30-31; Augenstein, “Rand”; Fahrney, History, pp. 1296-98. General 

Putt: Letter, Major General Donald Putt to Commanding General, Air Materiel Command, 

USAF, 21 August 1950.

10 Neufeld, Ballistic, pp. 44-50, 68-70.

11 Ibid., pp. 70-79.

12 Development of the Navaho, pp. 40-46; Journal of Guidance and Control, 

September-October 1981, pp. 455-57.

13 Neufeld, Ballistic, pp. 73, 77-78 (quote, p. 78).

14 Rhodes, Dark Sun, pp. 495, 499, 505 (quotes, p. 509). Taylor quote: McPhee, Curve, p. 77. 

Hiroshima: Rhodes, Atomic, p. 711.

15 Von Karman and Edson, Wind, pp. 300-01; Neufeld, Ballistic, p. 98.

16 Neufeld, Ballistic, pp. 95, 98-99, 102; quote: Chapman, Atlas, p. 73.

17 AIAA Paper 67-838, pp. 12-13; Neufeld, Ballistic, p. 102 (quotes, p. 259).

18 Rhodes, Dark Sun, pp. 482-84, 541-42 (quote, p. 541).

19 Neufeld, Ballistic, pp. 104-06, 117 (quotes, pp. 105, 106); Rhodes, Dark Sun, p. 542.

20 Neufeld, Ballistic, p. 117; Emme, History, p. 151. Dimensions: Ley, Rockets, p. 400.

21 Miller, X-Planes, ch. 7; letter, Smith DeFrance to NACA Headquarters, 26 November 1952.

22 NACA Report 1381 (quotes, pp. 11, 13).

23 Anderson, History, pp. 440-41.

24 Neufeld, Ballistic, p. 79.

25 Time, 13 June 1960, p. 70. 9,000 K: Journal of the Aeronautical Sciences, February 1958, p. 88. 

Quote: Rose, “Physical,” p. 1.

26 Rose, “Physical,” p. 1.

27 Hallion, Hypersonic, pp. xxiii, xxvi.

28 Time, 13 June 1960 (includes quote). Arthur Kantrowitz: Author interview, 22 June 2001. 
52

Facing the Heat Barrier: A History of Hypersonics

53

Nose Cones and Re-entry

Folder 18649, NASA Historical Reference Collection, NASA History Division, Washington, 

D.C. 20546.

29 Heppenheimer, Man-Made, pp. 286-91. Quote: Author interview, Arthur Kantrowitz, 22 June 

2001. Folder 18649, NASA Historical Reference Collection, NASA History Division, Washington, D.C. 20546.

30 Shapiro, Compressible, pp. 1007-11, 1027; Lukasiewicz, Experimental, ch. 13.

31 Comptes Rendus, Vol. 129 (1899), pp. 1128-30.

32 Proceedings of the Royal Society, Vol. A186 (1946), pp. 293-21.

33 Review of Scientific Instruments, November 1949, pp. 807-815 (quotes, pp. 813, 814).

34 Journal of Applied Physics, December 1952, pp. 1390-99 (quote, p. 1397).

35 Wegener, Peenemunde, p. 131; Nature, Vol. 171 (1953), pp. 395-96; NASA SP-4308, p. 134; 

quote: Hallion, Hypersonic, pp. xxvi-xxvii.

36 Rose, “Physical”; Kantrowitz, “Shock Tubes.”

37 Journal of the Aeronautical Sciences, February 1958, pp. 86-97. 

38 Rose, “Physical,” pp. 11-12.

39 Jet Propulsion, April 1956, pp. 259-269.

40 Journal of the Aeronautical Sciences, February 1958: pp. 73-85, 86-97. 

41 Brown et al., “Study,” pp. 13-17.

42 Neufeld, Ballistic, pp. 78-79, 117.

43 Emme, History, pp. 7-110; Ordway and Sharpe, Rocket Team, pp. 366-73.

44 Ley, Rockets, pp. 237, 245.

45 “Re-Entry Studies,” Vol. 1, pp. 21, 27, 29.

46 Grimwood and Strowd, History, pp. 10-13.

47 “Re-Entry Studies,” Vol. 1, pp. 2, 24-25, 31, 37-45, 61.

48 Journal of Spacecraft and Rockets, January-February 1982, pp. 3-11.

49 Ibid. Extended quote, p. 5.

50 Journal of the Aero/Space Sciences, May 1960, pp. 377-85. 

51 Kreith, Heat Transfer, pp. 538-45.

52 Journal of Spacecraft and Rockets, January-February 1982, p. 6. Quote: Author interview, Arthur 

Kantrowitz, 22 June 2001. Folder 18649, NASA Historical Reference Collection, NASA History Division, Washington, D.C. 20546.

53 NASA SP-440, p. 85; ARS Paper 724-58 (quote, p. 4).

54 ARS Paper 724-58.

55 NASA SP-4308, p. 133.

56 ARS Paper 724-58; Journal of the Aero/Space Sciences, July 1960, pp. 535-43 (quote, p. 535). 

57 Journal of Spacecraft and Rockets, January-February 1982, pp. 8-9.

58 ARS Paper 838-59.

59 “Re-Entry Test Vehicle X-17.” Over-the-top: NASA RP-1028, p. 443.

60 This flight is further discussed in ch. 6.

61 NASA RP-1028, p. 488.

62 Miller, X-Planes, pp. 215, 217.

63 Ibid., p. 213; Journal of Spacecraft and Rockets, January-February 1982, p. 6.

64 “Re-Entry Studies,” Vol. 1, pp. 69-71; Grimwood and Strowd, History, p. 156.

65 “Re-Entry Studies,” Vol. 1, pp. 61, 64, 71, 73, 98.

66 Grimwood and Strowd, History, p. 156.

67 Time, 18 November 1957, pp. 19-20 (includes quote).

68 “Re-Entry Studies,” Vol. 1, pp. 3, 5.

69 Journal of the British Interplanetary Society, May 1984, pp. 219-25; Hallion, Hypersonic, p. lxxii. 

70 Journal of the British Interplanetary Society, May 1984, p. 219. 

71 DTIC AD-320753, pp. 11-12.

72 DTIC AD-342828, pp. 16-19; AD-362539, pp. 11, 83.

73 Journal of the British Interplanetary Society, May 1984, p. 220. Trajectory data: DTIC AD320753, Appendix A. 

74 AIAA Journal, January 1963, pp. 41-45 (quote, p. 41).

75 Journal of the British Interplanetary Society, May 1984, p. 220. 

76 Journal of the British Interplanetary Society, May 1984, p. 220; Journal of Spacecraft and Rockets, 

January-February 1982, p. 8. 

77 Time, December 8, 1958, p. 15 (includes quote).

78 Hallion, Hypersonic, p. lxxii. Dimensions: DTIC AD-362539, p. 73. Weight: AD-832686, p. 

20-23.

79 DTIC AD-362539, pp. 11, 39.

80 Hallion, Hypersonic, p. lxxii.

81 DTIC AD-362539, pp. 11-12. Peak heating occurred approximately at maximum dynamic 

pressure, at Mach 14 (see p. 12). Altitude, 40,000 feet, is from Figs. 2.6.

82 DTIC AD-362539, pp. 9, 12-13 (quote, p. 9).

83 Author interview, Arthur Kantrowitz, 22 June 2001. Folder 18649, NASA Historical Reference 

Collection, NASA History Division, Washington, D.C. 20546

84 Author interview, Steven Weinberg, 3 January 1990. Folder 18649, NASA Historical Reference 

Collection, NASA History Division, Washington, D.C. 20546. Quote: Journal of Spacecraft 

and Rockets, January-February 1982, p. 3.

Across almost half a century, the X-15 program stands out to this day not only 

for its achievements but for its audacity. At a time when the speed record stood right 

at Mach 2, the creators of the X-15 aimed for Mach 7—and nearly did it.* Moreover, the accomplishments of the X-15 contrast with the history of an X-planes 

program that saw the X-1A and X-2 fall out of the sky due to flight instabilities, and 

in which the X-3 fell short in speed because it was underpowered.1

The X-15 is all the more remarkable because its only significant source of aerodynamic data was Becker’s 11-inch hypersonic wind tunnel. Based on that instrument alone, the Air Force and NACA set out to challenge the potential difficulties 

of hypersonic piloted flight. They succeeded, with this aircraft setting speed and 

altitude marks that were not surpassed until the advent of the space shuttle.

It is true that these agencies worked at a time of rapid advance, when performance was leaping forward at rates never approached either before or since. Yet 

there was more to this craft than a can-do spirit. Its designers faced specific technical 

issues and overcame them well before the first metal was cut.

The X-3 had failed because it proved infeasible to fit it with the powerful turbojet engines that it needed. The X-15 was conceived from the start as relying on 

rocket power, which gave it a very ample reserve.

Flight instability was already recognized as a serious concern. Using Becker’s 

hypersonic tunnel, the aerodynamicist Charles McLellan showed that the effectiveness of tail surfaces could be greatly increased by designing them with wedge-shaped 

profiles.2

The X-15 was built particularly to study problems of heating in high-speed 

flight, and there was the question of whether it might overheat when re-entering 

the atmosphere following a climb to altitude. Calculations showed that the heating 

would remain within acceptable bounds if the airplane re-entered with its nose high. 

This would present its broad underbelly to the oncoming airflow. Here was a new 

application of the Allen-Eggers blunt-body principle, for an airplane with its nose 

up effectively became blunt.

55

The X-15

3

*Official flight records are certified by the Federation Aeronautique Internationale. The cited 

accomplishments lacked this distinction, but they nevertheless represented genuine achievements.
56

Facing the Heat Barrier: A History of Hypersonics

57

The X-15

Hence in 1951 and 1952, it already was too late to initiate a new program aimed 

at building an X-plane that could provide timely support for the Navaho and XF103. The X-10 supported Navaho from 1954 to 1957, but it used turbojets rather 

than ramjets and flew at Mach 2.There was no quick and easy way to build aircraft 

capable of Mach 3, let alone Mach 4; the lagging X-2 was the only airplane that 

might do this, however belatedly. Yet it was already appropriate to look beyond the 

coming Mach 3 generation and to envision putative successors.

Maxwell Hunter, at Douglas Aircraft, argued that with fighter aircraft on their 

way to Mach 3, antiaircraft missiles would have to fly at Mach 5 to Mach 10.7

 In 

addition, Walter Dornberger, the wartime head of Germany’s rocket program, now 

was at Bell Aircraft. He was directing studies of Bomi, Bomber Missile, a twostage fully reusable rocket-powered bomber concept that was to reach 8,450 mph, 

or Mach 12.8

 At Convair, studies of intercontinental missiles included boost-glide 

concepts with much higher speeds.9

 William Dorrance, a company aerodynamicist, 

had not been free to disclose the classified Atlas concept to NACA but nevertheless 

declared that data at speeds up to Mach 20 were urgently needed.10 In addition, the 

Rand Corporation had already published reports that envisioned spacecraft in orbit. 

The documents proposed that such satellites could serve for weather observation 

and for military reconnaissance.11

At Bell Aircraft, Robert Woods, a co-founder of the company, took a strong 

interest in Dornberger’s ideas. Woods had designed the X-1, the X-1A that reached 

Mach 2.4, and the X-2. He also was a member of NACA’s influential Committee on 

Aerodynamics. At a meeting of this committee in October 1951, he recommended 

a feasibility study of a “V-2 research airplane, the objective of which would be to 

obtain data at extreme altitudes and speeds and to explore the problems of re-entry 

into the atmosphere.”12 He reiterated this recommendation in a letter to the committee in January 1952. Later that month, he received a memo from Dornberger 

that outlined an “ionospheric research plane,” capable of reaching altitudes of “more 

than 75 miles.”13

NACA Headquarters sent copies of these documents to its field centers. This 

brought responses during May, as several investigators suggested means to enhance 

the performance of the X-2. The proposals included a rocket-powered carrier aircraft with which this research airplane was to attain “Mach numbers up to almost 

10 and an altitude of about 1,000,000 feet,”14 which the X-2 had certainly never 

been meant to attain. A slightly more practical concept called for flight to 300,000 

feet.15 These thoughts were out in the wild blue, but they showed that people at least 

were ready to think about hypersonic flight.

Accordingly, at a meeting in June 1952, the Committee on Aerodynamics 

adopted a resolution largely in a form written by another of its members, the Air 

Force science advisor Albert Lombard:

The plane’s designers also benefited from a stroke of serendipity. Like any airplane, the X-15 was to reduce its weight by using stressed-skin construction; its 

outer skin was to share structural loads with internal bracing. Knowing the stresses 

this craft would encounter, the designers produced straightforward calculations to 

give the requisite skin gauges. A separate set of calculations gave the skin thicknesses 

that were required for the craft to absorb its heat of re-entry without weakening. 

The two sets of skin gauges were nearly the same! This meant that the skin could 

do double duty, bearing stress while absorbing heat. It would not have to thicken 

excessively, thus adding weight, to cope with the heat.

Yet for all the ingenuity that went into this preliminary design, NACA was a very 

small tail on a very large dog in those days, and the dog was the Air Force. NACA 

alone lacked the clout to build anything, which is why one sees military insignia 

on photos of the X-planes of that era. Fortuitously, two new inventions—the twinspool and the variable-stator turbojet—were bringing the Air Force face to face with 

a new era in flight speed. Ramjet engines also were in development, promising still 

higher speed. The X-15 thus stood to provide flight-test data of the highest importance—and the Air Force grabbed the concept and turned it into reality.

Origins of the X-15

Experimental aircraft flourished during the postwar years, but it was hard for 

them to keep pace with the best jet fighters. The X-1, for instance, was the first 

piloted aircraft to break the sound barrier. But only six months later, in April 1948, 

the test pilot George Welch did this in a fighter plane, the XP-86.3

 The layout of the 

XP-86 was more advanced, for it used a swept wing whereas the X-1 used a simple 

straight wing. Moreover, while the X-1 was a highly specialized research airplane, 

the XP-86 was a prototype of an operational fighter.

Much the same happened at Mach 2. The test pilot Scott Crossfield was the 

first to reach this mark, flying the experimental Douglas Skyrocket in November 

1953.4

 Just then, Alexander Kartveli of Republic Aviation was well along in crafting 

the XF-105. The Air Force had ordered 37 of them in March 1953. It first flew in 

December 1955; in June 1956 an F-105 reached Mach 2.15. It too was an operational fighter, in contrast to the Skyrocket of two and a half years earlier.

Ramjet-powered craft were to do even better. Navaho was to fly near Mach 3. An 

even more far-reaching prospect was in view at that same Republic Aviation, where 

Kartveli was working on the XF-103. It was to fly at Mach 3.7 with its own ramjet, 

nearly 2,500 miles per hour (mph), with a sustained ceiling of 75,000 feet.5

Yet it was already clear that such aircraft were to go forward in their programs 

without benefit of research aircraft that could lay groundwork. The Bell X-2 was 

in development as a rocket plane designed to reach Mach 3, but although first 

thoughts of it dated to 1945, the program encountered serious delays. The airplane 

did not so much as fly past Mach 1 until 1956.6
58

Facing the Heat Barrier: A History of Hypersonics

59

The X-15

sibility for trajectories and aerodynamic heating. Maxime Faget addressed issues of 

propulsion. Three other specialists covered the topics of structures and materials, 

piloting, configuration, stability, and control.20

A performance analysis defined a loaded weight of 30,000 pounds. Heavier 

weights did not increase the peak speed by much, whereas smaller concepts showed 

a marked falloff in this speed. Trajectory studies then showed that this vehicle could 

reach a range of speeds, from Mach 5 when taking off from the ground to Mach 

10 if launched atop a rocket-powered first stage. If dropped from a B-52 carrier, it 

would attain Mach 6.3.21

Concurrently with this work, prompted by a statement written by Langley’s 

Robert Gilruth, the Air Force’s Aircraft Panel recommended initiation of a research 

airplane that would reach Mach 5 to 7, along with altitudes of several hundred thousand feet. Becker’s group selected a goal of Mach 7, noting that this would permit 

investigation of “extremely wide ranges of operating and heating conditions.” By 

contrast, a Mach 10 vehicle “would require a much greater expenditure of time and 

effort” and yet “would add little in the fields of stability, control, piloting problems, 

and structural heating.”22

A survey of temperature-resistant superalloys brought selection of Inconel X for 

the primary aircraft structure. This was a proprietary alloy from the firm of International Nickel, comprising 72.5 percent nickel, 15 percent chromium, 1 percent 

columbium, and iron as most of the balance. Its principal constituents all counted 

among the most critical materials used in aircraft construction, being employed in 

small quantities for turbine blades in jet engines. But Inconel X was unmatched in 

temperature resistance, holding most of its strength and stiffness at temperatures as 

high as 1200oF.23

Could a Mach 7 vehicle re-enter the atmosphere without exceeding this temperature limit? Becker’s designers initially considered that during reentry, the airplane should point its nose in the direction of flight. This proved impossible; in 

Becker’s words, “the dynamic pressures quickly exceeded by large margins the limit 

of 1,000 pounds per square foot set by structural considerations, and the heating 

loads became disastrous.”

Becker tried to alleviate these problems by using lift during re-entry. According 

to his calculations, he obtained more lift by raising the nose—and the problem 

became far more manageable. He saw that the solution lay in having the plane enter 

the atmosphere with its nose high, presenting its flat undersurface to the air. It then 

would lose speed in the upper atmosphere, easing both the overheating and the 

aerodynamic pressure. The Allen-Eggers paper had been in print for nearly a year, 

and in Becker’s words, “it became obvious to us that what we were seeing here was a 

new manifestation of H. J. Allen’s ‘blunt-body’ principle. As we increased the angle 

of attack, our configuration in effect became more ‘blunt.’” Allen and Eggers had 

WHEREAS, The upper stratosphere is the important new flight region for 

military aircraft in the next decade and certain guided missiles are already 

under development to fly in the lower portions of this region, and

WHEREAS, Flight in the ionosphere and in satellite orbits in outer space 

has long-term attractiveness to military operations....

RESOLVED, That the NACA Committee on Aerodynamics recommends 

that (1) the NACA increase its program dealing with problems of unmanned 

and manned flight in the upper stratosphere at altitudes between 12 and 50 

miles, and at Mach numbers between 4 and 10, and (2) the NACA devote 

a modest effort to problems associated with unmanned and manned flights 

at altitudes from 50 miles to infinity and at speeds from Mach number 10 

to the velocity of escape from the Earth’s gravity.

Three weeks later, in mid-July, the NACA Executive Committee adopted essentially the same resolution, thus giving it the force of policy.16

Floyd Thompson, associate director of NACA-Langley, responded by setting up 

a three-man study team. Their report came out a year later. It showed strong fascination with boost-glide flight, going so far as to propose a commercial aircraft based on 

a boost-glide Atlas concept that was to match the standard fares of current airliners. 

On the more immediate matter of a high-speed research airplane, this group took 

the concept of a boosted X-2 as a point of departure, suggesting that such a vehicle 

could reach Mach 3.7. Like the million-foot X-2 and the 300,000-foot X-2, this lay 

beyond its thermal limits. Still, this study pointed clearly toward an uprated X-2 as 

the next step.17

The Air Force weighed in with its views in October 1953. A report from the 

Aircraft Panel of its Scientific Advisory Board (SAB) discussed the need for a new 

research airplane of very high performance. The panelists stated that “the time was 

ripe” for such a venture and that its feasibility “should be looked into.”18 With this 

plus the report of the Langley group, the question of such a research plane went on 

the agenda of the next meeting of NACA’s Interlaboratory Research Airplane Panel. 

It took place at NACA Headquarters in Washington in February 1954.

It lasted two days. Most discussions centered on current programs, but the issue 

of a new research plane indeed came up. The participants rejected the concept of 

an uprated X-2, declaring that it would be too small for use in high-speed studies. 

They concluded instead “that provision of an entirely new research airplane is desirable.”19

This decision led quickly to a new round of feasibility studies at each of the 

four NACA centers: Langley, Ames, Lewis, and the High-Speed Flight Station. The 

study conducted at Langley was particularly detailed and furnished much of the 

basis for the eventual design of the X-15. Becker directed the work, taking respon-
60

Facing the Heat Barrier: A History of Hypersonics

61

The X-15

through the clever use of available heat-resistant materials. In Becker’s study, the 

Allen-Eggers blunt-body principle reduced the re-entry heating to a level that Inconel X could accommodate.

The putative airplane still faced difficult issues of stability and control. Early in 

1954 these topics were in the forefront, for the test pilot Chuck Yeager had nearly 

crashed when his X-1A fell out of the sky due to a loss of control at Mach 2.44. This 

problem of high-speed instability reflected the natural instability, at all Mach numbers, of a simple wing-body vehicle that lacked tail surfaces. Such surfaces worked 

well at moderate speeds, like the feathers of an arrow, but lost effectiveness with 

increasing Mach. Yeager’s near-disaster had occurred because he had pushed just 

beyond a speed limit set by such considerations of stability. These considerations 

would be far more severe at Mach 7.26

Another Langley aerodynamicist, Charles McLellan, took up this issue by closely 

examining the airflow around a tail surface at high Mach. He drew on recent experimental results from the Langley 11-inch hypersonic tunnel, involving an airfoil with 

a cross section in the shape of a thin diamond. Analysis had indicated that most of 

the control effectiveness of this airfoil was generated by its forward wedge-shaped 

portion. The aft portion contributed little to its overall effectiveness because the pressures on that part of the surface were lower. Experimental tests had confirmed this.

McLellan now proposed to respond to the problem of hypersonic stability by 

using tail surfaces having airfoils that would be wedge-shaped along their entire 

length. In effect, such a surface would consist of a forward portion extending all the 

way to the rear. Subsequent tests in the 11-inch tunnel confirmed that this solution 

worked. Using standard thin airfoils, the new research plane would have needed tail 

surfaces nearly as large as the wings. The wedge shape, which saw use in the operational X-15, reduced their sizes to those of conventional tails.27

The group’s report, dated April 1954, contemplated flight to altitudes as great as 

350,000 feet, or 66 miles. (The X-15 went to 354,200 feet in 1963.)28 This was well 

above the sensible atmosphere, well into an altitude range where flight would be ballistic. This meant that at that early date, Becker’s study was proposing to accomplish 

piloted flight into space.

The Air Force and High-Speed Flight

This report did not constitute a design. However, it gave good reason to believe 

that such a design indeed was feasible. It also gave a foundation for briefings at 

which supporters of hypersonic flight research could seek to parlay the pertinent 

calculations into a full-blown program that would actually build and fly the new 

research planes. To do this, NACA needed support from the Air Force, which had 

a budget 300 times greater than NACA’s. For FY 1955 the Air Force budget was 

$16.6 billion; NACA’s was $56 million.29

developed their principle for missile nose cones, but it now proved equally useful 

when applied to a hypersonic airplane.24

The use of this principle now placed a structural design concept within reach. 

To address this topic, Norris Dow, the structural analyst, considered the use of a 

heat-sink structure. This was to use Inconel X skin of heavy gauge to absorb the heat 

and spread it through this metal so as to lower its temperature. In addition, the skin 

was to play a structural role. Like other all-metal aircraft, the nascent X-15 was to 

use stressed-skin construction. This gave the skin an optimized thickness so that it 

could carry part of the aerodynamic loads, thus reducing the structural weight.

Dow carried through a design exercise in which he initially ignored the issue of 

heating, laying out a stressed-skin concept built of Inconel X with skin gauges determined only by requirements of mechanical strength and stiffness. A second analysis 

then took note of the heating, calculating new gauges that would allow the skin to 

serve as a heat sink. It was clear that if those gauges were large, adding weight to the 

airplane, then it might be necessary to back off from the Mach 7 goal so as to reduce 

the input heat load, thereby reducing the required thicknesses.

When Dow made the calculations, he received a welcome surprise. He found that 

the weight and thickness of a heat-absorbing structure were nearly the same as those 

of a simple aerodynamic structure! This meant that a hypersonic airplane, designed 

largely from consideration of aerodynamic loads, could provide heat-sink thermal 

protection as a bonus. It could do this with little or no additional weight.25

This, more than anything, was the insight that made the X-15 possible. Designers such as Dow knew all too well that ordinary aircraft aluminum lost strength 

beyond Mach 2, due to aerodynamic heating. Yet if hypersonic flight was to mean 

anything, it meant choosing a goal such as Mach 7 and then reaching this goal 

X-15 skin gauges and design temperatures. Generally, the heaviest gauges were required to meet 

the most severe temperatures. (NASA)
62

Facing the Heat Barrier: A History of Hypersonics

63

The X-15

Fortunately, at that very 

moment the Air Force was face 

to face with two major technical innovations that were upsetting all conventional notions of 

military flight. They faced the 

immediate prospect that aircraft 

would soon be flying at temperatures at which aluminum would 

no longer suffice. The inventions that brought this issue to 

the forefront were the dual-spool 

turbojet and the variable-stator 

turbojet—which call for a digression into technical aspects of jet 

propulsion.

Jet engines have functioned 

at speeds as high as Mach 3.3. 

However, such an engine must 

accelerate to reach that speed and 

must remain operable to provide 

control when decelerating from 

that speed. Engine designers 

face the problem of “compressor 

stall,” which arises because compressors have numerous stages or 

rows of blades and the forward 

stages take in more air than the 

rear stages can accommodate. 

Gerhard Neumann of General 

Electric, who solved this problem, states that when a compressor stalls, the airflow 

pushes forward “with a big bang and the pilot loses all his thrust. It’s violent; we 

often had blades break off during a stall.”

An interim solution came from Pratt & Whitney, as the “twin-spool” engine. 

It separated the front and rear compressor stages into two groups, each of which 

could be made to spin at a proper speed. To do this, each group had its own turbine to provide power. A twin-spool turbojet thus amounted to putting one such 

engine inside another one. It worked; it prevented compressor stall, and it also gave 

high internal pressure that promoted good fuel economy. It thus was selected for 

long-range aircraft, including jet bombers and early commercial jet airliners. It also 

powered a number of fighters.

Gerhard Neumann’s engine for supersonic flight. Top, high performance appeared unattainable 

because when accelerating, the forward compressor stages pulled in more airflow than the rear ones 

could swallow. Center, Neumann approached this problem by working with the stators, stationary 

vanes fitted between successive rows of rotating compressor blades. Bottom, he arranged for stators 

on the front stages to turn, varying their angles to the flow. When set crosswise to the flow, as on 

the right, these variable stators reduced the amount of airflow that their compressor stages would 

pull in. This solved the problem of compressor stall, permitting flight at Mach 2 and higher. (Art 

by Don Dixon and Chris Butler)

Twin-spool turbojet, amounting to two engines in one. 

It avoided compressor stall because its low-pressure 

compressor rotated somewhat slowly during acceleration, and hence pulled in less air. (Art by Don Dixon 

and Chris Butler)
64

Facing the Heat Barrier: A History of Hypersonics

65

The X-15

To place this in perspective, one must note the highly nonuniform manner in 

which the Air Force increased the speed of its best fighters after the war. The advent 

of jet propulsion itself brought a dramatic improvement. The author Tom Wolfe 

notes that “a British jet, the Gloster Meteor, jumped the official world speed record 

from 469 to 606 in a single day.”32 That was an increase of nearly thirty percent, but 

after that, things calmed down. The Korean War-era F-86 could break the sound 

barrier in a dive, but although it was the best fighter in service during that war, it 

definitely counted as subsonic. When the next-generation F-100A flew supersonic 

in level flight in May 1953, the event was worthy of note.33

By then, though, both the F-104 and F-105 were on order and in development. 

A twin-spool engine was already powering the F-100A, while the F-104 was to fly 

with variable stators. At a stroke, then, the Air Force found itself in another great 

leap upward, with speeds that were not to increase by a mere thirty percent but were 

to double.

There was more. There had been much to learn about aerodynamics in crafting 

earlier jets; the swept wing was an important example of the requisite innovations. 

But the new aircraft had continued to use aluminum structures. Still, the F-104 and 

F-105 were among the last aircraft that were to be designed using this metal alone. 

At higher speeds, it would be necessary to use other materials as well.

Other materials were already part of mainstream aviation, even in 1954. The 

Bell X-2 had probably been the first airplane to be built with heat-resistant metals, 

mounting wings of stainless steel on a fuselage of the nickel alloy K Monel. This 

gave it a capability of Mach 3.5. Navaho and the XF-103 were both to be built of 

steel and titanium, while the X-7, a ramjet testbed, was also of steel.34 But all these 

craft were to fly near Mach 3, whereas the X-15 was to reach Mach 7. This meant 

that in an era of accelerating change, the X-15 was plausibly a full generation ahead 

of the most advanced designs that were under development.

The Air Force already had shown its commitment to support flight at high 

speed by building the Arnold Engineering Development Center (AEDC). Its background dated to the closing days of World War II, when leaders in what was then 

the Army Air Forces became aware that Germany had been well ahead of the United 

States in the fields of aerodynamics and jet propulsion. In March 1946, Brigadier 

General H. I. Hodes authorized planning an engineering center that would be the 

Air Force’s own.

This facility was to use plenty of electrical power to run its wind tunnels, and a 

committee selected three possible locations. One was Grand Coulee near Spokane, 

Washington, but was ruled out as being too vulnerable to air attack. The second 

was Arizona’s Colorado River, near Hoover Dam. The third was the hills north of 

Alabama, where the Tennessee Valley Authority had its own hydro dams. Senator 

Kenneth McKellar, the president pro tempore of the Senate and chairman of its 

But the twin-spool was relatively heavy, and there was much interest in avoiding 

compressor stall with a lighter solution. It came from Neumann in the form of the 

“variable-stator” engine. Within an engine’s compressor, one finds rows of whirling 

blades. One also finds “stators,” stationary vanes that receive airflow from those 

blades and direct the air onto the next set of blades. Neumann’s insight was that 

the stators could themselves be adjusted, varied in orientation. At moderate speeds, 

when a compressor was prone to stall, the stators could be set crosswise to the flow, 

blocking it in part. At higher speeds, close to an engine’s peak velocity, the stators 

could turn to present themselves edge-on to the flow. Very little of the airstream 

would be blocked, but the engine could still work as designed.30

The twin-spool approach had demanded nothing less than a complete redesign of 

the entire turbojet. The variable-stator approach was much neater because it merely 

called for modification of the forward stages of the compressor. It first flew as part of 

the Lockheed F-104, which was in development during 1953 and which then flew 

in March 1954. Early versions used engines that did not have variable stators, but 

the F-104A had them by 1958. In May of that year this aircraft reached 1,404 mph, 

setting a new world speed record, and set a similar altitude mark at 91,249 feet.31

The F-104, which used variable stators. (U. S. Air Force)
66

Facing the Heat Barrier: A History of Hypersonics

67

The X-15

Facilities such as Tunnels B and C could indeed attain hypersonic speeds, but 

the temperatures of the flows were just above the condensation point of liquid air. 

There was much interest in achieving far greater temperatures, both to add realism 

at speeds below Mach 10 and to obtain Mach numbers well beyond 10. Beginning 

in 1953, the physicist Daniel Bloxsom used the exploding-wire technique, in which 

a powerful electric pulse vaporizes a thin wire, to produce initial temperatures as 

high as 5900 K.

This brought the advent of a new high-speed flow facility: the hotshot tunnel. It 

resembled the shock tube, for the hot gas was to burst a diaphragm and then reach 

high speeds by expanding through a nozzle. But its run times were considerably 

longer, reaching one-twentieth of a second compared to less than a millisecond for 

the shock tube. The first such instrument, Hotshot 1, had a 16-inch test section 

and entered service early in 1956. In March 1957, the 50-inch Hotshot 2 topped 

“escape velocity.”42

Against this background, the X-15 drew great interest. It was to serve as a fullscale airplane at Mach 7, when the best realistic tests that AEDC could offer was 

full-scale engine test at Mach 4.75. Indeed, a speed of Mach 7 was close to the Mach 

8 of Tunnel B. The X-15 also could anchor a program of hypersonic studies that 

soon would have hotshot tunnels and would deal with speeds up to orbital velocity 

and beyond. And while previous X-planes were seeing their records broken by jet 

fighters, it would be some time before any other plane flew at such speeds.

The thermal environment of the latest aircraft was driving designers to the use of 

titanium and steel. The X-15 was to use Inconel X, which had still better properties. 

This nickel alloy was to be heat-treated and welded, thereby developing valuable 

shop-floor experience in its use. In addition, materials problems would be pervasive 

in building a working X-15. The success of a flight could depend on the proper 

choice of lubricating oil.

The performance of the X-15 meant that it needed more than good aerodynamics. The X-2 was already slated to execute brief leaps out of the atmosphere. Thus, 

in September 1956 test pilot Iven Kincheloe took it to 126,200 feet, an altitude 

at which his ailerons and tail surfaces no longer functioned.43 In the likely event 

that future interceptors were to make similar bold leaps, they would need reaction 

controls—which represented the first really new development in the field of flight 

control since the Wright Brothers.44 But the X-15 was to use such controls and 

would show people how to do it.

The X-15 would also need new flight instruments, including an angle-of-attack 

indicator. Pilots had been flying with turn-and-bank indicators for some time, with 

these gyroscopic instruments enabling them to determine their attitude while flying 

blind. The X-15 was to fly where the skies were always clear, but still it needed to 

determine its angle with respect to the oncoming airflow so that the pilot could set 

Armed Services Committee, won the new AEDC for his home state of Tennessee 

by offering to give the Air Force an existing military base, the 40,000-acre Camp 

Forrest. It was located near Tullahoma, far from cities and universities, but the Air 

Force was accustomed to operating in remote areas. It accepted this offer in April 

1948, with the firm of ARO, Inc. providing maintenance and operation.35

There was no interest in reproducing the research facilities of NACA, for the 

AEDC was to conduct its own activities. Engine testing was to be a specialty, and 

the first facility at this center was an engine test installation that had been “liberated” from the German firm of BMW. But the Air Force soon was installing its own 

equipment, achieving its first supersonic flow within its Transonic Model Tunnel early 

in 1953. Then, during 1954, events showed that AEDC was ready to conduct engineering development on a scale well beyond anything that NACA could envision.36

That year saw the advent of the 16-Foot Propulsion Wind Tunnel, with a test 

section 16 feet square. NACA had larger tunnels, but this one approached Mach 

3.5 and reached Mach 4.75 under special operating conditions. A Mach of 4.75 

had conventionally been associated with the limited run times of blowdown tunnels, but this tunnel, known as 16S, was a continuous-flow facility. It was unparalleled for exercising full-scale engines for realistic durations over the entire supersonic 

range.37

In December 1956 it tested the complete propulsion package of the XF-103, 

which had a turbojet with an afterburner that functioned as a ramjet. This engine 

had a total length of 39 feet. But the test section within 16S had a length of 40 feet, 

which gave room to spare.38 In addition, the similar Engine Test Facility accommodated the full-scale SRJ47 engine of Navaho, with a 51-inch diameter that made it 

the largest ramjet engine ever built.39

The AEDC also jumped into hypersonics with both feet. It already had an Engine 

Test Facility, a Gas Dynamics Facility (renamed the Von Karman Gas Dynamics 

Facility in 1959), and a Propulsion Wind Tunnel, the 16S. During 1955 it added 

a ramjet center to the Engine Test Facility, which many people regarded as a fourth 

major laboratory.40 Hypersonic wind tunnels were also on the agenda. Two 50-inch 

installations were in store, to operate respectively at Mach 8 and Mach 10. Both 

were continuous-flow facilities that used a 92,500-horsepower compressor system. 

Tunnel B, the Mach 8 facility, became operational in October 1958. Tunnel C, the 

Mach 10 installation, prevented condensation by heating its air to 1,450oF using 

a combustion heater and a 12-megawatt resistance heater. It entered operation in 

May 1960.41

The AEDC also conducted basic research in hypersonics. It had not intended 

to do that initially; it had expected to leave such studies to NACA, with its name 

reflecting its mission of engineering development. But the fact that it was off in the 

wilds of Tullahoma did not prevent it from attracting outstanding scientists, some 

of whom went on to work in hypersonics.
68

Facing the Heat Barrier: A History of Hypersonics

69

The X-15

“Our present research airplanes have developed startling performance only 

by the use of rocket engines and flying essentially in a vacuum. Testing 

airplanes designed for transonic flight speeds at Mach numbers between 

2 and 3 has proven, mainly, the bravery of the test pilots and the fact 

that where there is no drag, the rocket engine can propel even mediocre 

aerodynamic forms at high Mach numbers.

I am not aware of any aerodynamic or power plant improvements to airbreathing engines that have resulted from our very expensive research 

airplane program. Our modern tactical airplanes have been designed almost 

entirely on NACA and other wind-tunnel data, plus certain rocket model 

tests....”49

Drawing on Lockheed experience with the X-7, an unpiloted high-speed missile, 

he called instead for a similar unmanned test aircraft as the way to achieve Mach 7. 

However, he was a minority of one. Everyone else voted to support the committee’s 

resolution:

BE IT HEREBY RESOLVED, That the Committee on Aerodynamics 

endorses the proposal of the immediate initiation of a project to design and 

construct a research airplane capable of achieving speeds of the order of 

Mach number 7 and altitudes of several hundred thousand feet....50

The Air Force was also on board, and the next step called for negotiation of a 

Memorandum of Understanding, whereby the participants—which included the 

Navy—were to define their respective roles. Late in October representatives from 

the two military services visited Hugh Dryden at NACA Headquarters, bringing a 

draft of this document for discussion. It stated that NACA was to provide technical direction, the Air Force would administer design and construction, and the Air 

Force and Navy were to provide the funds. It concluded with the words, “Accomplishment of this project is a matter of national urgency.”51

The draft became the final MOU, with little change, and the first to sign it was 

Trevor Gardner. He was a special assistant to the Air Force Secretary and had midwifed the advent of Atlas a year earlier. James Smith, Assistant Secretary of the Navy 

for Air, signed on behalf of that service, while Dryden signed as well. These signatures all were in place two days before Christmas of 1954. With this, the groundwork was in place for the Air Force’s Air Materiel Command to issue a Request for 

Proposal and for interested aircraft companies to begin preparing their bids.52

As recently as February, all that anyone knew was that this new research aircraft, if it materialized, would be something other than an uprated X-2. The project 

up a proper nose-high attitude. This instrument would face the full heat load of reentry and had to work reliably.

It thus was not too much to call the X-15 a flying version of AEDC, and highlevel Air Force representatives were watching developments closely. In May 1954 

Hugh Dryden, Director of NACA, wrote a letter to Lieutenant General Donald 

Putt, who now was the Air Force’s Deputy Chief of Staff, Development. Dryden 

cited recent work, including that of Becker’s group, noting that these studies “will 

lead to specific preliminary proposals for a new research airplane.” Putt responded 

with his own letter, stating that “the Scientific Advisory Board has done some thinking in this area and has formally recommended that the Air Force initiate action on 

such a program.”45

The director of Wright Air Development Center (WADC), Colonel V. R. 

Haugen, found “unanimous” agreement among WADC reviews that the Langley 

concept was technically feasible. These specialists endorsed Langley’s engineering 

solutions in such areas as choice of material, structure, thermal protection, and 

stability and control. Haugen sent his report to the Air Research and Development 

Command (ARDC), the parent of WADC, in mid-August. A month later Major 

General F. B. Wood, an ARDC deputy commander, sent a memo to Air Force 

Headquarters, endorsing the NACA position and noting its support at WADC. He 

specifically recommended that the Air Force “initiate a project to design, construct, 

and operate a new research aircraft similar to that suggested by NACA without 

delay.”46

Further support came from the Aircraft Panel of the Scientific Advisory Board. 

In October it responded to a request from the Air Force Chief of Staff, General 

Nathan Twining, with its views:

“[A] research airplane which we now feel is ready for a program is one 

involving manned aircraft to reach something of the order of Mach 5 and 

altitudes of the order of 200,000 to 500,000 feet. This is very analogous 

to the research aircraft program which was initiated 10 years ago as a joint 

venture of the Air Force, the Navy, and NACA. It is our belief that a similar 

co-operative arrangement would be desirable and appropriate now.”47

The meetings contemplated in the Dryden-Putt correspondence were also under 

way. There had been one in July, at which a Navy representative had presented 

results of a Douglas Aircraft study of a follow-on to the Douglas Skyrocket. It was 

to reach Mach 8 and 700,000 feet.48

Then in October, at a meeting of NACA’s Committee on Aerodynamics, Lockheed’s Clarence “Kelly” Johnson challenged the entire postwar X-planes program. 

His XF-104 was already in flight, and he pulled no punches in his written statement:
70

Facing the Heat Barrier: A History of Hypersonics

71

The X-15

both with advanced heat-resistant metals and with the practical issues of powering 

piloted aircraft using liquid-fuel rocket engines. It even had an in-house group that 

was building such engines. Bell also was the home of the designers Robert Woods 

and Walter Dornberger, with the latter having presided over the V-2.

Dornberger’s Bomi concept already was introducing the highly useful concept of 

hot structures. These used temperature-resistant alloys such as stainless steel. Wings 

might be covered with numerous small and very hot metal panels, resembling shingles, that would radiate heat away from the aircraft. Overheating would be particularly severe along the leading edges of wings; these could be water-cooled. Insulation 

could protect an internal structure that would withstand the stresses and forces of 

flight; active cooling could protect a pilot’s cockpit and instrument compartment. 

Becker described these approaches as “the first hypersonic aircraft hot structures 

concepts to be developed in realistic meaningful detail.”54

Even so, Bell ranked third. Historian Dennis Jenkins writes that within the proposal, “almost every innovation they proposed was hedged in such a manner as to 

make the reader doubt that it would work. The proposal itself seemed rather poorly 

organized and was internally inconsistent (i.e., weights and other figures frequently 

differed between sections).”55 Yet the difficulties ran deeper and centered on the 

specifics of its proposed hot structure.

Bell adopted the insulated-structure approach, with the primary structure being 

of aluminum, the most familiar of aircraft materials and the best understood. Corrugated panels of Inconel X, mounted atop the aluminum, were to provide insulation. Freely-suspended panels of this alloy, contracting and expanding with ease, 

were to serve as the outer skin.

Yet this concept was quite unsuitable for the X-15, both on its technical merits 

and as a tool for research. A major goal of the program was to study aircraft structures at elevated temperatures, and this would not be possible with a primary structure of cool aluminum. There were also more specific deficiencies, as when Bell’s 

thermal analysis assumed that the expanding panels of the outer shell would prevent 

leakage of hot air from the boundary layer. However, the evaluation made the flat 

statement, “leakage is highly probable.” Aluminum might not withstand the resulting heating, with the loss of even one such panel leading perhaps to destructive 

heating. Indeed, the Bell insulated structure appeared so sensitive that it could be 

trusted to successfully complete only three of 13 reference flights.56

Another contender, Douglas Aircraft, had shared honors with Bell in building 

previous experimental aircraft. Its background included the X-3 and the Skyrocket, 

which meant that Douglas also had people who knew how to integrate a liquid 

rocket engine with an airplane. This company’s concept came in second.

Its design avoided reliance on insulated structures, calling instead for use of a 

heat sink. The material was to be a lightweight magnesium alloy that had excellent 

had taken form with considerable dispatch, and the key was the feasibility study 

of Becker’s group. An independent review at WADC confirmed its conclusions, 

whereupon Air Force leaders, both in uniform and in mufti, embraced the concept. 

Approval at the Pentagon then came swiftly.

In turn, this decisiveness demonstrated a willingness to take risks. It is hard today 

to accept that the Pentagon could endorse this program on the basis of just that one 

study. Moreover, the only hypersonic wind tunnel that was ready to provide supporting research was Becker’s 11-inch instrument; the AEDC hypersonic tunnels 

were still several years away from completion. But the Air Force was in no mood to 

hold back or to demand further studies and analyses.

This service was pursuing a plethora of initiatives in jet bombers, advanced fighters, and long-range missiles. Inevitably, some would falter or find themselves superseded, which would lead to charges of waste. However, Pentagon officials knew that 

the most costly weapons were the ones that America might need and not have in 

time of war. Cost-benefit analysis had not yet raised its head; Robert McNamara 

was still in Detroit as a Ford Motor executive, and Washington was not yet a city 

where the White House would deliberate for well over a decade before ordering the 

B-1 bomber into limited production. Amid the can-do spirit of the 1950s, the X-15 

won quick approval.

X-15: The Technology

Four companies competed for the main contract, covering design and construction of the X-15: Republic, Bell, Douglas, and North American. Each of them 

brought a substantial amount of hands-on experience with advanced aircraft. 

Republic, for example, had Alexander Kartveli as its chief designer. He was a highly 

imaginative and talented man whose XF-105 was nearly ready for first flight and 

whose XF-103 was in development. Republic had also built a rocket plane, the XF91. This was a jet fighter that incorporated the rocket engine of the X-1 for an extra 

boost in combat. It did not go into production, but it flew in flight tests.

Still, Republic placed fourth in the competition. Its concept rated “unsatisfactory” as a craft for hypersonic research, for it had a thin outer fuselage skin that 

appeared likely to buckle when hot. The overall proposal rated no better than average in a number of important areas, while achieving low scores in Propulsion System 

and Tanks, Engine Installation, Pilot’s Instruments, Auxiliary Power, and Landing 

Gear. In addition, the company itself was judged as no more than “marginal” in 

the key areas of Technical Qualifications, Management, and Resources. The latter 

included availability of in-house facilities and of an engineering staff not committed 

to other projects.53

Bell Aircraft, another contender, was the mother of research airplanes, having 

built the X-1 series as well as the X-2. This firm therefore had direct experience 
72

Facing the Heat Barrier: A History of Hypersonics

73

The X-15

impromptu fashion. The preliminary-design groups within industry were far more 

experienced, and it had appeared entirely possible that these experts, applying their 

seasoned judgment, might come up with better ideas. This did not happen. Indeed, 

the Bell and Douglas concepts failed even to meet an acceptable definition of the 

new research airplane. By contrast, the winning concept from North American 

amounted to a particularly searching affirmation of the work of Becker’s group.59

How had Bell and Douglas missed the boat? The government had set forth performance requirements, which these companies both had met. In the words of the 

North American proposal, “the specification performance can be obtained with very 

moderate structural temperatures.” However, “the airplane has been designed to 

tolerate much more severe heating in order to provide a practical temperature band 

within which exploration can be conducted.”

In Jenkins’s words, “the Bell proposal...was terrible—you walked away not 

entirely sure that Bell had committed themselves to the project. The exact opposite 

was true of the North American proposal. From the opening page you knew that 

North American understood what was trying to be accomplished with the X-15 

program and had attempted to design an airplane that would help accomplish the 

task—not just meet the performance specifications (which did not fully describe the 

intent of the program).”60 That intent was to build an aircraft that could accomplish 

research at 1,200oF and not merely meet speed and altitude goals.

The overall process of proposal evaluation cast the competing concepts in sharp 

relief, heightening deficiencies and emphasizing sources of potential difficulty. 

These proposals also received numerical scores, while another basis for comparison 

involved estimated program costs:

North American 81.5 percent $56.1 million

Douglas Aircraft 80.1 36.4

Bell Aircraft 75.5 36.3

Republic Aviation 72.2 47.0

North American’s concept thus was far from perfect, while Republic’s represented a serious effort. In addition, it was clear that the Air Force—which was to 

foot most of the bill—was willing to pay for what it would get. The X-15 program 

thus showed budgetary integrity, with the pertinent agencies avoiding the temptation to do it on the cheap.61

On 30 September 1955, letters went out to North American as well as to the 

unsuccessful bidders, advising them of the outcome of the competition. With this, 

engineers now faced the challenge of building and flying the X-15 as a practical 

exercise in hypersonic technology. Accordingly, it broke new ground in such areas as 

heat capacity. Indeed, its properties were so favorable that it would reach temperatures of only 600oF, while an Inconel X heat-sink airplane would go to 1,200oF.

Again, though, this concept missed the point. Managers wanted a vehicle that 

could cope successfully with temperatures of 1,200oF, to lay groundwork for operational fighters that could fly well beyond Mach 3. In addition, the concept had 

virtually no margin for temperature overshoots. Its design limit of 600oF was right 

on the edge of a regime of which its alloy lost strength rapidly. At 680oF, its strength 

could fall off by 90 percent. With magnesium being flammable, there was danger 

of fire within the primary structure itself, with the evaluation noting that “only 

a small area raised to the ignition temperature would be sufficient to destroy the 

aircraft.”57

Then there was North American, the home of Navaho. That missile had not 

flown, but its detailed design was largely complete and specified titanium in hot 

areas. This meant that that company knew something about using advanced metals. 

The firm also had a particularly strong rocket-engine group, which split off during 

1955 to form a new corporate division called Rocketdyne. Indeed, engines built by 

that association had already been selected for Atlas.58

North American became the winner. It paralleled the thinking at Douglas by 

independently proposing its own heat-sink structure, with the material being Inconel X. This concept showed close similarities to that of Becker’s feasibility study 

a year earlier. Still, this was not to say that the deck was stacked in favor of Becker’s approach. He and his colleagues had pursued conceptual design in a highly 

The North American X-15. (NASA)
74

Facing the Heat Barrier: A History of Hypersonics

75

The X-15

effectiveness between 90,000 

and 100,000 feet. The X-15 

therefore incorporated reaction controls, which were small 

thrusters fueled with hydrogen 

peroxide. Nose-mounted units 

controlled pitch and yaw. Other 

units, set near the wingtips, gave 

control of roll.

No other research airplane 

had ever flown with such thrusters, although the X-1B conducted early preliminary experiments and the X-2 came close to 

needing them in 1956. During 

a flight in September of that 

year, the test pilot Iven Kincheloe took it to 126,200 feet. At 

that altitude, its aerodynamic 

controls were useless. Kincheloe 

flew a ballistic arc, experiencing 

near-weightlessness for close to 

a minute. His airplane banked 

to the left, but he did not try to counter this movement, for he knew that his X-2 

could easily go into a deadly tumble.66

In developing reaction controls, an important topic for study involved determining the airplane handling qualities that pilots preferred. Initial investigations used 

an analog computer as a flight simulator. The “airplane” was disturbed slightly; 

a man used a joystick to null out the disturbance, achieving zero roll, pitch, and 

yaw. These experiments showed that pilots wanted more control authority for roll 

than for pitch or yaw. For the latter, angular accelerations of 2.5 degrees per second 

squared were acceptable. For roll, the preferred control effectiveness was two to four 

times greater.

Flight test came next. The X-2 would have served splendidly for this purpose, 

but only two had been built, with both being lost in accidents. At NACA’s HighSpeed Flight Station, investigators fell back on the X-1B, which was less capable 

but still useful. In preparation for its flights with reaction controls, the engineers 

built a simulator called the Iron Cross, which matched the dimensions and inertial 

characteristics of this research plane. A pilot, sitting well forward along the central 

arm, used a side-mounted control stick to actuate thrusters that used compressed 

Attitude control of a hypersonic airplane using aerodynamic controls and reaction controls. (U.S. Air Force)

metallurgy and fabrication, onboard instruments, reaction controls, pilot training, 

the pilot’s pressure suit, and flight simulation.62

Inconel X, a nickel alloy, showed good ductility when fully annealed and had some 

formability. When severely formed or shaped, though, it showed work-hardening, 

which made the metal brittle and prone to crack. Workers in the shop addressed this 

problem by forming some parts in stages, annealing the workpieces by heating them 

between each stage. Inconel X also was viewed as a weldable alloy, but some welds 

tended to crack, and this problem resisted solution for some time. The solution lay 

in making welds that were thicker than the parent material. After being ground flat, 

their surfaces were peened—bombarded with spherical shot—and rolled flush with 

the parent metal. After annealing, the welds often showed better crack resistance 

than the surrounding Inconel X.

A titanium alloy was specified for the internal structure of the wings. It proved 

difficult to weld, for it became brittle by reacting with oxygen and nitrogen in the 

air. It therefore was necessary to enclose welding fixtures within enclosures that 

could be purged with an inert gas such as helium and to use an oxygen-detecting 

device to determine the presence of air. With these precautions, it indeed proved 

possible to weld titanium while avoiding embrittlement.63

Greases and lubricants posed their own problems. Within the X-15, journal 

and antifriction bearings received some protection from heat and faced operating temperatures no higher than 600oF. This nevertheless was considerably hotter 

than engineers were accustomed to accommodating. At North American, candidate 

lubricants underwent evaluation by direct tests in heated bearings. Good greases 

protected bearing shafts for 20,000 test cycles and more. Poor greases gave rise to 

severe wearing of shafts after as few as 350 cycles.64

In contrast to conventional aircraft, the X-15 was to fly out of the sensible atmosphere and then re-enter, with its nose high. It also was prone to yaw while in nearvacuum. Hence, it needed a specialized instrument to determine angles of attack 

and of sideslip. This took form as the “Q-ball,” built by the Nortronics Division of 

Northrop Aircraft. It fitted into the tip of the X-15’s nose, giving it the appearance 

of a greatly enlarged tip of a ballpoint pen.

The ball itself was cooled with liquid nitrogen to withstand air temperatures as 

high as 3,500oF. Orifices set within the ball, along yaw and pitch planes, measuring 

differential pressures. A servomechanism rotated the ball to equalize these pressures by 

pointing the ball’s forward tip directly into the onrushing airflow. With the direction 

of this flow thus established, the pilot could null out any sideslip. He also could raise 

the nose to a desired angle of attack. “The Q-ball is a go-no go item,” the test pilot 

Joseph Walker told Time magazine in 1961. “Only if she checks okay do we go.”65

To steer the aircraft while in flight, the X-15 mounted aerodynamic controls. 

These retained effectiveness at altitudes well below 100,000 feet. However, they lost 
76

Facing the Heat Barrier: A History of Hypersonics

77

The X-15

was a variable gain, which changed automatically in response to flight conditions. 

Within the Air Force, the Flight Control Laboratory at WADC had laid groundwork with a program dating to 1955. Adaptive-gain controls flew aboard F-94 and 

F-101 test aircraft. The X-15 system, the Minneapolis Honeywell MH-96, made its 

first flight in December 1961.70

How did it work? When a pilot moved the control stick, as when changing the 

pitch, the existing value of gain in the pitch channel caused the aircraft to respond 

at a certain rate, measured by a rate gyro. The system held a stored value of the 

optimum pitch rate, which reflected preferred handling qualities. The adaptive-gain 

control compared the measured and desired rates and used the difference to determine a new value for the gain. Responding rapidly, this system enabled the airplane 

to maintain nearly constant control characteristics over the entire flight envelope.71

The MH-96 made it possible to introduce the X-15’s blended aerodynamic and 

reaction controls on the same control stick. This blending occurred automatically 

in response to the changing gains. When the gains in all three channels—roll, pitch, 

and yaw—reached 80 percent of maximum, thereby indicating an imminent loss of 

effectiveness in the aerodynamic controls, the system switched to reaction controls. 

During re-entry, with the airplane entering the sensible atmosphere, the system 

returned to aerodynamic control when all the gains dropped to 60 percent.72

The X-15 flight-control system thus stood three steps removed from the conventional stick-and-cable installations of World War II. It used hydraulically-boosted 

controls; it incorporated automatic stability augmentation; and with the MH-96, 

it introduced adaptive gain. Fly-by-wire systems lay ahead and represented the next 

steps, with such systems being built both in analog and digital versions.

Analog fly-by-wire systems exist within the F-16A and other aircraft. A digital 

system, as in the space shuttle, uses a computer that receives data both from the 

pilot and from the outside world. The pilot provides input by moving a stick or 

sidearm controller. These movements do not directly actuate the ailerons or rudder, 

as in days of old. Instead, they generate signals that tell a computer the nature of 

the desired maneuver. The computer then calculates a gain by applying control 

laws, which take account of the plane’s speed and altitude, as measured by onboard 

instruments. The computer then sends commands down a wire to hydraulic actuators co-mounted with the controls to move or deflect these surfaces so as to comply 

with the pilot’s wishes.73

The MH-96 fell short of such arrangements in two respects. It was analog, not 

digital, and it was a control system, not a computer. Like other systems executing 

automatic control, the MH-96 could measure an observed quantity such as pitch 

rate, compare it to a desired value, and drive the difference to zero. But the MH-96 

was wholly incapable of implementing a control law, programmed as an algebraic 

expression that required values of airspeed and altitude. Hence, while the X-15 with 

nitrogen. This simulator was mounted on a universal joint, which allowed it to 

move freely in yaw, pitch, and roll.

Reaction controls went into the X-1B late in 1957. The test pilot Neil Armstrong, who walked on the Moon 12 years later, made three flights in this research 

plane before it was grounded in mid-1958 due to cracks in its fuel tank. Its peak 

altitude during these three flights was 55,000 feet, where its aerodynamic controls 

readily provided backup. The reaction controls then went into an F-104, which 

reached 80,000 feet and went on to see much use in training X-15 pilots. When 

the X-15 was in flight, these pilots had to transition from aerodynamic controls to 

reaction controls and back again. The complete system therefore provided overlap. 

It began blending in the reaction controls at approximately 85,000 feet, with most 

pilots switching to reaction controls exclusively by 100,000 feet.67

Since the war, with aircraft increasing in both speed and size, it had become 

increasingly impractical for a pilot to exert the physical strength to operate a plane’s 

ailerons and elevators merely by moving the control stick in the cockpit. Hydraulically-boosted controls thus were in the forefront, resembling power steering in a car. 

The X-15 used such hydraulics, which greatly eased the workload on a test pilot’s 

muscles. These hydraulic systems also opened the way for stability augmentation 

systems of increasing sophistication.

Stability augmentation represented a new refinement of the autopilot. Conventional autopilots used gyroscopes to detect deviations from a plane’s straight and 

level course. These instruments then moved an airplane’s controls so as to null these 

deviations to zero. For high-performance jet fighters, the next step was stability 

augmentation. Such aircraft often were unstable in flight, tending to yaw or roll; 

indeed, designers sometimes enhanced this instability to make them more maneuverable. Still, it was quite wearying for a pilot to have to cope with this. A stability 

augmentation system made life in the cockpit much easier.

Such a system used rate gyros, which detected rates of movement in pitch, roll, 

and yaw at so many degrees per second. The instrument then responded to these 

rates, moving the controls somewhat like before to achieve a null. Each axis of this 

control had “gain,” defining the proportion or ratio between a sensed rate of angular motion and an appropriate deflection of ailerons or other controls. Fixed-gain 

systems worked well; there also were variable-gain arrangements, with the pilot setting the value of gain within the cockpit. This addressed the fact that the airplane 

might need more gain in thin air at high altitude, to deflect these surfaces more 

strongly.68

The X-15 program built three of these aircraft. The first two used a stability augmentation system that incorporated variable gain, although in practice these aircraft 

flew well with constant values of gain, set in flight.69 The third replaced it with a 

more advanced arrangement that incorporated something new: adaptive gain. This 
78

Facing the Heat Barrier: A History of Hypersonics

79

The X-15

The simulations indeed were realistic in their cockpit displays, but they left out 

an essential feature: the g-loads, produced both by rocket thrust and by deceleration 

during re-entry. In addition, a failure of the stability augmentation system, during 

re-entry, could allow the airplane to oscillate in pitch or yaw. This would change its 

drag characteristics, imposing a substantial cyclical force.

To address such issues, investigators installed a flight simulator within the gondola of a centrifuge at the Naval Air Development Center in Johnsville, Pennsylvania. The gondola could rotate on two axes while the centrifuge as a whole was turning. It not only produced g-forces, but 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 a yaw or pitch oscillation.78

Not all test pilots rode the centrifuge. William “Pete” Knight, who stood among 

the best, was one who did not. His training, coupled with his personal coolness 

and skill, enabled him to cope even with an extreme emergency. In 1967, during 

a planned flight to 250,000 feet, an X-15 experienced a complete electrical failure 

while climbing through 107,000 feet at Mach 4. This failure brought the shutdown 

of both auxiliary power units and hence of both hydraulic systems. Knight, the 

pilot, succeeded in restarting one of these units, which restored hydraulic power. He 

still had zero electrical power, but with his hydraulics, he now had both his aerodynamic and reaction controls. He rode his plane to a peak of 173,000 feet, re-entered 

the atmosphere, made a 180-degree turn, and glided to a safe landing on Mud Lake 

near Tonopah, Nevada.79

During such flights, as well as during some exercises in the centrifuge, pilots 

wore a pressure suit. Earlier models had already been good enough to allow the 

test pilot Marion Carl to reach 83,235 feet in the Douglas Skyrocket in 1953. Still, 

some of those versions left much to be desired. Time magazine, in 1952, discussed 

an Air Force model that allowed a pilot to breathe, but “with difficulty. His hands, 

not fully pressurized, swell up with blue venous blood. His throat is another trouble 

spot; the medicos have not yet learned how to pressurize a throat without strangling 

its owner.”80

The David G. Clark Company, a leading supplier of pressure suits for Air Force 

flight crews, developed a greatly improved model for the X-15. Such suits tended 

to become rigid and hard to bend when inflated. This is also true of a child’s long 

balloon, with an internal pressure that only slightly exceeds that of the atmosphere. 

The X-15 suit was to hold five pounds per square inch of pressure, or 720 pounds 

per square foot. The X-15 cockpit had its own counterbalancing pressure, but it 

could (and did) depressurize at high altitude. In such an event, the suit was to protect the test pilot rather than leave him immobile.

The solution used an innovative fabric that contracted in circumference while it 

stretched in length. With proper attention to the balance between these two effects, 

MH-96 stood three steps removed from the fighters of the recent war, it was two 

steps removed from the digital fly-by-wire control of the shuttle.

The X-15 also used flight simulators. These served both for pilot training and for 

development of onboard systems, including the reaction controls and the MH-96. 

The most important flight simulator was built by North American. 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.74

Flight simulators dated to the war. The famous Link Trainer introduced over 

half a million neophytes to their cockpits. The firm of Link Aviation added analog 

computers in 1949, within a trainer that simulated flight in a jet fighter.75 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.

Flight simulation showed its value quite early. An initial X-15 design proved 

excessively unstable and difficult to control. The cure lay in stability augmentation. 

A 1956 paper stated that this 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 ensure “consistent and successful entries” into the atmosphere.76

The North American simulator, which was transferred to the NACA Flight 

Research Center, 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, attempting to complete the 

normal mission or, at least, 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 simulated missions introduced in-flight failures 

of the radio or Q-ball. Premature engine shutdowns imposed a requirement for safe 

landing on an alternate lakebed, which was available for emergency use.77
80

Facing the Heat Barrier: A History of Hypersonics

81

The X-15

research aircraft made only 20 flights before the program ended, prematurely, with 

the loss of the second flight vehicle. The X-15 with XLR11s thus amounted to X2s that had been brought back from the dead, and that belatedly completed their 

intended flight program.

The Big Engine, the Reaction Motors XLR99, went into service in November 

1960. It launched a program of carefully measured steps that brought the fall of one 

Mach number after another. A month after the last flight with XLR11s, in March 

1961, the pilot Robert White took the X-15 past Mach 4. This was the first time a 

piloted aircraft had flown that fast, as White raised the speed mark by nearly a full 

Mach. Mach 5 fell, also to Robert White, four months later. In November 1961 

White did it again, as he reached Mach 6.04. Once flights began with the Big 

Engine, it took only 15 of them to reach this mark and to double the maximum 

Mach that had been reached with the X-2.

Altitude flights were also on the agenda. The X-15 climbed to 246,700 feet in 

April 1962, matched this mark two months later, and then soared to 314,750 feet in 

July 1962. Again White was in the cockpit, and the Federation Aeronautique Internationale, which keeps the world’s aviation records, certified this one as the absolute 

altitude record for its class. A year later, without benefit of the FAI, the pilot Joseph 

Walker reached 354,200 feet. He thus topped 100 kilometers, a nice round number 

that put him into space without question or cavil.84

The third era in the X-15’s history took shape as an extension of the second one. 

In November 1962, with this airplane’s capabilities largely demonstrated, a serious 

landing accident caused major damage and led to an extensive rebuild. The new aircraft, designated X-15A-2, retained the Big Engine but sported external tankage for 

a longer duration of engine burn. It also took on an ablative coating for enhanced 

thermal protection.

It showed anew the need for care in flight test. In mid-1962, and for that matter 

in 1966, the X-2’s best speed stood at 4,104 miles per hour, or Mach 5.92. (Mach 

number depends on both vehicle speed and air temperature. The flight to Mach 

6.04 reached 4,093 miles per hour.) Late in 1966, flying the X-15A-2 without the 

ablator, Pete Knight raised this to Mach 6.33. Engineers then applied the ablator 

and mounted a dummy engine to the lower fin, with Knight taking this craft to 

Mach 4.94 in August 1967. Then in October he tried for more.

But the X-15A-2, with both ablator and dummy engine, now was truly a new 

configuration. Further, it had only been certified with these additions in the flight 

to Mach 4.94 and could not be trusted at higher Mach. Knight took the craft to 

Mach 6.72, a jump of nearly two Mach numbers, and this proved to be too much. 

The ablator, when it came back, was charred and pitted so severely that it could 

not be restored for another flight. Worse, shock-impingement heating burned the 

engine off its pylon and seared a hole in the lower fin, disabling the propellant ejecthe suit maintained a constant volume when pressurized, enhancing a pilot’s freedom of movement. Gloves and boots were detachable and zipped to this fabric. The 

helmet was joined to the suit with a freely-swiveling ring that gave full mobility to 

the head. Oxygen flowed into the helmet; exhalant passed through valves in a neck 

seal and pressurized the suit. Becker later described it as “the first practical full-pressure suit for pilot protection in space.”81

Thus accoutered, protected for flight in near-vacuum, X-15 test pilots rode their 

rockets as they approached the edge of space and challenged the hypersonic frontier. 

They returned with results galore for project scientists—and for the nation.

X-15: Some Results

During the early 1960s, when the nation was agog over the Mercury astronauts, 

the X-15 pointed to a future in which piloted spaceplanes might fly routinely to 

orbit. The men of Mercury went water-skiing with Jackie Kennedy, but within their 

orbiting capsules, they did relatively little. Their flights were under automatic control, which left them as passengers along for the ride. Even a monkey could do it. 

Indeed, a chimpanzee named Ham rode a Redstone rocket on a suborbital flight 

in January 1961, three months before Alan Shepard repeated it before the gaze of 

an astonished world. Later that year another chimp, Enos, orbited the Earth and 

returned safely. The much-lionized John Glenn did this only later.82

In the X-15, by contrast, only people entered the cockpit. A pilot fired the rocket, 

controlled its thrust, and set the angle of climb. He left the atmosphere, soared high 

over the top of the trajectory, and then used reaction controls to set up his re-entry. 

All the while, if anything went wrong, he had to cope with it on the spot and work 

to save himself and the plane. He maneuvered through re-entry, pulled out of his 

dive, and began to glide. Then, while Mercury capsules were using parachutes to 

splash clumsily near an aircraft carrier, the X-15 pilot goosed his craft onto Rogers 

Dry Lake like a fighter.

All aircraft depend on propulsion for their performance, and the X-15’s engine 

installations allow the analyst to divide its career into three eras. It had been designed 

from the start to use the so-called Big Engine, with 57,000 pounds of thrust, but 

delays in its development brought a decision to equip it with two XLR11 rocket 

engines, which had served earlier in the X-1 series and the Douglas Skyrocket. 

Together they gave 16,000 pounds of thrust.

Flights with the XLR11s ran from June 1959 to February 1961. The best speed 

and altitude marks were Mach 3.50 in February 1961 and 136,500 feet in August 

1961. These closely matched the corresponding numbers for the X-2 during 1956: 

Mach 3.196, 126,200 feet.83 The X-2 program had been ill-starred—it had had 

two operational aircraft, both of which were destroyed in accidents. Indeed, these 
82

Facing the Heat Barrier: A History of Hypersonics

83

The X-15

But the X-15 brought surprises in boundary-layer flow and aerodynamic heating. There was reason to believe that this flow would remain laminar, being stabilized in this condition by heat flow out of the boundary layer. This offered hope, 

for laminar flow, as compared to turbulent, meant less skin-friction drag and less 

heating. Instead, the X-15 showed mostly turbulent boundary layers. These resulted 

from small roughnesses and irregularities in the aircraft skin surface, which tripped 

the boundary layers into turbulence. Such skin roughness commonly produced turbulent boundary layers on conventional aircraft. The same proved to be true at 

Mach 6.

The X-15 had a conservative thermal design, giving large safety margins to cope 

with the prevailing lack of knowledge. The turbulent boundary layers might have 

brought large increases in the heat-transfer rates, limiting the X-15’s peak speed. 

But in another surprise, these rates proved to be markedly lower than expected. As 

a consequence, the measured skin temperatures often were substantially less than 

had been anticipated (based on existing theory as well as on wind-tunnel tests). 

These flight results, confirmed by repeated measurements, were also validated with 

further wind-tunnel work. They resisted explanation by theory, but a new empirical 

model used these findings to give a more accurate description of hypersonic heattion system and threatening the craft’s vital hydraulics. No one ever tried to fly faster 

in the X-15.85

It soon retired with honor, for in close to 200 powered flights, it had operated 

as a true instrument of hypersonic research. Its flight log showed nearly nine hours 

above Mach 3, close to six hours above Mach 4, and 87 minutes above Mach 5.86

It served as a flying wind tunnel and made an important contribution by yielding 

data that made it possible to critique the findings of experiments performed in 

ground-based tunnels. Tunnel test sections were small, which led to concern that 

their results might not be reliable when applied to full-size hypersonic aircraft. Such 

discrepancies appeared particularly plausible because wind tunnels could not reproduce the extreme temperatures of hypersonic flight.

The X-15 set many of these questions to rest. In Becker’s words, “virtually all 

of the flight pressures and forces were found to be in excellent agreement with the 

low-temperature wind-tunnel predictions.”87 In addition to lift and drag, this good 

agreement extended as well to wind-tunnel values of “stability derivatives,” which 

governed the aircraft’s handling qualities and its response to the aerodynamic controls. Errors due to temperature became important only beyond Mach 10 and were 

negligible below such speeds.

B-52 mother ship with X-15A-2. The latter mounted a dummy scramjet and carried external tanks 

X-15 with dummy Hypersonic Research Engine mounted to the lower fin. (NASA) as well as ablative thermal protection. (NASA)
84

Facing the Heat Barrier: A History of Hypersonics

85

The X-15

Caution in flight test also proved beneficial in dealing with the auxiliary power 

units (APUs). The APU, built by General Electric, was a small steam turbine driven 

by hydrogen peroxide and rotating at 51,200 revolutions per minute. Each X-15 

airplane mounted two of them for redundancy, with each unit using gears to drive 

an electric alternator and a pump for the hydraulic system. Either APU could carry 

the full electrical and hydraulic load, but failure of both was catastrophic. Lacking 

hydraulic power, a pilot would have been unable to operate his aerodynamic controls.

Midway through 1962 a sudden series of failures in a main gear began to show 

up. On two occasions, a pilot experienced complete gear failure and loss of one 

APU, forcing him to rely on the second unit as a backup. Following the second such 

flight, the other APU gear also proved to be badly worn. The X-15 aircraft then 

were grounded while investigators sought the source of the problem.

They traced it to a lubricating oil, one type of which had a tendency to foam 

when under reduced pressure. The gear failures coincided with an expansion of the 

altitude program, with most of the flights above 100,000 feet having taken place 

during 1962 and later. When the oil turned to foam, it lost its lubricating properties. A different type had much less tendency to foam; it now became standard. 

Designers also enclosed the APU gearbox within a pressurized enclosure. Subsequent flights again showed reliable APU operation, as the gear failures ceased.92

Within the X-15 flight-test program, the contributions of its research pilots were 

decisive. A review of the first 44 flights, through November 1961, showed that 13 of 

them would have brought loss of the aircraft in the absence of a pilot and of redundancies in onboard systems. The actual record showed that all but one of these missions had been successfully flown, with the lone exception ending in an emergency 

landing that also went well.93

Still there were risks. The dividing line between a proficient flight and a disastrous one, between life and death for the pilot, could be narrow indeed, and the man 

who fell afoul of this was Major Mike Adams. His career in the cockpit dated to the 

Korean War. He graduated from the Experimental Test Pilot School, ranking first 

in his class, and then was accepted for the Aerospace Research Pilot School. Yeager 

himself was its director; his faculty included Frank Borman, Tom Stafford, and Jim 

McDivitt, all of whom went on to win renown as astronauts. Yeager and his selection board picked only the top one percent of this school’s applicants.94

Adams made his first X-15 flight in October 1966. The engine shut down prematurely, but although he had previously flown this craft only in a simulator, he 

successfully guided his plane to a safe landing on an emergency dry lakebed. A year 

later, in the fall of 1967, he trained for his seventh mission by spending 23 hours in 

the simulator. The flight itself took place on 15 November.

As he went over the top at 266,400 feet, his airplane made a slow turn to the 

right that left it yawing to one side by 15 degrees.95 Soon after, Adams made his 

ing. Because this model predicted less heating and lower temperatures, it permitted 

design of vehicles that were lighter in weight.88

An important research topic involved observation of how the X-15 itself would 

stand up to thermal stresses. The pilot Joseph Walker stated that when his craft 

was accelerating and heating rapidly, “the airplane crackled like a hot stove.” This 

resulted from buckling of the skin. The consequences at times could be serious, as 

when hot air leaked into the nose wheel well and melted aluminum tubing while in 

flight. On other occasions, such leaks destroyed the nose tire.89

Fortunately, such problems proved manageable. For example, the skin behind 

the wing leading edge showed local buckling during the first flight to Mach 5.3. The 

leading edge was a solid bar of Inconel X that served as a heat sink, with thin slots or 

expansion joints along its length. The slots tripped the local airflow into turbulence, 

with an accompanying steep rise in heat transfer. This created hot spots, which led 

to the buckling. The cure lay in cutting additional expansion slots, covering them 

with thin Inconel tabs, and fastening the skin with additional rivets. The wing leading edge faced particularly severe heating, but these modifications prevented buckling as the X-15 went beyond Mach 6 in subsequent flights.

Buckling indeed was an ongoing problem, and an important way to deal with 

it lay in the cautious step-by-step program of advance toward higher speeds. This 

allowed problems of buckling to appear initially in mild form, whereas a sudden 

leap toward record-breaking performance might have brought such problems in 

forms so severe as to destroy the airplane. This caution showed its value anew as 

buckling problems proved to lie behind an ongoing difficulty in which the cockpit 

canopy windows repeatedly cracked.

An initial choice of soda-lime glass for these windows gave way to alumino-silicate glass, which had better heat resistance. The wisdom of this decision became 

clear in 1961, when a soda-lime panel cracked in the course of a flight to 217,000 

feet. However, a subsequent flight to Mach 6.04 brought cracking of an aluminosilicate panel that was far more severe. The cause again was buckling, this time in 

the retainer or window frame. It was made of Inconel X; its buckle again produced 

a local hot spot, which gave rise to thermal stresses that even this heat-resistant glass 

could not withstand. The original retainers were replaced with new ones made of 

titanium, which had a significantly lower coefficient of thermal expansion. Again 

the problem disappeared.90

The step-by-step test program also showed its merits in dealing with panel flutter, wherein skin panels oscillated somewhat like a flag waving in the breeze. This 

brought a risk of cracking due to fatigue. Some surface areas showed flutter at conditions no worse than Mach 2.4 and dynamic pressure of 650 pounds per square 

foot, a rather low value. Wind-tunnel tests verified the flight results. Engineers reinforced the panels with skin doublers and longitudinal stiffeners to solve the problem. Flutter did not reappear, even at the much higher dynamic pressure of 2,000 

pounds per square foot.91
86

Facing the Heat Barrier: A History of Hypersonics

87

The X-15

1 Miller, X-Planes, chs. 4, 6, 7.

2 NACA RM L54F21.

3 Gunston, Fighters, pp. 156-57; Anderson, History, pp. 430, 432.

4 NASA SP-4303, pp. 67-69.

5 Gunston, Fighters, pp. 193-94, 199.

6 Miller, X-Planes, ch. 6.

7 NASA RP-1028, p. 237.

8 Spaceflight: November 1979, p. 435; July-August 1980, pp. 270-72. 

9 Neufeld, Ballistic, p. 70.

10 NASA RP-1028, p. 237.

11 Rand Corporation: Reports SM-11827, R-217.

12 Miller, X-Planes, p. 22; NASA SP-4305, p. 350. Quote: Minutes, Committee on Aerodynam Quote: Minutes, Committee on Aerodynamics, 4 October 1951, p. 16.

13 Letter, Woods to Committee on Aerodynamics, 8 January 1952; memo, Dornberger to Woods, 

18 January 1952 (includes quotes).

14 Drake and Carman, “Suggestion” (quote, p. 1).

15 NASA SP-4305, p. 354; memo, Stone to Chief of Research (Langley).

16 Drake and Carman, “Suggestion” (quote, p. 1).

17 Brown et al., “Study,” p. 58.

18 Astronautics & Aeronautics, February 1964, p. 54 (includes quotes).

19 Memo, J. R. Crowley to NACA centers, 11 December 1953; minutes, Interlaboratory Research 

Airplane Projects Panel, 4-5 February 1954 (quote, p. 12).

20 Martin, “History,” p. 4; Astronautics & Aeronautics,February, 1964, p. 54.

21 Becker et al., “Research,” pp. 5-8.

22 Martin, “History,” p. 3 with insert; Becker et al., “Research,” p. 24.

23 Becker et al., “Research,” p. 18. Alloy composition: memo, Rhode to Gilruth, 4 August 1954.

24 Hallion, Hypersonic, p. 386 (includes quotes).

25 Astronautics & Aeronautics, February 1964, p. 58; Brown, “Study,” pp. 38-39.

26 Astronautics & Aeronautics, February 1964, p. 56.

27 Ibid., pp. 53, 54, 56-57; NACA RM L54F21.

28 Becker et al., “Research,” pp. 8, 21.

29 Budgets: NACA: NASA SP-4305, p. 428. Air Force: Hansen, Almanac, p. 757.

30 Heppenheimer, Turbulent, pp. 149-51, 197-200 (quote, p. 197).

31 Gunston, Fighters, pp. 121-22.

32 Wolfe, Right Stuff, p. 39.

33 Gunston, Fighters, pp. 155-64, 170-76.

34 Miller, X-Planes, chs. 6, 11; Acta Astronautica, Volume 26 (1992), p. 743; Gunston, Fighters, 

pp. 193-94.

mistake. His instrument panel included an attitude indicator with a vertical bar. He 

could select between two modes of display, whereby this bar could indicate either 

sideslip angle or roll angle. He was accustomed to reading it as a yaw or sideslip 

angle—but he had set it to display roll.

“It is most probable that the pilot misinterpreted the vertical bar and flew it as a 

sideslip indicator,” the accident report later declared. Radio transmissions from the 

ground might have warned him of his faulty attitude, but the ground controllers 

had no data on yaw. Adams might have learned more by looking out the window, 

but he had been carefully trained to focus on his instruments. Three other cockpit 

indicators displayed the correct values of heading and sideslip angle, but he apparently kept his eyes on the vertical bar. He seems to have felt vertigo, which he had 

trained to overcome by concentrating on that single vertical needle.96

Mistaking roll for sideslip, he used his reaction controls to set up a re-entry with 

his airplane yawed at ninety degrees. This was very wrong; it should have been 

pointing straight ahead with its nose up. At Mach 5 and 230,000 feet, he went 

into a spin. He fought his way out of it, recovering from the spin at Mach 4.7 and 

120,000 feet. However, some of his instruments had been knocked badly awry. His 

inertial reference unit was displaying an altitude that was more than 100,000 feet 

higher than his true altitude. In addition, the MH-96 flight-control system made 

a fatal error.

It set up a severe pitch oscillation by operating at full gain, as it moved the horizontal stabilizers up and down to full deflection, rapidly and repeatedly. This system 

should have reduced its gain as the aircraft entered increasingly dense atmosphere, 

but instead it kept the gain at its highest value. The wild pitching produced extreme 

nose-up and nose-down attitudes that brought very high drag, along with decelerations as great as 15 g. Adams found himself immobilized, pinned in his seat by 

forces far beyond what his plane could withstand. It broke up at 62,000 feet, still 

traveling at Mach 3.9. The wings and tail came off; the fuselage fractured into three 

pieces. Adams failed to eject and died when he struck the ground.97

“We set sail on this new sea,” John Kennedy declared in 1962, “because there 

is new knowledge to be gained, and new rights to be won.” Yet these achievements 

came at a price, which Adams paid in full.98
88

Facing the Heat Barrier: A History of Hypersonics

89

The X-15

27 October 1961, p. 89.

66 NASA SP-4303, p. 76.

67 “Research-Airplane-Committee Report,” 1956, pp. 175-81; 1958, pp. 303-11. X-1B: NASA 

SP-4303, pp. 79-80, 294-95; SP-2000-4518, p. 77.

68 NASA TN D-1157, Figure 2.

69 NASA TN D-1157, TN D-1402.

70 AIAA Paper 93-0309, p. 2; Miller, X-Planes, p. 188. Start in 1955: NASA TM X-56008, p. 1.

71 NASA: SP-60, pp. 78-79; TN D-6208, pp. 7-8.

72 AIAA Paper 64-17, p. 1.

73 American Heritage of Invention & Technology, Winter 1992, pp. 19-24.

74 NASA TN D-6208, p. 10; AIAA Paper 63075, pp. 15-16.

75 American Heritage of Invention & Technology, Winter 1989, pp. 60-62.

76 Quotes: “Research-Airplane-Committee Report,” 1956, p. 84.

77 NASA: SP-60, pp. 37-38; TN D-1159, pp. 6-7.

78 “Research-Airplane-Committee Report,” 1958, pp. 107-16.

79 AIAA Paper 93-0309, p. 5; NASA SP-4303, pp. 118, 336.

80 NASA SP-4222, pp. 51, 113; quote: Time, 8 December 1952, 69.

81 NASA SP-4201, p. 228; “Research-Airplane-Committee Report,” 1958, pp. 117-27 (5 psi: p. 

117). Quote: NASA SP-4303, pp. 126-27.

82 NASA SP-4201, pp. 314-18, 402-07.

83 NASA SP-4303, pp. 315-16, 330; Miller, X-Planes, pp. 64, 191-192.

84 Miller, X-Planes, pp. 192-93.

85 Ibid., pp. 189-95; NASA SP-4303, pp. 121-22.

86 NASA CP-3105, p. 165.

87 Quote: Raumfahrtforschung, March-April 1969: 47.

88 Ibid., pp. 47-48; Hallion, Hypersonic, p. 164. 

89 Hunley, “Significance,” p. 4; Astronautics & Aeronautics, March 1964, p. 23. Quote: AIAA 

Paper 93-0309, p. 3.

90 AIAA Paper 93-0309, pp. 3, 9; NASA TN D-1278, pp. 8-9, 40-41; Raumfahrtforschung, 

March-April 1969, p. 49.

91 Astronautics & Aerospace Engineering, June 1963, p. 25; NASA TN D-1278, pp. 10-11.

92 Astronautics & Aerospace Engineering, June 1963, pp. 28-29.

93 NASA TN D-1278, pp. 26-28.

94 Aviation Week, 26 August 1968, p. 97; Yeager and Janos, Yeager, p. 267.

95 Aviation Week, 12 August 1968, p. 104; 26 August 1968, pp. 97, 100.

96 Aviation Week, 26 August 1968, p. 85-105 (quote, p. 101).

97 NASA SP-4303, pp. 123-25. Numbers: Aviation Week, 12 August 1968,pp. 104, 117.

98 Quote: New York Times, 13 September 1962, p. 16.

35 “History of the Arnold Engineering Development Center,” 1 August 1944 to 1 January 1951; 

Von Karman and Edson, Wind, pp. 298-300.

36 “History of the Arnold Engineering Development Center,” 1 January to 1 June 1953.

37 NASA RP-1132, pp. 199-199A.

38 Gunston, Fighters, pp. 193-94; Air Enthusiast, July-September 1978, pp. 206-07.

39 “History of the Arnold Engineering Development Center,” 1 July to 1 December 1957; “Standard Missile Characteristics: XSM-64 Navaho,” p. 3.

40 “History of the Arnold Engineering Development Center,” 1 July to 1 December 1955; 1 July 

to 1 December 1959.

41 DTIC AD-299774, pp. 1-2.

42 DTIC AD-098980; Quarterly Review: Summer 1959, p. 81; Fall 1959, p. 57.

43 NASA SP-4303, p. 76.

44 Glenn Curtiss, rather than the Wrights, was the first to use ailerons. But in subsequent litigation, the Wrights established that his innovation in fact was covered under their patent, which 

had been issued in 1906. See Heppenheimer, First Flight, ch. 9.

45 Letters, Dryden to Putt, 4 May 1954; Putt to Dryden, 26 May 1954.

46 Hallion, Hypersonic, pp. 2-5 (includes quotes).

47 Ibid., p. xxiii Extended quote, pp. xxvi, xxix.

48 Ibid., pp. 2-3; Martin, “History,” pp. 5-6; “Agenda: Meeting to Consider Need for New Research Aircraft,” 9 July 1954.

49 Minutes of Meeting, Committee on Aerodynamics, 4-5 October 1954. Extended quote: Appendix I, p. 2.

50 Ibid., pp. 15-18 (extended quote, pp. 17-18); Martin, “History,” p. 10.

51 Martin, “History,” pp. 13-14; Memorandum of Understanding, 23 December 1954 (quote, 

paragraph H).

52 NACA copy of signed Memorandum of Understanding with handwritten notes by Clotaire 

Wood; memo, Gardner to Assistant Secretary of the Navy for Air, 11 November 1954.

53 Gunston, Fighters, pp. 191-94; “Evaluation Report,” pp. 82, 96, 99, 104-05.

54 Hallion, Hypersonic, p. 384 (includes quote).

55 NASA SP-2004-4236, footnote, pp. 66-67.

56 “Evaluation Report,” pp. 23, 26, 99-100, 103; quote, p. 28.

57 Hallion, “American,” p. 298; “Evaluation Report,” pp. 44, 49, 50, 100-01 (quote, p. 49).

58 “Thirty Years” (Rocketdyne); Gibson, Navaho, pp. 36, 40-41.

59 “Evaluation Report,” p. 65; memo from Becker, 13 June 1955.

60 NASA SP-2004-4236, p. 72 (includes quotes; see also footnotes).

61 “Evaluation Report,” pp. 5, 6, 98.

62 Hallion, Hypersonic, p. 13.

63 “Research-Airplane-Committee Report,” 1958, pp. 243-73.

64 Ibid., p. 264.

65 Ibid., pp. 151-58; Miller, X-Planes, p. 200. 3,500oF: Hallion, Hypersonic, p. 158; quote: Time, 

Three new aircraft engines emerged from World War II: the turbojet, the ramjet, 

and the liquid rocket. The turbojet was not suitable for hypersonic flight, but the 

rocket and the ramjet both gave rise to related airbreathing concepts that seemed to 

hold promise.

Airbreathing rockets drew interest, but it was not possible to pump in outside 

air with a conventional compressor. Such rockets instead used liquid hydrogen fuel 

as a coolant, to liquefy air, with this liquid air being pumped to the engine. This 

arrangement wasted cooling power by also liquefying the air’s nonflammable hydrogen, and so investigators sought ways to remove this nitrogen. They wanted a flow 

of nearly pure liquid oxygen, taken from the air, for use as the oxidizer.

Ramjets provided higher flight speeds than turbojets, but they too had limits. 

Antonio Ferri, one of Langley’s leading researchers, took the lead in conceiving of 

ramjets that appeared well suited to flight at hypersonic and perhaps even orbital 

speeds, at least on paper. Other investigators studied combined-cycle engines. The 

ejector ramjet, for one, sought to integrate a rocket with a ramjet, yielding a single 

compact unit that might fly from a runway to orbit.

Was it possible to design a flight vehicle that in fact would do this? Ferri thought 

so, as did his colleague Alexander Kartveli of Republic Aviation. Air Force officials 

encouraged such views by sponsoring a program of feasibility studies called Aerospaceplane. Designers at several companies contributed their own ideas.

These activities unfolded within a world where work with conventional rockets 

was advancing vigorously.1

 In particular, liquid hydrogen was entering the mainstream of rocket engineering.2

 Ramjets also won acceptance as standard military 

engines, powering such missiles as Navaho, Bomarc, Talos, and the X-7. With this 

background, for a time some people believed that even an Aerospaceplane might 

prove feasible.

Ramjets As Military Engines

The ramjet and turbojet relied on fundamentally the same thermodynamic cycle. 

Both achieved compression of inflowing air, heated the compressed flow by burning 

91

First Thoughts of

Hypersonic Propulsion

4
92

Facing the Heat Barrier: A History of Hypersonics

93

First Thoughts of Hypersonic Propulsion

The X-7 craft were based at Holloman Air Force Base in New Mexico, which 

was an early center for missile flight test. The first flight took place in April 1951, 

with a ramjet of 20-inch diameter built by Wright Aeronautical. The X-7 soon 

took on the role of supporting developmental tests of a 28-inch engine built by the 

Marquardt Company and intended for use with the Bomarc missile. Flights with 

this 28-inch design began in December 1952 and achieved a substantial success 

the following April. The engine burned for some 20 seconds; the vehicle reached 

59,500 feet and Mach 2.6. This exceeded the Mach 2.44 of the X-1A rocket plane 

in December 1953, piloted by Chuck Yeager. Significantly, although the X-7 was 

unpiloted, it remained aerodynamically stable during this flight. By contrast, the X1A lost stability and fell out of the sky, dropping 51,000 feet before Yeager brought 

it back under control.4

The X-7 soon became a workhorse, running off some one hundred missions 

between 1955 and the end of the program in July 1960. It set a number of records, 

including range and flight time of 134 miles and 552 seconds, respectively. Its altitude mark of 106,000 feet, achieved with an airbreathing ramjet, compared with 

126,200 feet for the rocket-powered X-2 research airplane.5

Other achievements involved speed. The vehicle had been built of heat-treated 

4130 steel, with the initial goal being Mach 3. The program achieved this—and 

simply kept going. On 29 August 1957 it reached Mach 3.95 with a 28-inch Marquardt engine. Following launch from a B-50 at 33,700 feet, twin solid motors 

mounted beneath the wings boosted the craft to Mach 2.25. These boosters fell 

away; the ramjet ignited, and the vehicle began to climb at a 20-degree angle before 

leveling out at 54,500 feet. It then went into a very shallow dive. The engine continued to operate, as it ran for a total of 91 seconds, and acceleration continued until 

the craft attained its peak Mach at fuel exhaustion. It was recovered through use of 

its parachute and nose spike, and temperature-sensitive paints showed that it had 

experienced heating to more than 600°F. This heating also brought internal damage 

to the engine.6

Even so, the X-7 was not yet at the limit of its capabilities. Fitted with a 36-inch 

ramjet, again from Marquardt, it flew to Mach 4.31 on 17 April 1958. This time 

the drop from the B-50 came at 28,500 feet, with the engine igniting following 

rocket boost to Mach 1.99. It operated for 70.9 seconds, propelling the vehicle to 

a peak altitude of 60,000 feet. By then it was at Mach 3.5, continuing to accelerate as it transitioned to level flight. It reached its maximum Mach—and sustained 

a sharp drop in thrust three seconds later, apparently due to an engine tailpipe 

failure. Breakup of the vehicle occurred immediately afterward, with the X-7 being 

demolished.7

This flight set a record for airbreathing propulsion that stands to this day. Its 

speed of 2,881 miles per hour (mph) compares with the record for an afterburning 

fuel, and obtained thrust by allowing the hot airflow to expand through a nozzle. 

The turbojet achieved compression by using a rotating turbocompressor, which 

inevitably imposed a requirement to tap a considerable amount of power from its 

propulsive jet by placing the turbine within the flow. A ramjet dispensed with this 

turbomachinery, compressing the incoming flow by the simple method of processing it through a normal or oblique shock. This brought the promise of higher flight 

speed. However, a ramjet paid for this advantage by requiring an auxiliary boost, 

typically with a rocket, to accelerate it to speeds at which this shock could form.3

The X-7 served as a testbed for development of ramjet engines as mainstream 

propulsion units. With an initial concept that dated to December 1946, it took 

shape as a three-stage flight vehicle. The first stage, a B-29 and later a B-50 bomber, 

played the classic role of lifting the aircraft to useful altitudes. Such bombers also 

served in this fashion as mother ships for the X-1 series, the X-2, and in time the 

X-15. For the X-7, a solid propellant rocket served as the second stage, accelerating 

the test aircraft to a speed high enough for sustained ramjet operation. The ramjet 

engine, slung beneath its fuselage, provided further acceleration along with highspeed cruise. Recovery was achieved using a parachute and a long nose spike that 

pierced the ground like a lance. This enabled the X-7 to remain upright, which 

protected the airframe and engine for possible re-use.

The X-7. (U.S. Air Force)
94

Facing the Heat Barrier: A History of Hypersonics

95

First Thoughts of Hypersonic Propulsion

homed in on target drones that were more than 100 miles out to sea. The Bomarcs 

dived on them and made intercepts. The missiles were unarmed, but one of them 

actually rammed its target. A similar success occurred a year later when a Bomarc 

made a direct hit on a Regulus 2 supersonic target over the Gulf of Mexico. The missile first achieved operational status in September 1959. Three years later, Bomarc 

was in service at eight Air Force sites, with deployment of Canadian squadrons following. These missiles remained on duty until 1972.12

Paralleling Bomarc, the Navy 

pursued an independent effort 

that developed a ship-based 

antiaircraft missile named Talos, 

after a mythical defender of the 

island of Crete. It took shape at a 

major ramjet center, the Applied 

Physics Laboratory (APL) of 

Johns Hopkins University. Like 

Bomarc, Talos was nuclear-capable; Jane’s gave its speed as Mach 

2.5 and its range as 65 miles.

An initial version first flew in 

1952, at New Mexico’s White 

Sands Missile Range. A prototype of a nuclear-capable version made its own first flight in 

December 1953. The Korean 

War had sparked development 

of this missile, but the war ended 

in mid-1953 and the urgency 

diminished. When the Navy 

selected the light cruiser USS 

Galveston for the first operational deployment of Talos, the conversion of this ship 

became a four-year task. Nevertheless, Talos finally joined the fleet in 1958, with 

other cruisers installing it as well. It remained in service until 1980.13

One military ramjet project, that of Navaho, found itself broken in mid-stride. 

Although the ultimate version was to have intercontinental range, the emphasis 

during the 1950s was on an interim model with range of 2,500 miles, with the 

missile cruising at Mach 2.75 and 76,800 feet. The missile used a rocket-powered 

booster with liquid-fueled engines built by Rocketdyne. The airplane-like Navaho 

mounted two 51-inch ramjets from Wright Aeronautical, which gave it the capability to demonstrate long-duration supersonic cruise under ramjet power.14

The Talos. (National Archives and Records 

Administration)

turbojet of 2,193 mph, set in an SR-71 

in 1976.8

 Moreover, while the X-7 was 

flying to glory, the Bomarc program that 

it supported was rolling toward operational deployment.

The name “Bomarc” derives from the 

contractors Boeing and the Michigan 

Aeronautical Research Center, which 

conducted early studies. It was a singlestage, ground-launched antiaircraft missile that could carry a nuclear warhead. 

A built-in liquid-propellant rocket provided boost; it was replaced by a solid 

rocket in a later version. Twin ramjets 

sustained cruise at Mach 2.6. Range of 

the initial operational model was 250 

miles, later extended to 440 miles.9

Specifications for this missile were 

written in September 1950. In January 

1951 an Air Force letter contract designated Boeing as the prime contractor, 

with Marquardt Aircraft winning a subcontract to build its ramjet. The development of this engine went forward 

rapidly. In July 1953 officials of the Air 

Force’s Air Materiel Command declared 

that work on the 28-inch engine was 

essentially complete.10

Flight tests were soon under way. An 

Air Force review notes that a test vehicle 

“traveled 70 miles in 1.5 minutes to complete a most successful test of 17 June 

1954.” The missile “cruised at Mach 3+ for 15 seconds and reached an altitude 

of 56,000 feet.” In another flight in February 1955, it reached a peak altitude of 

72,000 feet as its ramjet burned for 245 seconds. This success brought a decision to 

order Bomarc into production. Four more test missiles flew with their ramjets later 

that year, with all four scoring successes.11

Other activity turned Bomarc from a missile into a weapon system, integrating it 

with the electronic Semi-Automatic Ground Environment (SAGE) that controlled 

air defense within North America. In October 1958, Bomarcs scored a spectacular 

success. Controlled remotely from a SAGE center 1,500 miles away, two missiles 

The Bomarc. (U.S. Air Force)
96

Facing the Heat Barrier: A History of Hypersonics

97

First Thoughts of Hypersonic Propulsion

Its engine, also from Wright Aeronautical, combined a turbojet and ramjet 

within a single package. The ramjet doubled as an afterburner, with internal doors 

closing off the ramjet’s long inlet duct. Conversion to pure ramjet operation took 

seven seconds. This turboramjet showed considerable promise. At Arnold Engineering Development Center, an important series of ground tests was slated to require 

as much as six weeks. They took only two weeks, with the engine running on the 

first day.

Unfortunately, the XF-103 outstayed its welcome. The project dated to 1951; it 

took until December 1956 to carry out the AEDC tests. Much of the reason for this 

long delay involved the plane’s highly advanced design, which made extensive use 

of titanium. Still, the Mach 1.8 XF-104 took less than 18 months to advance from 

the start of engineering design to first flight, and the XF-103 was not scheduled to 

fly until 1960. The Air Force canceled it in August 1957, and aviation writer Richard DeMeis pronounced its epitaph: “No matter how promising or outstanding an 

aircraft may be, if development takes inordinately long, the mortality rate rises in 

proportion.”16

Among the five cited programs, three achieved operational status, with the X-7 

actually outrunning its initial planned performance. The feasibility of Navaho was 

never in doubt; the inlet problem was one of engineering development, not one 

that would call its practicality into question. Only the XF-103 encountered serious problems of technology that lay beyond the state of the art. The ramjet of the 

1950s thus was an engine whose time had come, and which had become part of 

mainstream design.

The XF-103 in artist’s rendering. (U.S. Air Force)

Flight tests began in November 

1956, with launches of complete 

missiles taking place at Cape Canaveral. The first four were flops; none 

even got far enough to permit ignition of the ramjets. In mid-July of 

1957, three weeks after the first 

launch of an Atlas, the Air Force 

canceled Navaho. Lee Atwood, president of North American Aviation, 

recalls that Atlas indeed had greater 

promise: “Navaho would approach 

its target at Mach 3; a good antiaircraft missile might shoot it down. 

But Atlas would come in at Mach 

20. There was no way that anyone 

would shoot it down.”

There nevertheless was hardware 

for several more launches, and there 

was considerable interest in exercising the ramjets. Accordingly, seven 

more Navahos were launched during 

the following 18 months, with the 

best flight taking place in January 

1958.

The missile accelerated on rocket 

power and leveled off, the twin ramjet 

engines ignited, and it stabilized in 

cruise at 64,000 feet. It continued in this fashion for half an hour. Then, approaching the thousand-mile mark in range, its autopilot initiated a planned turnaround 

to enable this Navaho to fly back uprange. The turn was wide, and ground controllers responded by tightening it under radio control. This disturbed the airflow near 

the inlet of the right ramjet, which flamed out. The missile lost speed, its left engine 

also flamed out, and the vehicle fell into the Atlantic. It had been airborne for 42 

minutes, covering 1,237 miles.15

Because the program had been canceled and the project staff was merely flying 

off its leftover hardware, there were no funds to address what clearly was a serious 

inlet problem. Still, Navaho at least had flown. By contrast, another project—the 

Air Force’s XF-103 fighter, which aimed at Mach 3.7—never even reached the prototype stage.

The Navaho. (Smithsonian Institution 

No. 77-10905)
98

Facing the Heat Barrier: A History of Hypersonics

99

First Thoughts of Hypersonic Propulsion

Still, there were limits to a ramjet’s effectiveness. Above Mach 5, designers faced 

increasingly difficult demands for thermal protection of an airframe and for cooling 

of the ramjet duct. With the internal flow being very hot, it became more difficult 

to add still more heat by burning fuel, without overtaxing the materials or the cooling arrangements. If the engine were to run lean to limit the temperature rise in 

the combustor, its thrust would fall off. At still higher Mach levels, the issue of heat 

addition through combustion threatened to become moot. With high internal temperatures promoting dissociation of molecules of air, combustion reactions would 

not go to completion and hence would cease to add heat.

A promising way around this problem involved doing away with a requirement 

for subsonic internal flow. Instead this airflow was to be supersonic and was to sustain combustion. Right at the outset, this approach reduced the need for internal 

cooling, for this airflow would not heat up excessively if it was fast enough. This 

relatively cool internal airflow also could continue to gain heat through combustion. It would avoid problems due to dissociation of air or failure of chemical reactions in combustion to go to completion. On paper, there now was no clear upper 

limit to speed. Such a vehicle might even fly to orbit.

Comparative performance of scramjets and other engines. Airbreathers have very high performance because they are “energy machines,” which burn fuel to heat air. Rockets have much lower 

performance because they are “momentum machines,” which physically expel flows of mass. 

(Courtesy of William Escher)

There was a ramjet industry, featuring the firms of Marquardt Aviation and 

Wright Aeronautical. Facilities for developmental testing existed, not only at these 

companies but at NACA-Lewis and the Navy’s Ordnance Aerophysics Laboratory, 

which had a large continuous-flow supersonic wind tunnel. With this background, 

a number of investigators looked ahead to engines derived from ramjets that could 

offer even higher performance.

Origins of the Scramjet

The airflow within a ramjet was subsonic. This resulted from its passage through 

one or more shocks, which slowed, compressed, and heated the flow. This was true 

even at high speed, with the Mach 4.31 flight of the X-7 also using a subsonic-combustion ramjet. Moreover, because shocks become stronger with increasing Mach, 

ramjets could achieve greater internal compression of the flow at higher speeds. This 

increase in compression improved the engine’s efficiency.

The scramjet. Oblique shocks in the isolator prevent disturbances in the combustor from propagating upstream, where they would disrupt flow in the inlet. (Courtesy of Frederick Billig)
100

Facing the Heat Barrier: A History of Hypersonics

101

First Thoughts of Hypersonic Propulsion

Also at Lewis, the aerodynamicists Richard Weber and John MacKay published 

the first important open-literature study of theoretical scramjet performance in 

1958. Because they were working entirely with equations, they too bypassed the 

problem of attaining shock-free flow in a supersonic duct by simply positing that it 

was feasible. They treated the problem using one-dimensional gas dynamics, corresponding to flow in a duct with properties at any location being uniform across the 

diameter. They restricted their treatment to flow velocities from Mach 4 to 7.

They discussed the issue of maximizing the thrust and the overall engine efficiency. They also considered the merits of various types of inlet, showing that a suitable choice could give a scramjet an advantage over a conventional ramjet. Supersonic combustion failed to give substantial performance improvements or to lead 

to an engine of lower weight. Even so, they wrote that “the trends developed herein 

indicate that the [scramjet] will offer superior performance at higher hypersonic 

flight speeds.”20

An independent effort proceeded along similar lines at Marquardt, where investigators again studied scramjet performance by treating the flow within an engine 

duct using one-dimensional gasdynamic theory. In addition, Marquardt researchers 

carried out their own successful demonstration of supersonic combustion in 1957. 

They injected hydrogen into a supersonic airflow, with the hydrogen and the air 

having the same velocity. This work overcame objections from skeptics, who had 

argued that the work at NACA-Lewis had not truly demonstrated supersonic combustion. The Marquardt experimental arrangement was simpler, and its results were 

less equivocal.21

The Navy’s Applied Physics Laboratory, home of Talos, also emerged as an early 

center of interest in scramjets. As had been true at NACA-Lewis and at Marquardt, 

this group came to the concept by way of external burning under a supersonic wing. 

William Avery, the leader, developed an initial interest in supersonic combustion 

around 1955, for he saw the conventional ramjet facing increasingly stiff competition from both liquid rockets and afterburning turbojets. (Two years later such 

competition killed Navaho.) Avery believed that he could use supersonic combustion to extend the performance of ramjets.

His initial opportunity came early in 1956, when the Navy’s Bureau of Ordnance 

set out to examine the technological prospects for the next 20 years. Avery took on 

the task of assembling APL’s contribution. He picked scramjets as a topic to study, 

but he was well aware of an objection. In addition to questioning the fundamental 

feasibility of shock-free supersonic combustion in a duct, skeptics considered that 

a hypersonic inlet might produce large pressure losses in the flow, with consequent 

negation of an engine’s thrust.

Avery sent this problem through Talos management to a young engineer, James 

Keirsey, who had helped with Talos engine tests. Keirsey knew that if a hypersonic 

ramjet was to produce useful thrust, it would appear as a small difference between 

Yet while a supersonic-combustion ramjet offered tantalizing possibilities, right 

at the start it posed a fundamental issue: was it feasible to burn fuel in the duct of 

such an engine without producing shock waves? Such shocks could produce severe 

internal heating, destroying the benefits of supersonic combustion by slowing the 

flow to subsonic speeds. Rather than seeking to achieve shock-free supersonic combustion in a duct, researchers initially bypassed this difficulty by addressing a simpler problem: demonstration of combustion in a supersonic free-stream flow.

The earliest pertinent research appears to have been performed at the Applied 

Physics Laboratory (APL), during or shortly after World War II. Machine gunners in 

aircraft were accustomed to making their streams of bullets visible by making every 

twentieth round a tracer, which used a pyrotechnic. They hoped that a gunner could 

walk his bullets into a target by watching the glow of the tracers, but experience 

showed that the pyrotechnic action gave these bullets trajectories of their own. The 

Navy then engaged two research centers to look into this. In Aberdeen, Maryland, 

Ballistic Research Laboratories studied the deflection of the tracer rounds themselves. Near Washington, DC, APL treated the issue as a new effect in aerodynamics 

and sought to make use of it.

Investigators conducted tests in a Mach 1.5 wind tunnel, burning hydrogen at 

the base of a shell. A round in flight experienced considerable drag at its base, but 

the experiments showed that this combustion set up a zone of higher pressure that 

canceled the drag. This work did not demonstrate supersonic combustion, for while 

the wind-tunnel flow was supersonic, the flow near the base was subsonic. Still, this 

work introduced APL to topics that later proved pertinent to supersonic-combustion ramjets (which became known as scramjets).17

NACA’s Lewis Flight Propulsion Laboratory, the agency’s center for studies of 

engines, emerged as an early nucleus of interest in this topic. Initial work involved 

theoretical studies of heat addition to a supersonic flow. As early as 1950, the Lewis 

investigators Irving Pinkel and John Serafini treated this problem in a two-dimensional case, as in flow over a wing or past an axisymmetric body. In 1952 they 

specifically treated heat addition under a supersonic wing. They suggested that this 

might produce more lift than could be obtained by burning the same amount of 

fuel in a turbojet to power an airplane.18

This conclusion immediately raised the question of whether it was possible to 

demonstrate supersonic combustion in a wind tunnel. Supersonic tunnels produced airflows having very low pressure, which added to the experimental difficulties. However, researchers at Lewis had shown that aluminum borohydride could 

promote the ignition of pentane fuel at air pressures as low as 0.03 atmosphere. In 

1953 Robert Dorsch and Edward Fletcher launched a research program that sought 

to ignite pure borohydride within a supersonic flow. Two years later they declared 

that they had succeeded. Subsequent work showed that at Mach 3, combustion of 

this fuel under a wing more than doubled the lift.19
102

Facing the Heat Barrier: A History of Hypersonics

103

First Thoughts of Hypersonic Propulsion

an extension of an existing type such as the ramjet. The scramjet needed a really 

high-level advocate, to draw attention within the larger realms of aerodynamics and 

propulsion. The man who took on this role was Antonio Ferri.

He had headed the supersonic wind tunnel in Guidonia, Italy. Then in 1943 the 

Nazis took control of that country and Ferri left his research to command a band 

of partisans who fought the Nazis with considerable effectiveness. This made him a 

marked man, and it was not only Germans who wanted him. An American agent, 

Moe Berg, was also on his trail. Berg found him and persuaded him to come to the 

States. The war was still on and immigration was nearly impossible, but Berg persuaded William Donovan, the head of his agency, to seek support from President 

Franklin Roosevelt himself. Berg had been famous as a baseball catcher in civilian 

life, and when Roosevelt learned that Ferri now was in the hands of his agent, he 

remarked, “I see Berg is still catching pretty well.”23

At NACA-Langley after the war, he rose in management and became director 

of the Gas Dynamics Branch in 1949. He wrote an important textbook, Elements 

of Aerodynamics of Supersonic Flows (Macmillan, 1949). Holding a strong fondness 

for the academic world, he took a professorship at Brooklyn Polytechnic Institute 

in 1951, where in time he became chairman of his department. He built up an 

aerodynamics laboratory at Brooklyn Poly and launched a new activity as a consultant. Soon he was working for major companies, drawing so many contracts 

that his graduate students could not keep up with them. He responded in 1956 by 

founding a company, General Applied Science Laboratories (GASL). With financial 

backing from the Rockefellers, GASL grew into a significant center for research in 

high-speed flight.24

He was a formidable man. Robert Sanator, a former student, recalls that “you 

had to really want to be in that course, to learn from him. He was very fast. His 

mind was constantly moving, redefining the problem, and you had to be fast to 

keep up with him. He expected people to perform quickly, rapidly.” John Erdos, 

another ex-student, adds that “if you had been a student of his and later worked for 

him, you could never separate the professor-student relationship from your normal 

working relationship.” He remained Dr. Ferri to these people, never Tony, even 

when they rose to leadership within their companies.25

He came early to the scramjet. Taking this engine as his own, he faced its technical difficulties squarely and asserted that they could be addressed, giving examples of 

approaches that held promise. He repeatedly emphasized that scramjets could offer 

performance far higher than that of rockets. He presented papers at international 

conferences, bringing these ideas to a wider audience. In turn, his strong professional 

reputation ensured that he was taken seriously. He also performed experiments as he 

sought to validate his claims. More than anyone else, Ferri turned the scramjet from 

an idea into an invention, which might be developed and made practical.

two large quantities: gross thrust and total drag. In view of uncertainties in both 

these numbers, he was unable to state with confidence that such an engine would 

work. Still he did not rule it out, and his “maybe” gave Avery reason to pursue the 

topic further.

Avery decided to set up a scramjet group and to try to build an engine for test in a 

wind tunnel. He hired Gordon Dugger, who had worked at NACA-Lewis. Dugger’s 

first task was to decide which of several engine layouts, both ducted and unducted, 

was worth pursuing. He and Avery selected an external-burning configuration with 

the shape of a broad upside-down triangle. The forward slope, angled downward, 

was to compress the incoming airflow. Fuel could be injected at the apex, with 

the upward slope at the rear allowing the exhaust to expand. This approach again 

bypassed the problem of producing shock-free flow in a duct. The use of external 

burning meant that this concept could produce lift as well as thrust.

Dugger soon became concerned that this layout might be too simple to be effective. Keirsey suggested placing a very short cowl at the apex, thereby easing problems 

of ignition and combustion. This new design lent itself to incorporation within the 

wings of a large aircraft of reasonably conventional configuration. At low speeds the 

wide triangle could retract until it was flat and flush with the wing undersurface, 

leaving the cowl to extend into the free stream. Following acceleration to supersonic 

speed, the two shapes would extend and assume their triangular shape, then function as an engine for further acceleration.

Wind-tunnel work also proceeded at APL. During 1958 this center had a Mach 

5 facility under construction, and Dugger brought in a young experimentalist 

named Frederick Billig to work with it. His first task was to show that he too could 

demonstrate supersonic combustion, which he tried to achieve using hydrogen as 

his fuel. He tried electric ignition; an APL history states that he “generated gigantic 

arcs,” but “to no avail.” Like the NACA-Lewis investigators, he turned to fuels that 

ignited particularly readily. His choice, triethyl aluminum, reacts spontaneously, 

and violently, on contact with air.

“The results of the tests on 5 March 1959 were dramatic,” the APL history continues. “A vigorous white flame erupted over the rear of [the wind-tunnel model] 

the instant the triethyl aluminum fuel entered the tunnel, jolting the model against 

its support. The pressures measured on the rear surface jumped upward.” The device 

produced less than a pound of thrust. But it generated considerable lift, supporting 

calculations that had shown that external burning could increase lift. Later tests 

showed that much of the combustion indeed occurred within supersonic regions of 

the flow.22

By the late 1950s small scramjet groups were active at NACA-Lewis, Marquardt, 

and APL. There also were individual investigators, such as James Nicholls of the 

University of Michigan. Still it is no small thing to invent a new engine, even as 
104

Facing the Heat Barrier: A History of Hypersonics

105

First Thoughts of Hypersonic Propulsion

process can take place without the formation of shocks.” Parallel injection might 

take place by building the combustor with a step or sudden widening. The flow 

could expand as it passed the step, thereby avoiding a shock, while the fuel could be 

injected at the step.29

Ferri also made an intriguing contribution in dealing with inlets, which are critical to the performance of scramjets. He did this by introducing a new concept called 

“thermal compression.” One approaches it by appreciating that a process of heat 

addition can play the role of a normal shock wave. When an airflow passes through 

such a shock, it slows in speed and therefore diminishes in Mach, while its temperature and pressure go up. The same consequences occur when a supersonic airflow 

is heated. It therefore follows that a process of heat addition can substitute for a 

normal shock.30

Practical inlets use oblique shocks, which are two-dimensional. Such shocks 

afford good control of the aerodynamics of an inlet. If heat addition is to substitute 

for an oblique shock, it too must be two-dimensional. Heat addition in a duct 

is one-dimensional, but Ferri proposed that numerous small burners, set within a 

flow, could achieve the desired two-dimensionality. By turning individual burners 

on or off, and by regulating the strength of each one’s heating, he could produce the 

desired pattern of heating that in fact would accomplish the substitution of heating 

for shock action.31

Why would one want to do this? The nose of a hypersonic aircraft produces a 

strong bow shock, an oblique shock that accomplishes initial compression of the 

airflow. The inlet rides well behind the nose and features an enclosing cowl. The 

cowl, in turn, has a lip or outer rim. For best effectiveness, the inlet should sustain 

a “shock-on-lip” condition. The shock should not impinge within the inlet, for 

only the lip is cooled in the face of shock-impingement heating. But the shock also 

should not ride outside the inlet, or the inlet will fail to capture all of the shockcompressed airflow.

To maintain the shock-on-lip condition across a wide Mach range, an inlet 

requires variable geometry. This is accomplished mechanically, using sliding seals 

that must not allow leakage of very hot boundary-layer air. Ferri’s principle of thermal compression raised the prospect that an inlet could use fixed geometry, which 

was far simpler. It would do this by modulating its burners rather than by physically 

moving inlet hardware.

Thermal compression brought an important prospect of flexibility. At a given 

value of Mach, there typically was only one arrangement of a variable-geometry 

inlet that would produce the desired shock that would compress the flow. By contrast, the thermal-compression process might be adjusted at will simply by controlling the heating. Ferri proposed to do this by controlling the velocity of injection of 

the fuel. He wrote that “the heat release is controlled by the mixing process, [which] 

His path to the scramjet began during the 1950s, when his work as a consultant brought him into a friendship with Alexander Kartveli at Republic Aviation. 

Louis Nucci, Ferri’s longtime colleague, recalls that the two men “made good sparks. 

They were both Europeans and learned men; they liked opera and history.” They 

also complemented each other professionally, as Kartveli focused on airplane design 

while Ferri addressed difficult problems in aerodynamics and propulsion. The two 

men worked together on the XF-103 and fed off each other, each encouraging the 

other to think bolder thoughts. Among the boldest was a view that there were no 

natural limits to aircraft speed or performance. Ferri put forth this idea initially; 

Kartveli then supported it with more detailed studies.26

The key concept, again, was the scramjet. Holding a strong penchant for experimentation, Ferri conducted research at Brooklyn Poly. In September 1958, at a 

conference in Madrid, he declared that steady combustion, without strong shocks, 

had been accomplished in a supersonic airstream at Mach 3.0. This placed him 

midway in time between the supersonic-combustion demonstrations at Marquardt 

and at APL.27

Shock-free flow in a duct continued to loom as a major problem. The Lewis, 

Marquardt, and APL investigators had all bypassed this issue by treating external 

combustion in the supersonic flow past a wing, but Ferri did not flinch. He took 

the problem of shock-free flow as a point of departure, thereby turning the ducted 

scramjet from a wish into a serious topic for investigation.

In supersonic wind tunnels, shock-free flow was an everyday affair. However, the 

flow in such tunnels achieved its supersonic Mach values by expanding through a 

nozzle. By contrast, flow within a scramjet was to pass through a supersonic inlet 

and then be strongly heated within a combustor. The inlet actually had the purpose 

of producing a shock, an oblique one that was to slow and compress the flow while 

allowing it to remain supersonic. However, the combustion process was only too 

likely to produce unwanted shocks, which would limit an engine’s thrust and performance.

Nicholls, at Michigan, proposed to make a virtue of necessity by turning a combustor shock to advantage. Such a shock would produce very strong heating of the 

flow. If the fuel and air had been mixed upstream, then this combustor shock could 

produce ignition. Ferri would have none of this. He asserted that “by using a suitable design, formation of shocks in the burner can be avoided.”28

Specifically, he started with a statement by NACA’s Weber and MacKay on 

combustors. These researchers had already written that the combustor needed a 

diverging shape, like that of a rocket nozzle, to overcome potential limits on the 

airflow rate due to heat addition (“thermal choking”). Ferri proposed that within 

such a combustor, “fuel is injected parallel to the stream to eliminate formation of 

shocks.... The fuel gradually mixes with the air and burns...and the combustion 
106

Facing the Heat Barrier: A History of Hypersonics

107

First Thoughts of Hypersonic Propulsion

positioning of inlet spikes was always done mechanically. In turn, the movable spike 

represented a prime example of variable geometry.

Scramjets faced similar issues, particularly near Mach 4. Ferri’s thermal-compression principle applied here as well—and raised the prospect of an inlet that might 

fight against unstarts by using thermal rather than mechanical arrangements. An 

inlet with thermal compression then might use fixed geometry all the way to orbit, 

while avoiding unstarts in the bargain.

Ferri presented his thoughts publicly as early as 1960. He went on to give a far 

more detailed discussion in May 1964, at the Royal Aeronautical Society in London. 

This was the first extensive presentation on hypersonic propulsion for many in the 

audience, and attendees responded effusively.

One man declared that “this lecture opened up enormous possibilities. Where 

they had, for lack of information, 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.” 

A. D. Baxter, a Fellow of the Society, added that Ferri “had given them an insight 

into the prospects and possibilities of extending the speed range of the airbreathing 

engine far beyond what most of them had dreamed of; in fact, assailing the field 

which until recently was regarded as the undisputed regime of the rocket.”37

Not everyone embraced thermal compression. “The analytical basis was rather 

weak,” Marquardt’s Arthur Thomas commented. “It was something that he had in 

his head, mostly. There were those who thought it was a lot of baloney.” Nor did 

Ferri help his cause in 1968, when he published a Mach 6 inlet that offered “much 

better performance” at lower Mach “because it can handle much higher flow.” His 

paper contained not a single equation.38

But Fred Billig was one who accepted the merits of thermal compression and 

gave his own analyses. He proposed that at Mach 5, thermal compression could 

increase an engine’s specific impulse, an important measure of its performance, by 

61 percent. Years later he recalled Ferri’s “great capability for visualizing, a strong 

physical feel. He presented a full plate of ideas, not all of which have been realized.”39

Combined-Cycle Propulsion Systems

The scramjet used a single set of hardware and operated in two modes, sustaining supersonic combustion as well as subsonic combustion. The transition involved 

a process called “swallowing the shock.” In the subsonic mode, the engine held a 

train of oblique shocks located downstream of the inlet and forward of the combustor. When the engine went over to the supersonic-combustion mode, these shocks 

passed through the duct and were lost. This happened automatically, when the flight 

depends on the difference of velocity of the air and of the injected gas.” Shock-free 

internal flow appeared feasible: “The fuel is injected parallel to the stream to eliminate formation of shocks [and] the combustion process can take place without the 

formation of shocks.” He added,

“The preliminary analysis of supersonic combustion ramjets...indicates 

that combustion can occur in a fixed-geometry burner-nozzle combination 

through a large range of Mach numbers of the air entering the combustion 

region. Because the Mach number entering the burner is permitted to vary 

with flight Mach number, the inlet and therefore the complete engine does 

not require variable geometry. Such an engine can operate over a large 

range of flight Mach numbers and, therefore, can be very attractive as an 

accelerating engine.”32

There was more. As noted, the inlet was to produce a bow shock of specified 

character, to slow and compress the incoming air. But if the inflow was too great, 

the inlet would disgorge its shock. This shock, now outside the inlet, would disrupt 

the flow within the inlet and hence in the engine, with the drag increasing and the 

thrust falling off sharply. This was known as an unstart.

Supersonic turbojets, such as the Pratt & Whitney J58 that powered the SR-71 

to speeds beyond Mach 3, typically were fitted with an inlet that featured a conical 

spike at the front, a centerbody that was supposed to translate back and forth to 

adjust the shock to suit the flight Mach number. Early in the program, it often did 

not work.33 The test pilot James Eastham was one of the first to fly this spy plane, 

and he recalls what happened when one of his inlets unstarted.

“An unstart has your full and undivided attention, right then. The airplane 

gives a very pronounced yaw; then you are very preoccupied with getting 

the inlet started again. The speed falls off; you begin to lose altitude. You 

follow a procedure, putting the spikes forward and opening the bypass 

doors. Then you would go back to the automatic positioning of the spike—

which many times would unstart it again. And when you unstarted on one 

side, sometimes the other side would also unstart. Then you really had to 

give it a good massage.”34

The SR-71 initially used a spike-positioning system from Hamilton Standard. 

It proved unreliable, and Eastham recalls that at one point, “unstarts were literally 

stopping the whole program.”35 This problem was eventually overcome through 

development of a more capable spike-positioning system, built by Honeywell.36

Still, throughout the development and subsequent flight career of the SR-71, the 
108

Facing the Heat Barrier: A History of Hypersonics

109

First Thoughts of Hypersonic Propulsion

within the duct that reached speeds of 40 mph, which served his studies.41 Nearly a 

century later ejectors were used to evacuate chambers that conducted rocket-engine 

tests in near-vacuum, with the ejectors rapidly pumping out the rocket exhaust. 

Ejectors also flew, being used with both the F-104 and the SR-71. Significantly, 

the value of an ejector could increase with Mach. On the SR-71, for instance, it 

contributed only 14 percent of the total thrust at Mach 2.2, but accounted for 28 

percent at Mach 3.2.42

Jack Charshafian of Curtiss-Wright, director of development of ramjets for 

Navaho, filed a patent disclosure for an ejector rocket as early as 1947. By entraining outside air, it might run fuel-rich and still burn all its fuel. Ejector concepts 

also proved attractive to other aerospace pioneers, with patents being awarded to 

Alfred Africano, a founder of the American Rocket Society; to Hans von Ohain, an 

inventor of the turbojet; and to Helmut Schelp, who stirred early interest in military 

turbojets within the Luftwaffe.43

A conventional ramjet needed a boost to reach speeds at which its air-ramming 

effect would come into play, and the hardware requirements approached those of a 

complete and separate flight vehicle. The turbojet of the XF-103 exemplified what 

was necessary, as did the large rocket-powered booster of Navaho. But after 1960 

the ejector ramjet brought the prospect of a ramjet that could produce thrust even 

when on the runway. By placing small rocket engines in a step surrounding a duct, 

a designer could leave the duct free of hardware. It might even sustain supersonic 

combustion, with the engine converting to a scramjet.

The ejector ramjet also promised to increase the propulsive efficiency by improving the match between flight speed and exhaust speed. A large mismatch greatly 

reduces the effectiveness of a propulsive jet. There would be little effective thrust, 

Performance of an ejector. Even with minimal flow of entrained air, the pressure ratio is much lower 

than that of a turbojet. A pressure ratio of 1.5 implied low efficiency. (Courtesy of William Escher)

vehicle topped a critical speed and the engine continued to burn with no diminution of thrust.

The turboramjet arrangement of the XF-103 also operated in two modes, serving both as a turbojet and as a ramjet. Here, however, the engine employed two 

sets of hardware, which were physically separate. They shared a common inlet and 

nozzle, while the ramjet also served as the turbojet’s afterburner. But only one set of 

equipment operated at any given time. Moreover, they were mounted separately and 

were not extensively integrated.40

System integration was the key concept within a third class of prime mover: the 

combined-cycle engine, which sought to integrate two separate thrust-producing 

cycles within a single set of hardware. In contrast to the turboramjet of the XF-103, 

engines of this type merged their equipment rather than keeping them separate. 

Two important concepts that did this were the ejector ramjet, which gave thrust 

even when standing still, and the Liquid Air Cycle Engine (LACE), which was an 

airbreathing rocket.

The ejector ramjet amounted to a combined-cycle system derived from a conventional ramjet. It used the ejector principle, whereby the exhaust of a rocket motor, 

placed within a duct, entrains a flow of air through the duct’s inlet. This increases 

the thrust by converting thermal energy, within the hot exhaust, to mechanical 

energy. The entrained air slows and cools the exhaust. The effect is much the same as 

when designers improve the performance of a turbojet engine by installing a fan.

Ejectors date back to the nineteenth century. Horatio Phillips, a pioneer in aeronautical research, used a steam ejector after 1870 to build a wind tunnel. His ejector 

was a ring of pipe pierced with holes and set within a duct with the holes facing 

downstream. Pressurized steam, expanding through the holes, entrained an airflow 

Ejector ramjet. Primary flow from a ducted rocket entrains a substantial secondary flow of external 

air. (U.S. Air Force)
110

Facing the Heat Barrier: A History of Hypersonics

111

First Thoughts of Hypersonic Propulsion

a hydrogen-fueled rocket could give much better performance when using oxygen 

rather than air. With oxygen liquefying at 90 K while nitrogen becomes a liquid at 

77 K, at atmospheric pressure the prospect existed of using this temperature difference to leave the nitrogen unliquefied. Nor would it be useless; it could flow within 

a precooler, an initial heat exchanger that could chill the inflowing air while reserving the much colder liquid hydrogen for the main cooler.

It did not appear feasible in practice to operate a high-capacity LACE air liquefier with the precision in temperature that could achieve this. However, a promising approach called for use of fractional distillation of liquid air, as a variant of the 

process used in oil refineries to obtain gasoline from petroleum. The distillation 

process promised fine control, allowing the nitrogen to boil off while keeping the 

oxygen liquid. To increase the throughput, the distillation was to take place within 

a rotary apparatus that could impose high g-loads, greatly enhancing the buoyancy 

of the gaseous nitrogen. A LACE with such an air separator was called ACES, Air 

Collection and Enrichment System.47

When liquid hydrogen chilled and liquefied the nitrogen in air, that hydrogen 

went only partially to waste. In effect, it transferred its coldness to the nitrogen, 

which used it to advantage in the precooler. Still, there was a clear prospect of greater 

efficiency in the heat-transfer process if one could remove the nitrogen directly from 

the ram air. A variant of ACES promised to do precisely this, using chemical separation of oxygen. The process relied on the existence of metal oxides that could take 

up additional oxygen when heated by the hot ram air and then release this oxygen 

when placed under reduced pressure. Only the oxygen then was liquefied. This 

brought the increased efficiency, for the amount of liquid hydrogen used as a coolant was reduced. This enhanced efficiency also translated into conceptual designs 

for chemical-separation ACES units that could be lighter in weight and smaller in 

size than rotary-distillation counterparts.48

Turboramjets, ramjets, scramjets, LACE, ramLACE and scramLACE, ACES: 

with all these in prospect, designers of paper engines beheld a plenitude of possibilities. They also carried a strong mutual synergism. A scramjet might use a type of 

turboramjet for takeoff, again with the scramjet duct also functioning as an afterburner. Alternately, it might install an internal rocket and become a scramLACE. 

It could use ACES for better performance, while adopting the chemical-separation 

process to derichen the use of hydrogen.

It did not take long before engineers rose to their new opportunities by conceiving of new types of vehicles that were to use these engines, perhaps to fly to orbit as 

a single stage. Everyone in aerospace was well aware that it had taken only 30 years 

to progress from Lindbergh in Paris to satellites in space. The studies that explored 

the new possibilities amounted to an assertion that this pace of technical advance 

was likely to continue.

for instance, if one had a jet velocity of 10,000 feet per second while flying in the 

atmosphere at 400 feet per second. The ejector promised to avoid this by slowing 

the overall flow.

The ejector ramjet thus offered the enticing concept of a unified engine that 

could propel a single-stage vehicle from a runway to orbit. It would take off with 

ejector-boosted thrust from its rocket, accelerate through the atmosphere by using 

the combination as an ejector-boosted ramjet and scramjet, and then go over completely to rocket propulsion for the final boost to orbit.

Yet even with help from an ejector, a rocket still had a disadvantage. A ramjet or 

scramjet could use air as its oxidizer, but a rocket had to carry heavy liquid oxygen in 

an onboard tank. Hence, there also was strong interest in airbreathing rockets. Still, 

it was not possible to build such a rocket through a simple extension of principles 

applicable to the turbojet, for there was a serious mismatch between pressures available through turbocompression and those of a rocket’s thrust chamber.

In the SR-71, for instance, a combination of inlet compression and turbocompression yielded an internal pressure of approximately 20 pounds per square inch 

(psi) at Mach 3 and 80,000 feet. By contrast, internal pressures of rocket engines 

started in the high hundreds of psi and rapidly ascended into the thousands for 

high performance. Unless one could boost the pressure of ram air to that level, no 

airbreathing rocket would ever fly.44

The concept that overcame this difficulty was LACE. It dated to 1954, and Randolph Rae of the Garrett Corporation was the inventor. LACE used liquid hydrogen both as fuel and as a refrigerant, to liquefy air. The temperature of the liquid 

hydrogen was only 21 K, far below that at which air liquefies. LACE thus called for 

incoming ram air to pass through a heat exchanger that used liquid hydrogen as the 

coolant. The air would liquefy, and then a pump could raise its pressure to whatever 

value was desired. In this fashion, LACE bypassed the restrictions on turbocompression of gaseous air. In turn, the warmed hydrogen flowed to the combustion 

chamber to burn in the liquefied air.45

At the outset, LACE brought a problem. The limited thermal capacity of liquid 

hydrogen brought another mismatch, for the system needed eight times more liquid 

hydrogen to liquefy a given mass of air than could burn in that mass. The resulting 

hydrogen-rich exhaust still had a sufficiently high velocity to give LACE a prospective advantage over a hydrogen-fueled rocket using tanked oxygen. Even so, there 

was interest in “derichening” the fuel-air mix, by making use of some of this extra 

hydrogen. An ejector promised to address this issue by drawing in more air to burn 

the hydrogen. Such an engine was called a ramLACE or scramLACE.46

A complementary strategy called for removal of nitrogen from the liquefied air, 

yielding nearly pure liquid oxygen as the product. Nitrogen does not support combustion, constitutes some three-fourths of air by weight, and lacks the redeeming 

quality of low molecular weight that could increase the exhaust velocity. Moreover, 
112

Facing the Heat Barrier: A History of Hypersonics

113

First Thoughts of Hypersonic Propulsion

Space Plane was to take off from a runway, using LACE and ACES while pumping the oxygen-rich condensate directly to the LACE combustion chambers. It 

would climb to 40,000 feet and Mach 3, cut off the rocket, and continue to fly 

using hydrogen-fueled ramjets. It was to use ACES for air collection while cruising 

at Mach 5.5 and 66,000 feet, trading liquid hydrogen for oxygen-rich liquid air 

while taking on more than 600,000 pounds of this oxidizer. Now weighing more 

than a million pounds, Space Plane would reach Mach 7 on its ramjets, then shut 

them down and go over completely to rocket power. Drawing on its stored oxidizer, 

it could fly to orbit while carrying a payload of 38,000 pounds.

Convair’s Space Plane concept. (Art by Dennis Jenkins)

Aerospaceplane

“I remember when Sputnik was launched,” says Arthur Thomas, a leader in early 

work on scramjets at Marquardt. The date was 4 October 1957. “I was doing analysis of scramjet boosters to go into orbit. We were claiming back in those days that 

we could get the cost down to a hundred dollars per pound by using airbreathers.” 

He adds that “our job was to push the frontiers. We were extremely excited and 

optimistic that we were really on the leading edge of something that was going to 

be big.”49

At APL, other investigators proposed what may have been the first concept for 

a hypersonic airplane that merited consideration. In an era when the earliest jet 

airliners were only beginning to enter service, William Avery leaped beyond the 

supersonic transport to the hypersonic transport, at least in his thoughts. His colleague Eugene Pietrangeli developed a concept for a large aircraft with a wingspan of 

102 feet and length of 175 feet, fitted with turbojets and with the Dugger-Keirsey 

external-burning scramjet, with its short cowl, under each wing. It was to accelerate 

to Mach 3.6 using the turbojets, then go over to scramjet propulsion and cruise at 

Mach 7. Carrying 130 passengers, it was to cross the country in half an hour and 

achieve a range of 7,000 miles. Its weight of 600,000 pounds was nearly twice that 

of the Boeing 707 Intercontinental, largest of that family of jetliners.50

Within the Air Force, an important prelude to similar concepts came in 1957 

with Study Requirement 89774. It invited builders of large missiles to consider 

what modifications might make them reusable. It was not hard to envision that they 

might return to a landing on a runway by fitting them with wings and jet engines, 

but most such rocket stages were built of aluminum, which raised serious issues of 

thermal protection. Still, Convair at least had a useful point of departure. Its Atlas 

used stainless steel, which had considerably better heat resistance.51

The Convair concept envisioned a new version of this missile, fitted out as a 

reusable first stage for a launch vehicle. Its wings were to use the X-15’s structure. 

A crew compartment, set atop a rounded nose, recalled that company’s B-36 heavy 

bomber. To ease the thermal problem, designers were aware that this stage, having 

burned its propellants, would be light in weight. It therefore could execute a hypersonic glide while high in the atmosphere, losing speed slowly and diminishing the 

rate of heating.52

It did not take long before Convair officials began to view this reusable Atlas as 

merely a first step into space, for the prospect of LACE opened new vistas. Beginning late in 1957, using a combination of Air Force and in-house funding, the 

company launched paper studies of a new concept called Space Plane. It took shape 

as a large single-stage vehicle with highly-swept delta wings and a length of 235 feet. 

Propulsion was to feature a combination of ramjets and LACE with ACES, installed 

as separate engines, with the ACES being distillation type. The gross weight at takeoff, 450,000 pounds, was to include 270,000 pounds of liquid hydrogen.
114

Facing the Heat Barrier: A History of Hypersonics

115

First Thoughts of Hypersonic Propulsion

promise, taking this topic as his own while leaving the rockets to his deputy, Lieutenant Colonel Edward Hall. He launched the first Air Force studies of hypersonic 

propulsion as early as 1957. In October 1959 he chaired a session on scramjets at 

the Second USAF Symposium on Advanced Propulsion Concepts.

In the wake of this meeting, he built on the earlier SR-89774 efforts and launched 

a new series of studies called Aerospaceplane. It did not aim at anything so specific 

as a real airplane that could fly to orbit. Rather, it supported design studies and 

conducted basic research in advanced propulsion, seeking to develop a base for the 

evolution of such craft in the distant future. Marquardt and GASL became heavily 

involved, as did Convair, Republic, North American, GE, Lockheed, Northrop, and 

Douglas Aircraft.58

The new effort broadened the scope of the initial studies, while encouraging 

companies to pursue their concepts to greater depth. Convair, for one, had issued 

single-volume reports on Space Plane in October 1959, April 1960, and December 

1960. In February 1961 it released an 11-volume set of studies, with each of them 

addressing a specific topic such as Aerodynamic Heating, Propulsion, Air Enrichment Systems, Structural Analysis, and Materials.59

Aerospaceplane proved too hot to keep under wraps, as a steady stream of disclosures presented concept summaries to the professional community and the general 

public. Aviation Week, hardly shy in these matters, ran a full-page article in October 

1960:

USAF PLANS RADICAL SPACE PLANE

Studies costing $20 million sought in next budget, Earth-to-orbit vehicle 

would need no large booster.60

At the Los Angeles Times, the aerospace editor Marvin Miles published headlined 

stories of his own. The first appeared in November:

LOCKHEED WORKING ON PLANE ABLE TO GO INTO ORBIT 

ALONE

Air Force Interested in Project61

Two months later another of his articles ran as a front-page headline:

HUGE BOOSTER NOT NEEDED BY AIR FORCE SPACE PLANE

Proposed Wing Vehicle Would Take Off, Return Like Conventional Craft

It particularly cited Convair’s Space Plane, with a Times artist presenting a view 

of this craft in flight.62

The concept was born in exuberance. Its planners drew on estimates “that by 

1970 the orbital payload accumulated annually would be somewhere between two 

million and 20 million pounds.” Most payloads were to run near 10,000 pounds, 

thereby calling for a schedule of three flights per day. Still the concept lacked an 

important element, for if scramjets were nowhere near the state of the art, at Convair they were not even the state of the imagination.53 Space Plane, as noted, used 

ramjets with subsonic combustion, installing them in pods like turbojets on a B-52. 

Scramjets lay beyond the thoughts of other companies as well. Thus, Northrop 

expected to use LACE with its Propulsive Fluid Accumulator (PROFAC) concept, 

which also was to cruise in the atmosphere while building up a supply of liquefied 

air. Like Space Plane, PROFAC also specified conventional ramjets.54

But Republic Aviation was home to the highly imaginative Kartveli, with Ferri 

being just a phone call away. Here the scramjet was very much a part of people’s 

thinking. Like the Convair designers, Kartveli looked ahead to flight to orbit with 

a single stage. He also expected that this goal was too demanding to achieve in a 

single jump, and he anticipated that intermediate projects would lay groundwork. 

He presented his thoughts in August 1960 at a national meeting of the Institute of 

Aeronautical Sciences.55

The XF-103 had been dead and buried for three years, but Kartveli had crafted 

the F-105, which topped Mach 2 as early as 1956 and went forward into production. He now expected to continue with a Mach 2.3 fighter-bomber with enough 

power to lift off vertically as if levitating and to cruise at 75,000 feet. Next on the 

agenda was a strategic bomber powered by nuclear ramjets, which would use atomic 

power to heat internal airflow, with no need to burn fuel. It would match the peak 

speed of the X-7 by cruising at Mach 4.25, or 2,800 mph, and at 85,000 feet.56

Kartveli set Mach 7, or 5,000 mph, as the next goal. He anticipated achieving 

this speed with another bomber that was to cruise at 120,000 feet. Propulsion was 

to come from two turbojets and two ramjets, with this concept pressing the limits 

of subsonic combustion. Then for flight to orbit, his masterpiece was slated for 

Mach 25. It was to mount four J58 turbojets, modified to burn hydrogen, along 

with four scramjets. Ferri had convinced him that such engines could accelerate 

this craft all the way to orbit, with much of the gain in speed taking place while 

flying at 200,000 feet. A small rocket engine might provide a final boost into space, 

but Kartveli placed his trust in Ferri’s scramjets, planning to use neither LACE nor 

ACES.57

These concepts drew attention, and funding, from the Aero Propulsion Laboratory at Wright-Patterson Air Force Base. Its technical director, Weldon Worth, had 

been closely involved with ramjets since the 1940s. Within a world that the turbojet 

had taken by storm, he headed a Nonrotating Engine Branch that focused on ramjets and liquid-fuel rockets. Indeed, he regarded the ramjet as holding the greater 
116

Facing the Heat Barrier: A History of Hypersonics

117

First Thoughts of Hypersonic Propulsion

On other propulsion systems:

When reduced to a common basis and compared with the best of current 

technology, all assumed large advances in the state-of-the-art.... On the 

basis of the best of current technology, none of the schemes could deliver 

useful payloads into orbits.

On vehicle design:

We are gravely concerned that too much emphasis may be placed on the 

more glamorous aspects of the Aerospace Plane resulting in neglect of 

what appear to be more conventional problems. The achievement of low 

structural weight is equally important...as is the development of a highly 

successful propulsion system.

Regarding scramjets, the panel was not impressed with claims that supersonic 

combustion had been achieved in existing experiments:

These engine ideas are based essentially upon the feasibility of diffusion 

deflagration flames in supersonic flows. Research should be immediately 

initiated using existing facilities...to substantiate the feasibility of this type 

of combustion.

The panelists nevertheless gave thumbs-up to the Aerospaceplane effort as a continuing program of research. Their report urged a broadening of topics, placing 

greater emphasis on scramjets, structures and materials, and two-stage-to-orbit configurations. The array of proposed engines were “all sufficiently interesting so that 

research on all of them should be continued and emphasized.”65

As the studies went forward in the wake of this review, new propulsion concepts 

continued to flourish. Lockheed was in the forefront. This firm had initiated company-funded work during the spring of 1959 and had a well-considered single-stage 

concept two years later. An artist’s rendering showed nine separate rocket nozzles at 

its tail. The vehicle also mounted four ramjets, set in pods beneath the wings.

Convair’s Space Plane had used separated nitrogen as a propellant, heating it in 

the LACE precooler and allowing it to expand through a nozzle to produce thrust. 

Lockheed’s Aerospace Plane turned this nitrogen into an important system element, 

with specialized nitrogen rockets delivering 125,000 pounds of thrust. This certainly did not overcome the drag produced by air collection, which would have 

turned the vehicle into a perpetual motion machine. However, the nitrogen rockets 

made a valuable contribution.66

Participants in the new studies took to the work with enthusiasm matching that 

of Arthur Thomas at Marquardt. Robert Sanator, a colleague of Kartveli at Republic, recalls the excitement: “This one had everything. There wasn’t a single thing in 

it that was off-the-shelf. Whatever problem there was in aerospace—propulsion, 

materials, cooling, aerodynamics—Aerospaceplane had it. It was a lifetime work 

and it had it all. I naturally jumped right in.”63

Aerospaceplane also drew attention from the Air Force’s Scientific Advisory 

Board, which set up an ad hoc committee to review its prospects. Its chairman, 

Alexander Flax, was the Air Force’s chief scientist. Members specializing in propulsion included Ferri, along with Seymour Bogdonoff of Princeton University, a leading experimentalist; Perry Pratt of Pratt & Whitney, who had invented the twinspool turbojet; NASA’s Alfred Eggers; and the rocket specialist George P. Sutton. 

There also were hands-on program managers: Robert Widmer of Convair, builder 

of the Mach 2 B-58 bomber; Harrison Storms of North American, who had shaped 

the X-15 and the Mach 3 XB-70 bomber.64

This all-star group came away deeply skeptical of the prospects for Aerospaceplane. Its report, issued in December 1960, addressed a number of points and gave 

an overall assessment:

The proposed designs for Aerospace Plane...appear to violate no physical 

principles, but the attractive performance depends on an estimated 

combination of optimistic assumptions for the performance of components 

and subsystems. There are practically no experimental data which support 

these assumptions.

On LACE and ACES:

We consider the estimated LACE-ACES performance very optimistic. In 

several cases complete failure of the project would result from any significant 

performance degradation from the present estimates.... Obviously the 

advantages claimed for the system will not be available unless air can be 

condensed and purified very rapidly during flight. The figures reported 

indicate that about 0.8 ton of air per second would have to be processed. 

In conventional, i.e., ordinary commercial equipment, this would require 

a distillation column having a cross section on the order of 500 square 

feet.... It is proposed to increase the capacity of equipment of otherwise 

conventional design by using centrifugal force. This may be possible, but as 

far as the Committee knows this has never been accomplished.
118

Facing the Heat Barrier: A History of Hypersonics

119

First Thoughts of Hypersonic Propulsion

For takeoff, Lockheed expected to use Turbo-LACE. This was a LACE variant 

that sought again to reduce the inherently hydrogen-rich operation of the basic 

system. Rather than cool the air until it was liquid, Turbo-Lace chilled it deeply but 

allowed it to remain gaseous. Being very dense, it could pass through a turbocompressor and reach pressures in the hundreds of psi. This saved hydrogen because less 

was needed to accomplish this cooling. The Turbo-LACE engines were to operate 

at chamber pressures of 200 to 250 psi, well below the internal pressure of standard 

rockets but high enough to produce 300,000 pounds of thrust by using turbocompressed oxygen.67

Republic Aviation continued to emphasize the scramjet. A new configuration 

broke with the practice of mounting these engines within pods, as if they were 

turbojets. Instead, this design introduced the important topic of engine-airframe 

integration by setting forth a concept that amounted to a single enormous scramjet 

fitted with wings and a tail. A conical forward fuselage served as an inlet spike. The 

inlets themselves formed a ring encircling much of the vehicle. Fuel tankage filled 

most of its capacious internal volume.

This design study took two views regarding the potential performance of its 

engines. One concept avoided the use of LACE or ACES, assuming again that this 

craft could scram all the way to orbit. Still, it needed engines for takeoff so turboramjets were installed, with both Pratt & Whitney and General Electric providing 

candidate concepts. Republic thus was optimistic at high Mach but conservative at 

low speed.

Republic’s Aerospaceplane concept showed extensive engine-airframe integration. 

(Republic Aviation)

Lockheed’s Aerospaceplane concept. The alternate hypersonic in-flight refueling system approach 

called for propellant transfer at Mach 6. (Art by Dennis Jenkins)
120

Facing the Heat Barrier: A History of Hypersonics

121

First Thoughts of Hypersonic Propulsion

hydrogen from tanks. It showed stable combustion, delivering thrust as high as 

5,700 pounds.72

Within the Air Force, the SAB’s Ad Hoc Committee on Aerospaceplane continued to provide guidance along with encouraging words. A review of July 1962 was 

less skeptical in tone than the one of 18 months earlier, citing “several attractive 

arguments for a continuation of this program at a significant level of funding”:

It will have the military advantages that accrue from rapid response times 

and considerable versatility in choice of landing area. It will have many of 

the advantages that have been demonstrated in the X-15 program, namely, 

a real pay-off in rapidly developing reliability and operational pace that 

comes from continuous re-use of the same hardware again and again. It 

may turn out in the long run to have a cost effectiveness attractiveness...

the cost per pound may eventually be brought to low levels. Finally, the 

Aerospaceplane program will develop the capability for flights in the 

atmosphere at hypersonic speeds, a capability that may be of future use to 

the Defense Department and possibly to the airlines.73

Single-stage-to-orbit (SSTO) was on the agenda, a topic that merits separate 

comment. The space shuttle is a stage-and-a-half system; it uses solid boosters plus a 

main stage, with all engines burning at liftoff. It is a measure of progress, or its lack, 

in astronautics that the Soviet R-7 rocket that launched the first Sputniks was also 

stage-and-a-half.74 The concept of SSTO has tantalized designers for decades, with 

these specialists being highly ingenious and ready to show a can-do spirit in the face 

of challenges.

This approach certainly is elegant. It also avoids the need to launch two rockets 

to do the work of one, and if the Earth’s gravity field resembled that of Mars, SSTO 

would be the obvious way to proceed. Unfortunately, the Earth’s field is considerably stronger. No SSTO has ever reached orbit, either under rocket power or by 

using scramjets or other airbreathers. The technical requirements have been too 

severe.

The SAB panel members attended three days of contractor briefings and reached 

a firm conclusion: “It was quite evident to the Committee from the presentation of 

nearly all the contractors that a single stage to orbit Aerospaceplane remains a highly 

speculative effort.” Reaffirming a recommendation from its 1960 review, the group 

urged new emphasis on two-stage designs. It recommended attention to “development of hydrogen fueled turbo ramjet power plants capable of accelerating the first 

stage to Mach 6.0 to 10.0.... Research directed toward the second stage which 

will ultimately achieve orbit should be concentrated in the fields of high pressure 

hydrogen rockets and supersonic burning ramjets and air collection and enrichment 

systems.”75

The other design introduced LACE and ACES both for takeoff and for final 

ascent to orbit and made use of yet another approach to derichening the hydrogen. 

This was SuperLACE, a concept from Marquardt that placed slush hydrogen rather 

than standard liquid hydrogen in the main tank. The slush consisted of liquid that 

contained a considerable amount of solidified hydrogen. It therefore stood at the 

freezing point of hydrogen, 14 K, which was markedly lower than the 21 K of liquid 

hydrogen at the boiling point.68

SuperLACE reduced its use of hydrogen by shunting part of the flow, warmed 

in the LACE heat exchanger, into the tank. There it mixed with the slush, chilling 

again to liquid while melting some of the hydrogen ice. Careful control of this flow 

ensured that while the slush in the tank gradually turned to liquid and rose toward 

the 21 K boiling point, it did not get there until the air-collection phase of a flight 

was finished. As an added bonus, the slush was noticeably denser than the liquid, 

enabling the tank to hold more fuel.69

LACE and ACES remained in the forefront, but there also was much interest 

in conventional rocket engines. Within the Aerospaceplane effort, this approach 

took the name POBATO, Propellants On Board At Takeoff. These rocket-powered 

vehicles gave points of comparison for the more exotic types that used LACE and 

scramjets, but here too people used their imaginations. Some POBATO vehicles 

ascended vertically in a classic liftoff, but others rode rocket sleds along a track while 

angling sharply upward within a cradle.70

In Denver, the Martin Company took rocket-powered craft as its own, for this 

firm expected that a next-generation launch vehicle of this type could be ready far 

sooner than one based on advanced airbreathing engines. Its concepts used vertical 

liftoff, while giving an opening for the ejector rocket. Martin introduced a concept 

of its own called RENE, Rocket Engine Nozzle Ejector (RENE), and conducted 

experiments at the Arnold Engineering Development Center. These tests went forward during 1961, using a liquid rocket engine, with nozzle of 5-inch diameter set 

within a shroud of 17-inch width. Test conditions corresponded to flight at Mach 

2 and 40,000 feet, with the shrouds or surrounding ducts having various lengths 

to achieve increasingly thorough mixing. The longest duct gave the best performance, increasing the rated 2,000-pound thrust of the rocket to as much as 3,100 

pounds.71

A complementary effort at Marquardt sought to demonstrate the feasibility of 

LACE. The work started with tests of heat exchangers built by Garrett AiResearch 

that used liquid hydrogen as the working fluid. A company-made film showed dark 

liquid air coming down in a torrent, as seen through a porthole. Further tests used 

this liquefied air in a small thrust chamber. The arrangement made no attempt to 

derichen the hydrogen flow; even though it ran very fuel-rich, it delivered up to 275 

pounds of thrust. As a final touch, Marquardt crafted a thrust chamber of 18-inch 

diameter and simulated LACE operation by feeding it with liquid air and gaseous 
122

Facing the Heat Barrier: A History of Hypersonics

123

First Thoughts of Hypersonic Propulsion

oxidizer prior to separation, but to reach from Mach 3 to orbital speed, the second 

stage had to be simple indeed. Steinhoff envisioned a long vehicle resembling a torpedo, powered by hydrogen-burning rockets but lacking wings and thermal protection. It was not reusable and would not reenter, but it would be piloted. A project 

report stated, “Crew recovery is accomplished by means of a reentry capsule of the 

Gemini-Apollo class. The capsule forms the nose section of the vehicle and serves as 

the crew compartment for the entire vehicle.”78

ROLS appears in retrospect as a mirror image of NASA’s eventual space shuttle, 

which adopted a technically simple booster—a pair of large solid-propellant rockets—while packaging the main engines and most other costly systems within a fullyrecoverable orbiter. By contrast, ROLS used a simple second stage and a highly 

intricate first stage, in the form of a large delta-wing airplane that mounted eight 

turbojet engines. Its length of 335 feet was more than twice that of a B-52. Weighing 825,000 pounds at takeoff, ROLS was to deliver a payload of 30,000 pounds 

to orbit.79

Such two-stage concepts continued to emphasize ACES, while still offering a role 

for LACE. Experimental test and development of these concepts therefore remained 

on the agenda, with Marquardt pursuing further work on LACE. The earlier tests, 

during 1960 and 1961, had featured an off-the-shelf thrust chamber that had seen 

use in previous projects. The new work involved a small LACE engine, the MA117, 

that was designed from the start as an integrated system.

LACE had a strong suit in its potential for a very high specific impulse, Isp. This 

is the ratio of thrust to propellant flow rate and has dimensions of seconds. It is a 

key measure of performance, is equivalent to exhaust velocity, and expresses the 

engine’s fuel economy. Pratt & Whitney’s RL10, for instance, burned hydrogen and 

oxygen to give thrust of 15,000 pounds with an Isp of 433 seconds.80 LACE was an 

airbreather, and its Isp could be enormously higher because it took its oxidizer from 

the atmosphere rather than carrying it in an onboard tank. The term “propellant 

flow rate” referred to tanked propellants, not to oxidizer taken from the air. For 

LACE this meant fuel only.

The basic LACE concept produced a very fuel-rich exhaust, but approaches 

such as RENE and SuperLACE promised to reduce the hydrogen flow substantially. Indeed, such concepts raised the prospect that a LACE system might use an 

optimized mixture ratio of hydrogen and oxidizer, with this ratio being selected to 

give the highest Isp. The MA117 achieved this performance artificially by using a 

large flow of liquid hydrogen to liquefy air and a much smaller flow for the thrust 

chamber. Hot-fire tests took place during December 1962, and a company report 

stated that “the system produced 83% of the idealized theoretical air flow and 81% 

of the idealized thrust. These deviations are compatible with the simplifications of 

the idealized analysis.”81

Convair, home of Space Plane, had offered single-stage configurations as early 

as 1960. By 1962 its managers concluded that technical requirements placed such 

a vehicle out of reach for at least the next 20 years. The effort shifted toward a 

two-stage concept that took form as the 1964 Point Design Vehicle. With a gross 

takeoff weight of 700,000 pounds, the baseline approach used turboramjets to reach 

Mach 5. It cruised at that speed while using ACES to collect liquid oxygen, then 

accelerated anew using ramjets and rockets. Stage separation occurred at Mach 8.6 

and 176,000 feet, with the second stage reaching orbit on rocket power. The payload was 23,000 pounds with turboramjets in the first stage, increasing to 35,000 

pounds with the more speculative SuperLACE.

The documentation of this 1964 Point Design, filling 16 volumes, was issued 

during 1963. An important advantage of the two-stage approach proved to lie in 

the opportunity to optimize the design of each stage for its task. The first stage was 

a Mach 8 aircraft that did not have to fly to orbit and that carried its heavy wings, 

structure, and ACES equipment only to staging velocity. The second-stage design 

showed strong emphasis on re-entry; it had a blunted shape along with only modest 

requirements for aerodynamic performance. Even so, this Point Design pushed the 

state of the art in materials. The first stage specified superalloys for the hot underside 

along with titanium for the upper surface. The second stage called for coated refractory metals on its underside, with superalloys and titanium on its upper surfaces.76

Although more attainable than its single-stage predecessors, the Point Design 

still relied on untested technologies such as ACES, while anticipating use in aircraft 

structures of exotic metals that had been studied merely as turbine blades, if indeed 

they had gone beyond the status of laboratory samples. The opportunity nevertheless existed for still greater conservatism in an airbreathing design, and the man 

who pursued it was Ernst Steinhoff. He had been present at the creation, having 

worked with Wernher von Braun on Germany’s wartime V-2, where he headed up 

the development of that missile’s guidance. After 1960 he was at the Rand Corporation, where he examined Aerospaceplane concepts and became convinced that 

single-stage versions would never be built. He turned to two-stage configurations 

and came up with an outline of a new one: ROLS, the Recoverable Orbital Launch 

System. During 1963 he took the post of chief scientist at Holloman Air Force Base 

and proceeded to direct a formal set of studies.77

The name of ROLS had been seen as early as 1959, in one of the studies that 

had grown out of SR-89774, but this concept was new. Steinhoff considered that 

the staging velocity could be as low as Mach 3. At once this raised the prospect that 

the first stage might take shape as a modest technical extension of the XB-70, a large 

bomber designed for flight at that speed, which at the time was being readied for 

flight test. ROLS was to carry a second stage, dropping it from the belly like a bomb, 

with that stage flying on to orbit. An ACES installation would provide the liquid 
124

Facing the Heat Barrier: A History of Hypersonics

125

First Thoughts of Hypersonic Propulsion

The double arrows indicate reversibility. The oxidation reactions were exothermic, occurring at approximately 1,600°F for barium and 1,800°F for cobalt. The 

reduction reactions, which released the oxygen, were endothermic, allowing the 

oxides to cool as they yielded this gas.

Dynatech’s separator unit consisted of a long rotating drum with its interior 

divided into four zones using fixed partitions. A pebble bed of oxide-coated particles 

lined the drum interior; containment screens held the particles in place while allowing the drum to rotate past the partitions with minimal leakage. The zones exposed 

the oxide alternately to high-pressure ram air for oxidation and to low pressure for 

reduction. The separation was to take place in flight, at speeds of Mach 4 to Mach 5, 

but an inlet could slow the internal airflow to as little as 50 feet per second, increasing the residence time of air within a unit. The company proposed that an array of 

such separators weighing just under 10 tons could handle 2,000 pounds per second 

of airflow while producing liquid oxygen of 65 percent purity.85

Ten tons of equipment certainly counts within a launch vehicle, even though it 

included the weight of the oxygen liquefaction apparatus. Still it was vastly lighter 

than the alternative: the rotating distillation system. The Linde Division of Union 

Carbide pursued this approach. Its design called for a cylindrical tank containing 

the distillation apparatus, measuring nine feet long by nine feet in diameter and 

rotating at 570 revolutions per minute. With a weight of 9,000 pounds, it was to 

process 100 pounds per second of liquefied air—which made it 10 times as heavy 

as the Dynatech system, per pound of product. The Linde concept promised liquid 

oxygen of 90 percent purity, substantially better than the chemical system could 

offer, but the cited 9,000-pound weight left out additional weight for the LACE 

equipment that provided this separator with its liquefied air.86

A study at Convair, released in October 1963, gave a clear preference to the 

Dynatech concept. Returning to the single-stage Space Plane of prior years, Convair 

engineers considered a version with a weight at takeoff of 600,000 pounds, using 

either the chemical or the distillation ACES. The effort concluded that the Dynatech 

separator offered a payload to orbit of 35,800 using barium and 27,800 pounds 

with cobalt. The Linde separator reduced this payload to 9,500 pounds. Moreover, 

because it had less efficiency, it demanded an additional 31,000 pounds of hydrogen 

fuel, along with an increase in vehicle volume of 10,000 cubic feet.87

The turn toward feasible concepts such as ROLS, along with the new emphasis 

on engineering design and test, promised a bright future for Aerospaceplane studies. 

However, a commitment to serious research and development was another matter. 

Advanced test facilities were critical to such an effort, but in August 1963 the Air 

Force canceled plans for a large Mach 14 wind tunnel at AEDC. This decision gave 

a clear indication of what lay ahead.88

A year earlier Aerospaceplane had received a favorable review from the SAB Ad 

Hoc Committee. The program nevertheless had its critics, who existed particularly 

The best performance run delivered 0.783 pounds per second of liquid air, 

which burned a flow of 0.0196 pounds per second of hydrogen. Thrust was 73 

pounds; Isp reached 3,717 seconds, more than eight times that of the RL10. Tests 

of the MA117 continued during 1963, with the best measured values of Isp topping 

4,500 seconds.82

In a separate effort, the Marquardt manager Richard Knox directed the preliminary design of a much larger LACE unit, the MA116, with a planned thrust of 

10,000 pounds. On paper, it achieved substantial derichening by liquefying only 

one-fifth of the airflow and using this liquid air in precooling, while deeply cooling 

the rest of the airflow without liquefaction. A turbocompressor then was to pump 

this chilled air into the thrust chamber. A flow of less than four pounds per second 

of liquid hydrogen was to serve both as fuel and as primary coolant, with the anticipated Isp exceeding 3,000 seconds.83

New work on RENE also flourished. The Air Force had a cooperative agreement with NASA’s Marshall Space Flight Center, where Fritz Pauli had developed a 

subscale rocket engine that burned kerosene with liquid oxygen for a thrust of 450 

pounds. Twelve of these small units, mounted to form a ring, gave a basis for this 

new effort. The earlier work had placed the rocket motor squarely along the centerline of the duct. In the new design, the rocket units surrounded the duct, leaving 

it unobstructed and potentially capable of use as an ejector ramjet. The cluster was 

tested successfully at Marshall in September 1963 and then went to the Air Force’s 

AEDC. As in the RENE tests of 1961, the new configuration gave a thrust increase 

of as much as 52 percent.84

While work on LACE and ejector rockets went forward, ACES stood as a particularly critical action item. Operable ACES systems were essential for the practical 

success of LACE. Moreover, ACES had importance distinctly its own, for it could 

provide oxidizer to conventional hydrogen-burning rocket engines, such as those 

of ROLS. As noted earlier, there were two techniques for air separation: by chemical methods and through use of a rotating fractional distillation apparatus. Both 

approaches went forward, each with its own contractor.

In Cambridge, Massachusetts, the small firm of Dynatech took up the challenge 

of chemical separation, launching its effort in May 1961. Several chemical reactions appeared plausible as candidates, with barium and cobalt offering particular 

promise:

2BaO2

 →

← 2BaO + O2

2Co3O4 →

← 6CoO + O2
126

Facing the Heat Barrier: A History of Hypersonics

127

First Thoughts of Hypersonic Propulsion

The design concepts of that era were meant to offer glimpses of possible futures, 

but for this Astrorocket, the future was only seven years off. It clearly foreshadowed 

a class of two-stage fully reusable space shuttles, fitted with delta wings, that came to 

the forefront in NASA-sponsored studies of 1971. The designers at Martin were not 

clairvoyant; they drew on the background of Dyna-Soar and on studies at NASAAmes of winged re-entry vehicles. Still, this concept demonstrated that some design 

exercises were returning to the mainstream.92

Further work on ACES also proceeded, amid unfortunate results at Dynatech. 

That company’s chemical separation processes had depended for success on having a 

very large area of reacting surface within the pebble-bed air separators. This appeared 

achievable through such means as using finely divided oxide powders or porous 

particles impregnated with oxide. But the research of several years showed that the 

oxide tended to sinter at high temperatures, markedly diminishing the reacting surface area. This did not make chemical separation impossible, but it sharply increased 

the size and weight of the equipment, which robbed this approach of its initially 

strong advantage over the Linde distillation system. This led to abandonment of 

Dynatech’s approach.93

Linde’s system was heavy and drastically less elegant than Dynatech’s alternative, but it amounted largely to a new exercise in mechanical engineering and went 

forward to successful completion. A prototype operated in test during 1966, and 

Martin’s Astrorocket. (U.S. Air Force)

within the SAB’s Aerospace Vehicles and Propulsion panels. In October 1963 they 

issued a report that dealt with proposed new bombers and vertical-takeoff-andlanding craft, as well as with Aerospaceplane, but their view was unmistakable on 

that topic:

The difficulties the Air Force has encountered over the past three years in 

identifying an Aerospaceplane program have sprung from the facts that 

the requirement for a fully recoverable space launcher is at present only 

vaguely defined, that today’s state-of-the-art is inadequate to support any 

real hardware development, and the cost of any such undertaking will be 

extremely large.... [T]he so-called Aerospaceplane program has had such 

an erratic history, has involved so many clearly infeasible factors, and has 

been subject to so much ridicule that from now on this name should be 

dropped. It is also recommended that the Air Force increase the vigilance 

that no new program achieves such a difficult position.89

Aerospaceplane lost still more of its rationale in December, as Defense Secretary 

Robert McNamara canceled Dyna-Soar. This program was building a mini-space 

shuttle that was to fly to orbit atop a Titan III launch vehicle. This craft was well 

along in development at Boeing, but program reviews within the Pentagon had 

failed to find a compelling purpose. McNamara thus disposed of it.90

Prior to this action, it had been possible to view Dyna-Soar as a prelude to operational vehicles of that general type, which might take shape as Aerospaceplanes. The 

cancellation of Dyna-Soar turned the Aerospaceplane concept into an orphan, a 

long-term effort with no clear relation to anything currently under way. In the wake 

of McNamara’s decision, Congress deleted funds for further Aerospaceplane studies, 

and Defense Department officials declined to press for its restoration within the FY 

1964 budget, which was under consideration at that time. The Air Force carried 

forward with new conceptual studies of vehicles for both launch and hypersonic 

cruise, but these lacked the focus on advanced airbreathing propulsion that had 

characterized Aerospaceplane.91

There nevertheless was real merit to some of the new work, for this more realistic 

and conservative direction pointed out a path that led in time toward NASA’s space 

shuttle. The Martin Company made a particular contribution. It had designed no 

Aerospaceplanes; rather, using company funding, its technical staff had examined 

concepts called Astrorockets, with the name indicating the propulsion mode. Scramjets and LACE won little attention at Martin, but all-rocket vehicles were another 

matter. A concept of 1964 had a planned liftoff weight of 1,250 tons, making it 

intermediate in size between the Saturn I-B and Saturn V. It was a two-stage fullyreusable configuration, with both stages having delta wings and flat undersides. 

These undersides fitted together at liftoff, belly to belly.
128

Facing the Heat Barrier: A History of Hypersonics

129

First Thoughts of Hypersonic Propulsion

1 See, for instance, Emme, History; Ley, Rockets; McDougall, Heavens; Neufeld, Ballistic; NASA 

SP-4206; and Heppenheimer, Countdown.

2 NASA SP-4404; SP-4221, pp. 105-08, 423-25.

3 Johns Hopkins APL Technical Digest, Vol. 13, No. 1 (1992), p. 57; Lay, Thermodynamics, pp. 

574-76.

4 Miller, X-Planes, ch. 11.

5 Ibid., pp. 116-17. X-2 record: NASA SP-4303, p. 316.

6 Peterson, “Evaluation.” 4130 steel: Miller, X-Planes, p. 119.

7 Ritchie, “Evaluation.”

8 Miller, X-Planes, p. 117; Crickmore, SR-71, pp. 154-55.

9 “Standard Missile Characteristics: IM-99A Bomarc.” Its name: Cornett, “Overview,” p. 109. 

Extended range: Jane’s, 1967-1968, p. 490.

10 “Development of Bomarc,” pp. 10-12; Bagwell, “History,” pp. 12-13.

11 Bagwell, “History,” p. 17 (includes quotes); Pfeifer, “Bomarc,” p. 1; Report Boeing D-11532 

(Boeing), pp. 43-44.

12 Jane’s, 1960-61, p. 447; Cornett, “Overview,” pp. 118, 124-25.

13 Johns Hopkins APL Technical Digest, Vol. 3, No. 2 (1982), pp. 117-22; Jane’s, 1960-61, p. 446.

14 “Standard Missile Characteristics: XSM-64 Navaho”; Gibson, Navaho, pp. 35-36, 45-48.

15 Gibson, Navaho, pp. 63-77; “Development of the SM-64,” 1954-58, pp. 100-07. Author interview, J. Leland Atwood, 18 July 1988. Folder 18649, NASA Historical Reference Collection, 

NASA History Division, Washington, D.C. 20546.

16 Air Enthusiast, July-September 1978, pp. 198-213 (quote, p. 213). XF-103, XF-104: Gunston, 

Fighters, pp. 121-22, 193. 

17 Author interview, Frederick Billig, 27 June 1987. Folder 18649, NASA Historical Reference 

Collection, NASA History Division, Washington, D.C. 20546. See also CASI 88N-12321,

p. 12-6.

18 Fletcher, “Supersonic”; NACA: TN 2206, RM E51K26.

19 Fletcher, “Supersonic.”

20 NACA TN 4386 (quote, p. 22).

21 Fletcher, “Supersonic,” p. 737.

22 Johns Hopkins APL Technical Digest, Vol. 11, Nos. 3 and 4 (1990), pp. 319-35 (quotes,

pp. 328-29).

23 “Statement Regarding the Military Service of Antonio Ferri”; Kaufman et al., Moe Berg, pp. 

182-88 (quote, p. 186).

24 Journal of Aircraft, January-February 1968, p. 3; Journal of the Royal Aeronautical Society. September 1964, p. 575; Johns Hopkins Magazine, December 1988 p. 53; author interview, Louis 

Nucci, 24 June 1987. 

25 Author interviews: John Erdos, 22 July 1985. Folder 18649, NASA Historical Reference Collection, NASA History Division, Washington, D.C. 20546; Robert Sanator, 1 February 1988. 

NASA Historical Reference Collection, NASA History Division, Washington, D.C. 20546.

while limits to the company’s installed power capacity prevented the device from 

processing the rated flow of 100 pounds of air per second, it handled 77 pounds per 

second, yielding a product stream of oxygen that was up to 94 percent pure. Studies 

of lighter-weight designs also proceeded. In 1969 Linde proposed to build a distillation air separator, rated again at 100 pounds per second, weighing 4,360 pounds. 

This was only half the weight allowance of the earlier configuration.94

In the end, though, Aerospaceplane failed to identify new propulsion concepts 

that held promise and that could be marked for mainstream development. The 

program’s initial burst of enthusiasm had drawn on a view that the means were in 

hand, or soon would be, to leap beyond the liquid-fuel rocket as the standard launch 

vehicle and to pursue access to orbit using methods that were far more advanced. 

The advent of the turbojet, which had swiftly eclipsed the piston engine, was on 

everyone’s mind. Yet for all the ingenuity behind the new engine concepts, they 

failed to deliver. What was worse, serious technical review gave no reason to believe 

that they could deliver.

In time it would become clear that hypersonics faced a technical wall. Only 

limited gains were achievable in airbreathing propulsion, with single-stage-to-orbit 

remaining out of reach and no easy way at hand to break through to the really 

advanced performance for which people hoped.
130

Facing the Heat Barrier: A History of Hypersonics

131

First Thoughts of Hypersonic Propulsion

53 Jenkins, Space Shuttle, p. 54; Convair ZP-M-095 (quote, p. ix).

54 Los Angeles Times, 15 January 1961, pp. 1, 3.

55 Republic Aviation News, 9 September 1960, pp. 1, 5.

56 Ibid. F-105: Gunston, Fighters, pp. 194-200.

57 Republic Aviation News, 9 September 1960, pp. 1, 5.

58 Author interview, Colonel Edward Hall, 25 January 1989. Folder 18649, NASA Historical Reference Collection, NASA History Division, Washington, D.C. 20546. Also Hallion, Hypersonic, 

p. 751.

59 DTIC AD-342992, p. i. Nine of these 11 reports are available from DTIC: AD-342992 

through AD-343000.

60 Aviation Week, 31 October 1960, p. 26.

61 Los Angeles Times, 3 November 1960, p. 3A.

62 Los Angeles Times, 15 January 1961, pp. 1, 3.

63 Author interview, Robert Sanator, February 1988. Folder 18649, NASA Historical Reference 

Collection, NASA History Division, Washington, D.C. 20546.

64 “Memorandum of the Scientific Advisory Board Ad Hoc Committee on Aerospace Plane,” 

December 1960.

65 Ibid. (includes all quotes.)

66 Los Angeles Times, 3 November 1960, p. 3A; Lockheed LAC 571375, cover and chart 6.

67 Lockheed LAC 571375, chart 6. For Turbo-LACE, see Heppenheimer, Hypersonic, p. 165.

68 Jenkins, Space Shuttle, pp. 55-57; DTIC ADC-053100, AD-351039 through AD-351041, AD351581 through AD-351584.

69 AIAA Paper 86-1680.

70 Jenkins, Space Shuttle, p. 56.

71 Martin M-63-1 (RENE, pp. IV-2 to IV-5).

72 DTIC AD-318628, AD-322199; film, “Liquid Air Cycle Engine” (Marquardt, March 1961).

73 “Report of the Scientific Advisory Board Ad Hoc Committee on Aerospaceplane,” 23-25 July 

1962.

74 NASA SP-2000-4408, p. 131; Heppenheimer, Countdown, pp. 110, 119.

75 “Report of the Scientific Advisory Board Ad Hoc Committee on Aerospaceplane,” 23-25 July 

1962 (includes quotes).

76 AIAA Paper 92-3499, p. 4.

77 Neufeld, Rocket, p. 101; Cornett, “History,” pp. 1-3.

78 Hallion, Hypersonic, p. 949; Cornett, “History,” pp. 2, 7-8. Quote: DTIC AD-359545, p. 11.

79 Cornett, “History,” pp. 7, 11.

80 For RL10, see NASA SP-4206, p. 139.

81 DTIC AD-335212, pp. 5-7 (quote, p. 1).

82 Ibid., pp. 18-19; CASI 75N-75668, pp. ii, 15.

83 DTIC AD-350952, AD-384302 (performance, p. 2).

26 Author interview, Louis Nucci, 24 June 1988 (includes quote). Folder 18649, NASA Historical 

Reference Collection, NASA History Division, Washington, D.C. 20546.

27 ISABE Paper 97-7004, p. 1.

28 Quote: Journal of the Royal Aeronautical Society, September 1964, p. 577. 

29 ISABE Paper 97-7004, p. 2. Quote: Astronautics & Aeronautics, August 1964, p. 33.

30 Shapiro, Compressible, pp. 198-99.

31 Ferri and Fox, “Analysis.”

32 Quotes: Astronautics & Aeronautics, August 1964, p. 33; Journal of the Royal Aeronautical Society, 

September 1964, p. 577. 

33 Shapiro, Compressible, pp. 147-51; Lockheed Horizons, Winter 1981/1982, 11-12; Crickmore, 

SR-71, pp. 26, 94-95.

34 Author interview, James Eastham, 1 May 1988. Folder 18649, NASA Historical Reference Collection, NASA History Division, Washington, D.C. 20546.

35 Ibid.

36 Crickmore, SR-71, pp. 125-26.

37 Ferri, “Possible Directions”; Journal of the Royal Aeronautical Society, September 1964, pp. 575-

97 (quotes, pp. 595, 597). 

38 Ferri and Fox, “Analysis” (quotes, p. 1111). Arthur Thomas quotes: Arthur Thomas, author 

interview, 24 September 1987. Folder 18649, NASA Historical Reference Collection, History 

Division, Washington, D.C. 20546.

39 Journal of Spacecraft and Rockets, September 1968, pp. 1076-81. Quote: Fred Billig interview, 

16 October 1987. Folder 18649, NASA Historical Reference Collection, NASA History Division, Washington, D.C. 20546. 

40 Air Enthusiast, July-September 1978, p. 210.

41 Heppenheimer, First Flight, pp. 63, 142.

42 Gunston, Fighters, p. 121; Crickmore, SR-71, pp. 94-95.

43 U.S. Patents 2,735,263 (Charshafian), 2,883,829 (Africano), 3,172,253 (Schelp), 3,525,474 

(Von Ohain). For Ohain and Schelp, see Schlaifer and Heron, Development, pp. 377-79,

383-86.

44 Crickmore, SR-71, pp. 94-95.

45 U.S. Patents 2,922,286, 3,040,519, 3,040,520 (all to Rae).

46 AIAA Paper 86-1680.

47 AIAA Paper 92-3499.

48 DTIC AD-351239.

49 Author interview, Arthur Thomas, 24 September 1987. Folder 18649, NASA Historical Reference Collection, NASA History Division, Washington, D.C. 20546.

50 Johns Hopkins APL Technical Digest, Vol. 11, Nos. 3 and 4 (1990), pp. 332-33; Boeing 707s: 

Pedigree, pp. 55-57.

51 Jenkins, Space Shuttle, p. 52; DTIC AD-372320.

52 “Reusable Space Launch Vehicle Systems Study” (Convair); Convair Report GD/C-DCJ 65-

004.
132

Facing the Heat Barrier: A History of Hypersonics

84 AIAA Paper 99-4896, p. 4; DTIC AD-359763, p. II-2.

85 DTIC AD-336169, pp. 1, 13, 26-28.

86 DTIC AD-351207, pp. iii, 3-4.

87 DTIC AD-345178, pp. 22, 397.

88 Cornett, “History,” p. 2.

89 “Report of the USAF Scientific Advisory Board Aerospace Vehicles/Propulsion Panels on Aerospaceplane, VTOL, and Strategic Manned Aircraft,” 24 October 1963; extended quote also in 

Hallion, Hypersonic, p. 951.

90 NASA SP-4221, pp. 49-54. For general overviews of Dyna-Soar, see Hallion, Hypersonic, Case 

II; Miller, X-Planes, ch. 24; DTIC ADA-303832.

91 Hallion, Hypersonic, pp. 951-52; Jenkins, Space Shuttle, pp. 66-67; Cornett, “History,” p. 4.

92 Hallion, Hypersonic, pp. 952-54; Jenkins, Space Shuttle, p. 62.

93 DTIC: AD-353898, AD-461220.

94 1966 work: AIAA Paper 92-3499, pp. 5-6; DTIC AD-381504, p. 10. 1969 proposal: DTIC 

AD-500489, p. iv.
The classic spaceship has wings, and throughout much of the 1950s both NACA 

and the Air Force struggled to invent such a craft. Design studies addressed issues 

as fundamental as whether this hypersonic rocket plane should have one particular 

wing-body configuration, or whether it should be upside down. The focus of the 

work was Dyna-Soar, a small version of the space shuttle that was to ride to orbit 

atop a Titan III. It brought remarkable engineering advances, but Pentagon policy 

makers, led by Defense Secretary Robert McNamara, saw it as offering little more 

than technical development, with no mission that could offer a military justification. In December 1963 he canceled it.

Better prospects attended NASA’s effort in manned spaceflight, which culminated in the Apollo piloted flights to the Moon. Apollo used no wings; rather, it 

relied on a simple cone that used the Allen-Eggers blunt-body principle. Still, its 

demands were stringent. It had to re-enter successfully with twice the energy of an 

entry from Earth orbit. Then it had to navigate a corridor, a narrow range of altitudes, to bleed off energy without either skipping back into space or encountering 

g-forces that were too severe. By doing these things, it showed that hypersonics was 

ready for this challenge.

Winged Spacecraft and Dyna-Soar

Boost-glide rockets, with wings, entered the realm of advanced conceptual design 

with postwar studies at Bell Aircraft called Bomi, Bomber Missile. The director of 

the work, Walter Dornberger, had headed Germany’s wartime rocket development 

program and had been in charge of the V-2. The new effort involved feasibility 

studies that sought to learn what might be done with foreseeable technology, but 

Bomi was a little too advanced for some of Dornberger’s colleagues. Historian Roy 

Houchin writes that when Dornberger faced “abusive and insulting remarks” from 

an Air Force audience, he responded by declaring that his Bomi would be receiving 

more respect if he had had the chance to fly it against the United States during the 

war. In Houchin’s words, “The silence was deafening.”1

133

Widening Prospects

for Re-entry

5
134

Facing the Heat Barrier: A History of Hypersonics

135

Widening Prospects for Re-entry

range. Basic feasibility then lay even farther in the future, but the Air Force’s interest in the Atlas ICBM meant that it wanted missiles of longer range, even though 

shorter-range designs could be available sooner. An intercontinental Bomi at least 

could be evaluated as a potential alternative to Atlas, and it might find additional 

roles such as strategic reconnaissance.3

In April 1954, with that ICBM very much in the ascendancy, WADC awarded 

Bell its desired study contract. Bomi now had an Air Force designation, MX-2276. 

Bell examined versions of its two-stage concept with 4,000- and 6,000-mile ranges 

while introducing a new three-stage configuration with the stages mounted bellyto-back. Liftoff thrust was to be 1.2 million pounds, compared with 360,000 for 

the three-engine Atlas. Bomi was to use a mix of liquid oxygen and liquid fluorine, 

the latter being highly corrosive and hazardous, whereas Atlas needed only liquid 

oxygen, which was much safer. The new Bomi was to reach 22,000 feet per second, 

slightly less than Atlas, but promised a truly global glide range of 12,000 miles. Even 

so, Atlas clearly was preferable.4

But the need for reconnaissance brought new life to the Bell studies. At WADC, 

in parallel with initiatives that were sparking interest in unpiloted reconnaissance 

satellites, officials defined requirements for Special Reconnaissance System 118P. 

These called initially for a range of 3,500 miles at altitudes above 100,000 feet. 

Bell won funding in September 1955, as a follow-on to its recently completed MX2276 activity, and proposed a two-stage vehicle with a Mach 15 glider. In March 

1956 the company won a new study contract for what now was called Brass Bell. It 

took shape as a fairly standard advanced concept of the mid-1950s, with a liquidfueled expendable first stage boosting a piloted craft that showed sharply swept delta 

wings. The lower stage was conventional in design, burning Atlas propellants with 

uprated Atlas engines, but the glider retained the company’s preference for fluorine. 

Officials at Bell were well aware of its perils, but John Sloop at NACA-Lewis was 

successfully testing a fluorine rocket engine with 20,000 pounds of thrust, and this 

gave hope.5

The Brass Bell study contract went into force at a moment when prospects for 

boost-glide were taking a sharp step upward. In February 1956 General Thomas 

Power, head of the Air Research and Development Command (ARDC), stated 

that the Air Force should stop merely considering such radical concepts and begin 

developing them. High on his list was a weapon called Robo, Rocket Bomber, for 

which several firms were already conducting in-house work as a prelude to funded 

study contracts. Robo sought to advance beyond Brass Bell, for it was to circle the 

globe and hence required near-orbital speed. In June ARDC Headquarters set forth 

System Requirement 126 that defined the scope of the studies. Convair, Douglas, 

and North American won the initial awards, with Martin, Bell, and Lockheed later 

participating as well.

The initial Bomi concept, dating back to 1951, took form as an in-house effort. 

It called for a two-stage rocket, with both stages being piloted and fitted with delta 

wings. The lower stage was mostly of aluminum, with titanium leading edges and 

nose; the upper stage was entirely of titanium and used radiative cooling. With an 

initial range of 3,500 miles, it was to come over the target above 100,000 feet and 

at speeds greater than Mach 4. Operational concepts called for bases in England or 

Spain, targets in the western Soviet Union, and a landing site in northern Africa.2

During the spring of 1952, Bell officials sought funds for further study from 

Wright Air Development Center (WADC). A year passed, and WADC responded 

with a firm no. The range was too short. Thermal protection and onboard cooling 

raised unanswered questions. Values assumed for L/D appeared highly optimistic, 

and no information was available on stability, control, or aerodynamic flutter at the 

proposed speeds. Bell responded by offering to consider higher speeds and greater 

The Bomi concept. (Art by Dennis Jenkins)
136

Facing the Heat Barrier: A History of Hypersonics

137

Widening Prospects for Re-entry

Director of Development Planning, and from Brigadier General Homer Boushey, 

Deputy Director of Research and Development. NACA’s John Crowley, Associate 

Director for Research, gave strong approval to the proposed test vehicle, viewing it 

as a logical step beyond the X-15. On 25 November, having secured support from 

his superiors, Boushey issued Development Directive 94, allocating $3 million to 

proceed with more detailed studies following a selection of contractors.10

The new concept represented another step in the sequence that included Eugen 

Sänger’s Silbervogel, his suborbital skipping vehicle, and among live rocket craft, the 

X-15. It was widely viewed as a tribute to Sänger, who was still living. It took the 

name Dyna-Soar, which drew on “dynamic soaring,” Sänger’s name for his skipping 

technique, and which also stood for “dynamic ascent and soaring flight,” or boostglide. Boeing and Martin emerged as the finalists in June 1958, with their roles 

being defined in November 1959. Boeing was to take responsibility for the winged 

spacecraft. Martin, described as the associate contractor, was to provide the Titan 

missile that would serve as the launch vehicle.11

The program now demanded definition of flight modes, configuration, structure, and materials. The name of Sänger was on everyone’s lips, but his skipping 

flight path had already proven to be uncompetitive. He and his colleague Bredt 

had treated its dynamics, but they had not discussed the heating. That task fell to 

NACA’s Allen and Eggers, along with their colleague Stanford Neice.

Top and side views of Dyna-Soar. (U.S. Air Force)

The X-15 by then was well along in design, but it clearly was inadequate for the 

performance requirements of Brass Bell and Robo. This raised the prospect of a new 

and even more advanced experimental airplane. At ARDC Headquarters, Major 

George Colchagoff took the initiative in pursuing studies of such a craft, which took 

the name HYWARDS: Hypersonic Weapons Research and Development Supporting System. In November 1956 the ARDC issued System Requirement 131, thereby 

placing this new X-plane on the agenda as well.6

The initial HYWARDS concept called for a flight speed of Mach 12. However, 

in December Bell Aircraft raised the speed of Brass Bell to Mach 18. This increased 

the boost-glide range to 6,300 miles, but it opened a large gap between the performance of the two craft, inviting questions as to the applicability of HYWARDS 

results. In January a group at NACA-Langley, headed by John Becker, weighed in 

with a report stating that Mach 18, or 18,000 feet per second, was appropriate for 

HYWARDS. The reason was that “at this speed boost gliders approached their peak 

heating environment. The rapidly increasing flight altitudes at speeds above Mach 

18 caused a reduction in the heating rates.”7

With the prospect now strong that Brass Bell and HYWARDS would have the 

same flight speed, there was clear reason not to pursue them as separate projects but 

to consolidate them into a single program. A decision at Air Force Headquarters, 

made in March 1957, accomplished this and recognized their complementary characters. They still had different goals, with HYWARDS conducting flight research 

and Brass Bell being the operational reconnaissance system, but HYWARDS now 

was to stand as a true testbed.8

Robo still was a separate project, but events during 1957 brought it into the 

fold as well. In June an ad hoc review group, which included members from ARDC 

and WADC, looked at Robo concepts from contractors. Robert Graham, a NACA 

attendee, noted that most proposals called for “a boost-glide vehicle which would fly 

at Mach 20-25 at an altitude above 150,000 feet.” This was well beyond the state of 

the art, but the panel concluded that with several years of research, an experimental 

craft could enter flight test in 1965, an operational hypersonic glider in 1968, and 

Robo in 1974.9

On 10 October—less than a week after the Soviets launched their first Sputnik—

ARDC endorsed this three-part plan by issuing a lengthy set of reports, “Abbreviated Systems Development Plan, System 464L—Hypersonic Strategic Weapon 

System.” It looked ahead to a research vehicle capable of 18,000 feet per second and 

350,000 feet, to be followed by Brass Bell with the same speed and 170,000 feet, 

and finally Robo, rated at 25,000 feet per second and 300,000 feet but capable of 

orbital flight.

The ARDC’s Lieutenant Colonel Carleton Strathy, a division chief and a strong 

advocate of program consolidation, took the proposed plan to Air Force Headquarters. He won endorsement from Brigadier General Don Zimmerman, Deputy 
138

Facing the Heat Barrier: A History of Hypersonics

139

Widening Prospects for Re-entry

per second. His colleague Peter Korycinski worked with Becker to develop heating 

analyses of flat-top and flat-bottom candidates, with Roger Anderson and others 

within Langley’s Structures Division providing estimates for the weight of thermal 

protection.

A simple pair of curves, plotted on graph paper, showed that under specified 

assumptions the flat-bottom weight at that velocity was 21,400 pounds and was 

increasing at a modest rate at higher speeds. The flat-top weight was 27,600 pounds 

and was rising steeply. Becker wrote that the flat-bottom craft placed its fuselage 

“in the relatively cool shielded region on the top or lee side of the wing—i.e., the 

wing was used in effect as a partial heat shield for the fuselage.... This ‘flat-bottomed’ design had the least possible critical heating area...and this translated into 

least circulating coolant, least area of radiative heat shields, and least total thermal 

protection in flight.”15

These approaches—flat-top at Ames, flat-bottom at Langley—brought a debate 

between these centers that continued through 1957. At Ames, the continuing strong 

interest in high L/D reflected an ongoing emphasis on excellent supersonic aerodynamics for military aircraft, which needed high L/D as a matter of course. To ease 

the heating problem, Ames held for a time to a proposed speed of 11,000 feet per 

second, slower than the Langley concept but lighter in weight and more attainable 

in technology while still offering a considerable leap beyond the X-15. Officials 

at NACA diplomatically described the Ames and Langley HYWARDS concepts 

respectively as “high L/D” and “low heating,” but while the debate continued, there 

remained no standard approach to the design of wings for a hypersonic glider.16

There was a general expectation that such a craft would require active cooling. 

Bell Aircraft, which had been studying Bomi, Brass Bell, and lately Robo, had the 

most experience in the conceptual design of such arrangements. Its Brass Bell of 

1957, designed to enter its glide at 18,000 feet per second and 170,000 feet in altitude, featured an actively cooled insulated hot structure. The primary or load-bearing structure was of aluminum and relied on cooling in a closed-loop arrangement 

that used water-glycol as the coolant. Wing leading edges had their own closed-loop 

cooling system that relied on a mix of sodium and potassium metals. Liquid hydrogen, pumped initially to 1,000 pounds per square inch, flowed first through a heat 

exchanger and cooled the heated water-glycol, then proceeded to a second heat 

exchanger to cool the hot sodium-potassium. In an alternate design concept, this 

gas cooled the wing leading edges directly, with no intermediate liquid-metal coolant loop. The warmed hydrogen ran a turbine within an onboard auxiliary power 

unit and then was exhausted overboard. The leading edges reached a maximum 

temperature of 1,400oF, for which Inconel X was a suitable material.17

During August of that year Becker and Korycinski launched a new series of studies that further examined the heating and thermal protection of their flat-bottom 

In 1954, following their classic analysis of ballistic re-entry, Eggers and Allen 

turned their attention to comparison of this mode with boost-glide and skipping 

entries. They assumed the use of active cooling and found that boost-glide held the 

advantage:

The glide vehicle developing lift-drag ratios in the neighborhood of 4 is far 

superior to the ballistic vehicle in ability to convert velocity into range. It 

has the disadvantage of having far more heat convected to it; however, it 

has the compensating advantage that this heat can in the main be radiated 

back to the atmosphere. Consequently, the mass of coolant material may 

be kept relatively low.

A skip vehicle offered greater range than the alternatives, in line with Sänger’s 

advocacy of this flight mode. But it encountered more severe heating, along with 

high aerodynamic loads that necessitated a structurally strong and therefore heavy 

vehicle. Extra weight meant extra coolant, with the authors noting that “ultimately the coolant is being added to cool coolant. This situation must obviously be 

avoided.” They concluded that “the skip vehicle is thought to be the least promising 

of the three types of hypervelocity vehicle considered here.”12

Following this comparative assessment of flight modes, Eggers worked with his 

colleague Clarence Syvertson to address the issue of optimum configuration. This 

issue had been addressed for the X-15; it was a mid-wing airplane that generally 

resembled the high-performance fighters of its era. In treating Dyna-Soar, following 

the Robo review of mid-1957, NACA’s Robert Graham wrote that “high-wing, midwing and low-wing configurations were proposed. All had a highly swept wing, and 

a small angle cone as the fuselage or body.” This meant that while there was agreement on designing the fuselage, there was no standard way to design the wing.13

Eggers and Syvertson proceeded by treating the design problem entirely as an 

exercise in aerodynamics. They concluded that the highest values of L/D were attainable by using a high-wing concept with the fuselage mounted below as a slender 

half-cone and the wing forming a flat top. Large fins at the wing tips, canted sharply 

downward, directed the airflow under the wings downward and increased the lift. 

Working with a hypersonic wind tunnel at NACA-Ames, they measured a maximum 

L/D of 6.65 at Mach 5, in good agreement with a calculated value of 6.85.14

This configuration had attractive features, not the least of which was that the 

base of its half-cone could readily accommodate a rocket engine. Still, it was not 

long before other specialists began to argue that it was upside down. Instead of 

having a flat top with the fuselage below, it was to be flipped to place the wing below 

the fuselage, giving it a flat bottom. This assertion came to the forefront during 

Becker’s HYWARDS study, which identified its preferred velocity as 18,000 feet 
140

Facing the Heat Barrier: A History of Hypersonics

141

Widening Prospects for Re-entry

boost-glider Brass Bell and for the manned rocket-powered bomber Robo. But the 

rationale for both projects became increasingly questionable during the early 1960s. 

The hypersonic Brass Bell gave way to a new concept, the Manned Orbiting Laboratory (MOL), which was to fly in orbit as a small space station while astronauts 

took reconnaissance photos. Robo fell out of the picture completely, for the success 

of the Minuteman ICBM, which used solid propellant, established such missiles as 

the nation’s prime strategic force. Some people pursued new concepts that continued to hold out hope for Dyna-Soar applications, with satellite interception standing in the forefront. The Air Force addressed this with studies of its Saint project, 

but Dyna-Soar proved unsuitable for such a mission.20

Dyna-Soar was a potentially superb technology demonstrator, but Defense Secretary Robert McNamara took the view that it had to serve a military role in its 

own right or lead to a follow-on program with clear military application. The cost 

of Dyna-Soar was approaching a billion dollars, and in October 1963 he declared 

that he could not justify spending such a sum if it was a dead-end program with no 

ultimate purpose. He canceled it on 10 December, noting that it was not to serve as 

a cargo rocket, could not carry substantial payloads, and could not stay in orbit for 

Full-scale model of Dyna-Soar, on display at an Air Force exhibition in 1962. The scalloped pattern on the base was intended to suggest Sänger’s skipping entry. (Boeing Company archives)

glider. They found that for a glider of global range, flying with angle of attack of 45 

degrees, an entry trajectory near the upper limit of permissible altitudes gave peak 

uncooled skin temperatures of 2,000°F. This appeared achievable with improved 

metallic or ceramic hot structures. Accordingly, no coolant at all was required!18

This conclusion, published early 

in 1959, influenced the configuration of subsequent boost-glide vehicles—Dyna-Soar, the space shuttle—much as the Eggers-Allen paper 

of 1953 had defined the blunt-body 

shape for ballistic entry. Preliminary and unpublished results were 

in hand more than a year prior to 

publication, and when the prospect 

emerged of eliminating active cooling, the concepts that could do this 

were swept into prominence. They 

were of the flat-bottom type, with 

Dyna-Soar being the first to proceed 

into mainstream development.

This uncooled configuration 

proved robust enough to accommodate substantial increases in flight 

speed and performance. In April 

1959 Herbert York, the Defense 

Director of Research and Engineering, stated that Dyna-Soar was to fly 

at 15,000 miles per hour. This was 

well above the planned speed of Brass 

Bell but still below orbital velocity. 

During subsequent years the booster 

changed from Martin’s Titan I to the more capable Titan II and then to the powerful 

Titan III-C, which could easily boost it to orbit. A new plan, approved in December 

1961, dropped suborbital missions and called for “the early attainment of orbital 

flight.” Subsequent planning anticipated that Dyna-Soar would reach orbit with 

the Titan III upper stage, execute several circuits of the Earth, and then come down 

from orbit by using this stage as a retrorocket.19

After that, though, advancing technical capabilities ran up against increasingly 

stringent operational requirements. The Dyna-Soar concept had grown out of 

HYWARDS, being intended initially to serve as a testbed for the reconnaissance 

Artist’s rendering showing Dyna-Soar boosted 

by a Titan III launch vehicle. (Boeing Company 

archives)
142

Facing the Heat Barrier: A History of Hypersonics

143

Widening Prospects for Re-entry

long durations. He approved MOL as a new program, thereby giving the Air Force 

continuing reason to hope that it would place astronauts in orbit, but stated that 

Dyna-Soar would serve only “a very narrow objective.”21

At that moment the program called for production of 10 flight vehicles, and 

Boeing had completed some 42 percent of the necessary tasks. McNamara’s decision therefore was controversial, particularly because the program still had highlevel supporters. These included Eugene Zuckert, Air Force Secretary; Alexander 

Flax, Assistant Secretary for Research and Development; and Brockway McMillan, 

Zuckert’s Under Secretary and Flax’s predecessor as Assistant Secretary. Still, McNamara gave more attention to Harold Brown, the Defense Director of Research and 

Engineering, who made the specific proposal that McNamara accepted: to cancel 

Dyna-Soar and proceed instead with MOL.22

Dyna-Soar never flew. The program had expended $410 million when canceled, 

but the schedule still called for another $373 million, and the vehicle was still some 

two and a half years away from its first flight. Even so, its technology remained available for further development, contributing to the widening prospects for reentry 

that marked the era.23

The Technology of Dyna-Soar

Its thermal environment during re-entry was less severe than that of an ICBM 

nose cone, allowing designers to avoid not only active structural cooling but ablative thermal protection as well. This meant that it could be reusable; it did not have 

to change out its thermal protection after every flight. Even so, its environment 

imposed temperatures and heat loads that pervaded the choice of engineering solutions throughout the vehicle.

Dyna-Soar used radiatively-cooled hot structure, with the primary or load-bearing structure being of Rene 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, imposing differential expansion on separate beams, 

caused these members to rotate at the pins. This accommodated the gradients without imposing thermal stress.

Rene 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 psi at room temperature, fell off only slightly at 1,200oF 

and retained useful values at 1,800oF. It could be processed as sheet, strip, wire, 

tubes, and forgings. Used as the primary structure of Dyna-Soar, it supported a 

design specification that indeed called for reusability. The craft was to withstand at 

least four re-entries under the most severe conditions permitted.

As an alloy, Rene 41 had a standard composition of 19 percent chromium, 11 

percent cobalt, 10 percent molybdenum, 3 percent titanium, and 1.5 percent aluArtist’s rendering showing Dyna-Soar in orbit. (Boeing Company archives)
144

Facing the Heat Barrier: A History of Hypersonics

145

Widening Prospects for Re-entry

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.27

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 where temperatures during re-entry did not 

exceed 2,700oF, while TZM was used for temperatures between 2,700 and 3,000oF. 

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 3,000oF radiated 43 percent more heat than one at 2,700oF.28

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.

Development testing featured use of an oxyacetylene torch, operated with excess 

oxygen, which heated small samples of coated refractory sheet to temperatures as 

high as 3,000oF, measured by optical pyrometer. Test durations ran as long as four 

hours, with a published review noting that failures of specimens “were easily detected 

by visual observation as soon as they occurred.” This work showed that although 

TZM had better oxidation resistance than D-36, both coated alloys could resist 

oxidation for more than two hours at 3,000oF. This exceeded design requirements. 

Similar tests applied stress to hot samples by hanging weights from them, thereby 

demonstrating their ability to withstand stress of 3,100 psi, again at 3,000oF.29

Other tests showed that complete panels could withstand aerodynamic flutter. 

This issue was important; a report of the Aerospace Vehicles Panel of the Air Force 

Scientific Advisory Board (SAB)—a panel on panels, as it were—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 by 4-foot Unitary 

facility at Langley that covered a range of Mach 1.6 to 2.8 and the 11 by 11-foot 

Unitary installation at Ames for Mach 1.2 to 1.4. Heaters warmed test samples to 

840oF as investigators started with steel panels and progressed to versions fabricated 

from Rene 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; actual flutter could 

easily have brought failure through fatigue, with parts of a specimen being blown 

through the tunnel at supersonic speed. The work therefore proceeded by starting 

minum, 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 permitted full aging without cracking, with the fabricated assemblies showing no significant tendency to develop cracks.24

As a structural material, the relatively mature state of Rene 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, reradiating the heat while withstanding temperatures as 

high as 3,000oF. 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, and 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 due to 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. For a time it appeared as a highly promising candidate for all the other 

panels.25

Wing design also promoted its use, for the craft mounted a delta wing with a 

leading-edge sweep of 73 degrees. Though built for hypersonic re-entry from orbit, 

it resembled the supersonic delta wings of contemporary aircraft such as the B-58 

bomber. However, this wing was designed using the Eggers-Allen blunt-body principle, with the leading edge being curved or 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.26

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, again at the most extreme temperatures. D-36 indeed was chosen to cover 
146

Facing the Heat Barrier: A History of Hypersonics

147

Widening Prospects for Re-entry

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 a vacuum and at high temperatures, to form a thin protective layer of silicon carbide. Near the stagnation point, the temperatures during re-entry would exceed that 

level. This brought the concept of a nose cap with siliconized graphite as the primary material, with an insulating 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 through the full time of re-entry.

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 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 re-entry. A ramjet, fueled with JP-4 and 

running with air from a wind tunnel, produced an exhaust with velocity up to 

4,500 feet per second and temperature up to 3,500°F. It also generated acoustic 

levels above 170 decibels, 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 using 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 

attached to the vehicle with a forged ring and clamp that also were of coated TZM. 

The cap as a whole relied on radiative cooling. It was designed to be reusable; like 

the primary structure, it was to withstand four re-entries under the most severe 

conditions permitted.34

at modest dynamic pressures, 400 and 500 pounds per square foot, and advancing 

over 18 months to levels that exceeded the design requirement of close to 1,400 

pounds per square foot. The Boeing report concluded that the success of this test 

program, which ran through mid-1962, “indicates that an adequate panel flutter 

capability has been achieved.”30

Between the outer panels and the inner primary structure, a corrugated skin of 

Rene 41 served as a substructure. On the upper wing surface and upper fuselage, 

where temperatures were no higher than 2,000°F, the thermal-protection panels 

were also of Rene 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 2,000oF, 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

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 shortduration overtemperatures of 2,900°F. However, the nose cap faced temperatures 

of 3,680°F, along with a peak heat flux of 143 BTU per square foot-second. This 

cap had a radius of curvature of 7.5 inches, making it far less blunt than the 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, and so 

the program conducted two independent development efforts that used separate 

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. However, this metal could not be welded, 

while no known coating could protect it against oxidation. Attention then turned 

to nonmetallic materials, including ceramics.
148

Facing the Heat Barrier: A History of Hypersonics

149

Widening Prospects for Re-entry

ers and cooled these loops. Brass Bell had called for its warmed hydrogen to flow 

through a turbine, operating the onboard Auxiliary Power Unit. Dyna-Soar used 

an arrangement that differed only slightly: a catalytic bed to combine the stream 

of warm hydrogen with oxygen that again came from an onboard supply. This 

produced gas that drove the turbine of the Dyna-Soar APU, which provided both 

hydraulic and electric power.

A cooled hydraulic system also was necessary to move the control surfaces as on 

a conventional aircraft. 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 water-glycol loop that removed its heat by cooling with hydrogen. 

Major hydraulic system components, including pumps, were mounted within an 

actively cooled compartment. Control-surface actuators, along with their 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 hydraulic 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 cobalt-base 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. However, Dyna-Soar needed ball 

bearings qualified for 1,600°F and obtained them as spheres of Rene 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 Dyna-Soar 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. Its emittance had the low value of 0.4 at 2,500°F, which meant that waveguides beneath 

these windows faced thermal damage even though they were made of columbium 

alloy. A mix of oxides of cobalt, aluminum, and nickel gave a suitable coating when 

fired at 3,000°F, raising the emittance to approximately 0.8.40

The pilot needed his own windows. The three main ones, facing forward, were 

the largest yet planned for a manned spacecraft. They had double panes of fused 

silica, with infrared-reflecting coatings 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, the frames were fitted with springs of Rene 41. The windows 

also needed thermal protection, and so they were covered with a shield of D-36. 

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 feet per second and an acoustic level of 154 decibels. 

Both installations were capable of long-duration testing, reproducing conditions 

during re-entries that could last for 30 minutes.35

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 both 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 was to 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, insulation. Active cooling also played 

a role, reducing temperatures within the pilot’s compartment and two equipment 

bays. These used a “water wall,” which mounted absorbent material between sheetmetal 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 hot spots.

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 cooling loops resembling those proposed for Brass Bell of 1957. In 

Dyna-Soar, liquid hydrogen from an onboard supply flowed through heat exchang-
150

Facing the Heat Barrier: A History of Hypersonics

151

Widening Prospects for Re-entry

Heat Shields for Mercury and Corona

In November 1957, a month after the first Sputnik reached orbit, the Soviets 

again startled the world by placing a much larger satellite into space, which held 

the dog Laika as a passenger. This clearly presaged the flight of cosmonauts, and the 

question then was how the United States would respond. No plans were ready at the 

moment, but whatever America did, it would have to be done quickly.

HYWARDS, the nascent Dyna-Soar, was proceeding smartly. In addition, at 

North American Aviation the company’s chief engineer, Harrison Storms, was in 

Washington, DC, with a concept designated X-15B. Fitted with thermal protection 

for return from orbit, it was to fly into space atop a cluster of three liquid-fueled 

boosters for an advanced Navaho, each with thrust of 415,000 pounds.44 However, 

neither HYWARDS nor the X-15B could be ready soon. Into this breach stepped 

Maxime Faget of NACA-Langley, who had already shown a talent for conceptual 

design during the 1954 feasibility study that led to the original X-15.

In 1958 he was a branch chief within Langley’s Pilotless Aircraft Research Division. Working on speculation, amid full awareness that the Army or Air Force might 

win the man-in-space assignment, he initiated a series of paper calculations and 

wind-tunnel tests of what he described as a “simple nonlifting satellite vehicle which 

follows a ballistic path in reentering the atmosphere.” He noted that an “attractive 

feature of such a vehicle is that the research and production experiences of the ballistic-missile programs are applicable to its design and construction,” and “since it 

follows a ballistic path, there is a minimum requirement for autopilot, guidance, or 

control equipment.”45

In seeking a suitable shape, Faget started with the heat shield. Invoking the AllenEggers principle, he at first considered a flat face. However, it proved to trap heat by 

interfering with the rapid airflow that could carry this heat away. This meant that 

there was an optimum bluntness, as measured by radius of curvature.

Calculating thermal loads and heat-transfer rates using theories of Lees and of 

Fay and Riddell, and supplementing these estimates with experimental data from 

his colleague William Stoney, he considered a series of shapes. The least blunt was 

a cone with a rounded tip that faced the airflow. It had the highest heat input and 

the highest peak heating rate. A sphere gave better results in both areas, while the 

best estimates came with a gently rounded surface that faced the flow. It had only 

two-thirds the total heat input of the rounded cone—and less than one-third the 

peak heating rate. It also was the bluntest shape of those considered, and it was 

selected.46

With a candidate heat-shield shape in hand, he turned his attention to the complete manned capsule. An initial concept had the shape of a squat dome that was 

recessed slightly from the edge of the shield, like a circular Bundt cake that does 

not quite extend to the rim of its plate. The lip of this heat shield was supposed to 

The cockpit was supposed to be jettisoned following re-entry, 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. The test pilot Neil Armstrong flew approaches and 

landings with a modified Douglas F5D fighter and showed that it was possible to 

land Dyna-Soar safely with side vision only.41

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 Waspaloy nickel steel and mounted wire bristles of Rene 

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 Rene 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 slideout.42

Through such means, the Dyna-Soar program took long strides toward establishing hot structures as a technology suitable for operational use during re-entry 

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 Rene 41 superalloy 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

TEMPERATURE LIMITS BEFORE AND AFTER DYNA-SOAR (in °F)

Element 1958 1963

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

Meanwhile, while Dyna-Soar was going forward within the Air Force, NASA 

had its own approaches to putting man in space.
152

Facing the Heat Barrier: A History of Hypersonics

153

Widening Prospects for Re-entry

The second Corona launch, in April 1959, flew successfully and became the 

world’s first craft to return safely from orbit. It was supposed to come down near 

Hawaii, and a ground controller transmitted a command to have the capsule begin 

re-entry at a particular time. However, he forgot to press a certain button. The 

director of the recovery effort, Lieutenant Colonel Charles “Moose” Mathison, then 

learned that it would actually come down near the Norwegian island of Spitzbergen.

Mathison telephoned a friend in Norway’s air force, Major General Tufte Johnsen, and told him to watch for a small spacecraft that was likely to be descending 

by parachute. Johnsen then phoned a mining company executive on the island and 

had him send out ski patrols. A three-man patrol soon returned with news: They 

had seen the orange parachute as the capsule drifted downward near the village of 

Barentsburg. That was not good because its residents were expatriate Russians. General Nathan Twining, Chairman of the Joint Chiefs, summarized the craft’s fate in a 

memo: “From concentric circular tracks found in the snow at the suspected impact 

point and leading to one of the Soviet mining concessions on the island, we strongly 

suspect that the Soviets are in possession of the capsule.”51

Meanwhile, NASA’s Maxime Faget was making decisions concerning thermal 

protection for his own program, which now had the name Project Mercury. He 

was well aware of ablation but preferred heat sink. It was heavier, but he doubted 

that industrial contractors could fabricate an ablative heat shield that had adequate 

reliability.52

The suitability of ablation could not be tested by flying a subscale heat shield 

atop a high-speed rocket. Nothing less would do than to conduct a full-scale test 

using an Atlas ICBM as a booster. This missile was still in development, but in 

December 1958 the Air Force Ballistic Missile Division agreed to provide one Atlas 

C within six months, along with eight Atlas Ds over the next several years. This 

made it possible to test an ablative heat shield for Mercury as early as September 

1959.53

The contractor for this shield was General Electric. The ablative material, phenolic-fiberglass, lacked the excellent insulating properties of Teflon or phenolicnylon. Still, it had flown successfully as a ballistic-missile nose cone. The project 

engineer Aleck Bond adds that “there was more knowledge and experience with 

fiberglass-phenolic than with other materials. A great deal of ground-test information was available.... There was considerable background and experience in the 

fabrication, curing, and machining of assemblies made of Fiberglass.” These could 

be laid up and cured in an autoclave.54

The flight test was called Big Joe, and it showed conservatism. The shield was 

heavy, with a density of 108 pounds per cubic foot, but designers added a large 

safety factor by specifying that it was to be twice as thick as calculations showed to 

be necessary. The flight was to be suborbital, with range of 1,800 miles but was to 

produce separated flow over the afterbody to reduce its heating. When tested in a 

wind tunnel, however, it proved to be unstable at subsonic speeds.

Faget’s group eliminated the open lip and exchanged the domed afterbody for a 

tall cone with a rounded tip that was to re-enter with its base end forward. It proved 

to be stable in this attitude, but tests in the 11-inch Langley hypersonic wind tunnel 

showed that it transferred too much heat to the afterbody. Moreover, its forward tip 

did not give enough room for its parachutes. This brought a return to the domed 

afterbody, which now was somewhat longer and had a cylinder on top to stow the 

chutes. Further work evolved the domed shape into a funnel, a conic frustum that 

retained the cylinder. This configuration provided a basis for design of the Mercury 

and later of the Gemini capsules, both of which were built by the firm of McDonnell Aircraft.47

Choice of thermal protection quickly emerged as a critical issue. Fortunately, the 

thermal environment of a re-entering satellite proved to be markedly less demanding 

than that of an ICBM. The two vehicles were similar in speed and kinetic energy, 

but an ICBM was to slam back into the atmosphere at a steep angle, decelerating 

rapidly due to drag and encountering heating that was brief but very severe. Reentry from orbit was far easier, taking place over a number of minutes. Indeed, 

experimental work 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 prospective new role of satellite reentry.48

Heat of ablation, in BTU per pound, had been a standard figure of merit. For 

satellite entry, however, with little energy being carried away by ablation, it could 

be irrelevant. Phenolic glass, a fine ICBM material with a measured heat of 9,600 

BTU per pound, was unusable for a satellite because it had an unacceptably high 

thermal conductivity. This meant that the prolonged thermal soak of re-entry could 

have time enough to fry a spacecraft. 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.49

Such results showed that it was not necessary to reopen the problem of thermal 

protection for satellite entry. With appropriate caveats, the experience and research 

techniques of the ICBM problem could carry over to this new realm. This background made it possible for the Central Intelligence Agency to build operational 

orbital re-entry vehicles at a time when nose cones for Atlas were still in flight test.

This happened beginning in 1958, when Richard Bissell, a senior manager 

within the CIA, launched a highly classified reconnaissance program called Corona. 

General Electric, which was building nose cones for Atlas, won a contract to build 

the film-return capsule. The company selected ablation as the thermal-protection 

method, with phenolic nylon as the ablative material.50
154

Facing the Heat Barrier: A History of Hypersonics

155

Widening Prospects for Re-entry

would subject its astronauts to excessive g-forces. Too shallow a re-entry meant that 

it would show insufficient loss of speed within the upper atmosphere and would fly 

back into space, to make a final entry and then land at an unplanned location. For 

a simple ballistic trajectory, this “corridor” was as little as seven miles wide, from 

top to bottom.59

At the outset, these issues raised two problems that were to be addressed in flight 

test. The heat shield had to be qualified, in tests that resembled those of the X-17 

but took place at much higher velocity. In addition, it was necessary to show that a 

re-entering spacecraft could maneuver with some precision. It was vital to broaden 

the corridor, and the only way to do this was to use lift. This meant demonstrating successful maneuvers that had to be planned in advance, using data from tests 

in ground facilities at near-orbital speeds, when such facilities were most prone to 

error.

Apollo’s Command Module, which was to execute the re-entry, lacked wings. 

Still, spacecraft of this general type could show lift-to-drag ratios of 0.1 or 0.2 by 

flying at a nonzero angle of attack, thereby tilting the heat shield and turning it into 

a lifting surface. Such values were far below those achievable with wings, but they 

brought useful flexibility during re-entry by permitting maneuver, thereby achieving a more accurate splashdown.

As early as 1958, Faget and his colleagues had noted three methods for trimming 

a capsule to a nonzero angle. Continuous thrust from a reaction-control system 

could do this, tilting the craft from its equilibrium attitude. A drag flap could do 

it as well by producing a modest amount of additional air resistance on one side of 

the vehicle. The simplest method required no onboard mechanism that might fail 

in flight and that expended no reaction-control propellant. It called for nothing 

more than a nonsymmetrical distribution of weight within the spacecraft, creating 

an offset in the location of the center of gravity. During re-entry, this offset would 

trim the craft to a tilted attitude, again automatically, due to the extra weight on one 

side. An astronaut could steer his capsule by using attitude control to roll it about its 

long axis, thereby controlling the orientation of the lift vector.60

This center-of-gravity offset went into the Gemini capsules that followed those 

of Project Mercury. The first manned Gemini flight carried the astronauts Virgil 

“Gus” Grissom and John Young on a three-orbit mission in March 1965. Following 

re-entry, they splashed down 60 miles short of the carrier USS Intrepid, which was 

on the aim point. This raised questions as to the adequacy of the preflight hypersonic wind-tunnel tests that had provided estimates of the spacecraft L/D used in 

mission planning.

The pertinent data had come from only two facilities. The Langley 11-inch 

tunnel had given points near Mach 7, while an industrial hotshot installation covered Mach 15 to 22, which was close to orbital speed. The latter facility lacked 

simulate a re-entry from orbit that was relatively steep and therefore demanding, 

producing higher temperatures on the face of the shield and on the afterbody.55

Liftoff came after 3 a.m., a time chosen to coincide with dawn in the landing 

area so as to give ample daylight for search and recovery. “The night sky lit up and 

the beach trembled with the roar of the Rocketdyne engines,” notes NASA’s history 

of Project Mercury. Two of those engines were to fall away during ascent, but they 

remained as part of the Atlas, increasing its weight and reducing its peak velocity by 

some 3,000 feet per second. What was more, the capsule failed to separate. It had 

an onboard attitude-control system that was to use spurts of compressed nitrogen 

gas to turn it around, to enter the atmosphere blunt end first. But this system used 

up all its nitrogen trying fruitlessly to swing the big Atlas that remained attached. 

Separation finally occurred at an altitude of 345,000 feet, while people waited to 

learn what would happen.56

The capsule performed better than planned. Even without effective attitude control, its shape and distribution of weights gave it enough inherent stability to turn 

itself around entirely through atmospheric drag. Its reduced speed at re-entry meant 

that its heat load was only 42 percent of the planned value of 7,100 BTU per square 

foot. But a particularly steep flight-path angle gave a peak heating rate of 77 percent 

of the intended value, thereby subjecting the heat shield to a usefully severe test. 

The capsule came down safely in the Atlantic, some 500 miles short of the planned 

impact area, but the destroyer USS Strong was not far away and picked it up a few 

hours later.

Subsequent examination showed that the heating had been uniform over the face 

of the heat shield. This shield had been built as an ablating laminate with a thickness 

of 1.075 inches, supported by a structural laminate half as thick. However, charred 

regions extended only to a depth of 0.20 inch, with further discoloration reaching 

to 0.35 inch. Weight loss due to ablation came to only six pounds, in line with 

experimental findings that had shown that little ablation indeed would occur.57

The heat shield not only showed fine thermal performance, it also sustained no 

damage on striking the water. This validated the manufacturing techniques used in 

its construction. The overall results from this flight test were sufficiently satisfactory 

to justify the choice of ablation for Mercury. This made it possible to drop heat sink 

from consideration and to go over completely to ablation, not only for Mercury but 

for Gemini, which followed.58 

Gemini and Apollo

An Apollo spacecraft, returning from the Moon, had twice the kinetic energy of 

a flight in low orbit and an aerodynamic environment that was nearly three times 

as severe. Its trajectory also had to thread a needle in its accuracy. Too steep a return 
156

Facing the Heat Barrier: A History of Hypersonics

157

Widening Prospects for Re-entry

The 1962 estimates of Apollo L/D proved to be off by 20 percent, with the 

trim angle being in error by 3 degrees.64 As with the Gemini data, these results 

showed anew that one could not obtain reliable data by working with a limited 

range of facilities. But when investigators broadened their reach to use more facilities, and sought accuracy through such methods as elimination of model-support 

errors, they indeed obtained results that matched flight test. This happened twice, 

with both Gemini and Apollo, with researchers finally getting the accurate estimates 

they needed.

These studies dealt with aerodynamic data at hypervelocity. In a separate series, 

other flights sought data on the re-entry environment that could narrow the range 

of acceptable theories of hypervelocity heating. Two such launches constituted Project Fire, which flew spacecraft that were approximately two feet across and had the 

general shape of Apollo’s Command Module. Three layers of beryllium served as 

calorimeters, with measured temperature rises corresponding to total absorbed heat. 

Three layers of phenolic-asbestos alternated with those layers to provide thermal 

protection. Windows of fused quartz, which is both heat-resistant and transparent 

over a broad range of optical wavelengths, permitted radiometers to directly observe 

the heat flux due to radiation, at selected locations. These included the nose, where 

heating was most intense.

The Fire spacecraft rode atop Atlas boosters, with flights taking place in April 

1964 and May 1965. Following cutoff of the Atlas, an Antares solid-fuel booster, 

modified from the standard third stage of the Scout booster, gave the craft an additional 17,000 feet per second and propelled it into the atmosphere at an angle of 

nearly 15 degrees, considerably steeper than the range of angles that were acceptable 

for an Apollo re-entry. This increased the rate of heating and enhanced the contribution from radiation. Each beryllium calorimeter gave useful data until its outer 

surface began to melt, which took only 2.5 seconds as the heating approached its 

maximum. When decelerations due to drag reached specified levels, an onboard 

controller ejected the remnants of each calorimeter in turn, along with its underlying layer of phenolic-asbestos. Because these layers served as insulation, each ejection exposed a cool beryllium surface as well as a clean set of quartz windows.

Fire 1 entered the atmosphere at 38,000 feet per second, markedly faster than 

the 35,000 feet per second of operational Apollo missions. Existing theories gave a 

range in estimates of total peak heating rate from 790 to 1,200 BTU per square footsecond. The returned data fell neatly in the middle of this range. Fire 2 did much 

the same, re-entering at 37,250 feet per second and giving a measured peak heating 

rate of just over 1,000 BTU per square foot-second. Radiative heating indeed was 

significant, amounting to some 40 percent of this total. But the measured values, 

obtained by radiometer, were at or below the minimum estimates obtained using 

existing theories.65

instruments of adequate precision and had produced data points that showed a large 

scatter. Researchers had averaged and curve-fit the measurements, but it was clear 

that this work had introduced inaccuracies.61

During that year flight data became available from the Grissom-Young mission 

and from three others, yielding direct measurements of flight angle of attack and 

L/D. To resolve the discrepancies, investigators at the Air Force’s Arnold Engineering Development Center undertook further studies using two additional facilities. 

Tunnel F, a hotshot, had a 100-inch-diameter test section and reached Mach 20, 

heating nitrogen with an electric arc and achieving run times of 0.05 to 0.1 seconds. 

Tunnel L was a low-density, continuous-flow installation that also used arc-heated 

nitrogen. The Langley 11-inch data was viewed as valid and was retained in the 

reanalysis.

This work gave an opportunity to benchmark data from continuous-flow and 

hotshot tunnels against flight data, at very high Mach numbers. Size did not matter, 

for the big Tunnel F accommodated a model at one-fifteenth scale that incorporated much detail, whereas Tunnel L used models at scales of 1/120 and 1/180, the 

latter being nearly small enough to fit on a tie tack. Even so, the flight data points 

gave a good fit to curves derived using both tunnels. Billy Griffith, supervising the 

tests, concluded: “Generally, excellent agreement exists” between data from these 

sources.

The preflight data had brought estimated values of L/D that were too high by 60 

percent. This led to a specification for the re-entry trim angle that proved to be off 

by 4.7 degrees, which produced the miss at splashdown. Julius Lukasiewicz, longtime head of the Von Karman Gas Dynamics Facility at AEDC, later added that if 

AEDC data had been available prior to the Grissom-Young flight, “the impact point 

would have been predicted to within ± 10 miles.”62

The same need for good data reappeared during Apollo. The first of its orbital 

missions took place during 1966, flying atop the Saturn I-B. The initial launch, 

designated AS-201, flew suborbitally and covered 5,000 miles. A failure in the reaction controls produced uncontrolled lift during entry, but the craft splashed down 

38 miles from its recovery ship. AS-202, six months later, was also suborbital. It 

executed a proper lifting entry—and undershot its designated aim point by 205 

miles. This showed that its L/D had also been mispredicted.63

Estimates of the Apollo L/D had relied largely on experimental data taken during 

1962 at Cornell Aeronautical Laboratory and Mach 15.8, and at AEDC and Mach 

18.7. Again these measurements lacked accuracy, and once more Billy Griffith of 

AEDC stepped forward to direct a comprehensive set of new measurements. In 

addition to Tunnels F and L, used previously, the new work used Tunnels A, B, and 

C, which with the other facilities covered a range from Mach 3 to 20. To account 

for effects due to model supports in the wind tunnels, investigators also used a gun 

range that fired small models as free-flight projectiles, at Mach 6.0 to 8.5.
158

Facing the Heat Barrier: A History of Hypersonics

159

Widening Prospects for Re-entry

Saturn I-B that went no further than earth orbit, and which included AS-201 and 

-202, the next flight lifted off in November 1967. It used a Saturn V to simulate a 

true lunar return.

No larger rocket had ever flown. This one was immense, standing 36 stories 

tall. The anchorman Walter Cronkite gave commentary from a nearby CBS News 

studio, and as this behemoth thundered upward atop a dazzling pillar of yellowwhite flame, Cronkite shouted, “Oh, my God, our building is shaking! Part of the 

roof has come in here!” The roar was as loud as a major volcanic eruption. People 

saw the ascent in Jacksonville, 150 miles away.70

Heat-shield qualification stood as a major goal. The upper stages operated in 

sequence, thrusting the spacecraft to an apogee of 11,242 miles. It spent several 

hours coasting, oriented with the heat shield in the cold soak of shadow to achieve 

the largest possible thermal gradient around the shield. Re-ignition of the main 

engine pushed the spacecraft into re-entry at 35,220 feet per second relative to the 

atmosphere of the rotating Earth. Flying with an effective L/D of 0.365, it came 

down 10 miles from the aim point and only six miles from the recovery ship, close 

enough for news photos that showed a capsule in the water with one of its chutes 

still billowing.

The heat shield now was ready for the Moon, for it had survived a peak heating 

rate of 425 BTU per square foot-second and a total heat load of 37,522 BTU per 

pound. Operational lunar flights imposed loads and heating rates that were markedly less demanding. In the words of Pavlosky and St. Leger, “the thermal protection 

subsystem was overdesigned.”71

A 1968 review took something of an offhand view of what once had been seen 

as an extraordinarily difficult problem. This report stated that thermal performance 

of ablative material “is one of the lesser criteria in developing a TPS.” Significant 

changes had been made to enhance access for inspection, relief of thermal stress, 

manufacturability, performance near windows and other penetrations, and control 

of the center of gravity to achieve design values of L/D, “but never to obtain better 

thermal performance of the basic ablator.”72

Thus, on the eve of the first lunar landing, specialists in hypersonics could look 

at a technology of re-entry whose prospects had widened significantly. A suite of 

materials now existed that were suitable for re-entry from orbit, having high emissivity to keep the temperature down, along with low thermal conductivity to prevent overheating during the prolonged heat soak. Experience had shown how careful research in ground facilities could produce reliable results and could permit 

maneuvering entry with accuracy in aim. This had been proven to be feasible for 

missions as demanding as lunar return.

Dyna-Soar had not flown, but it introduced metallic hot structures that brought 

the prospect of reusability. It also introduced wings for high L/D and particular 

Earlier work had also shown that radiative heating was no source of concern. 

The new work also validated the estimates of total heating that had been used in 

designing the Apollo heat shield. A separate flight test, in August 1964, placed a 

small vehicle—the R-4—atop a five-stage version of the Scout. As with the X-17, 

this fifth stage ignited relatively late in the flight, accelerating the test vehicle to its 

peak speed when it was deep in the upper atmosphere. This speed, 28,000 feet per 

second, was considerably below that of an Apollo entry. But the increased air density 

subjected this craft to a particularly high heating rate.66

This was a materials-testing flight. The firm of Avco had been developing ablators of lower and lower weight and had come up with its 5026-39 series. They used 

epoxy-novolac as the resin, with phenolic microballoons added to the silica-fiber 

filler of an earlier series. Used with a structural honeycomb made of phenolic reinforced with fiberglass, it cut the density to 35 pounds per cubic foot and, with subsequent improvements, to as little as 31 pounds per cubic foot. This was less than 

three-tenths the density of the ancestral phenolic-fiberglass of Mercury—which 

merely orbited the Earth and did not fly back from the Moon.67

The new material had the designation Avcoat 5026-39G. The new flight sought 

to qualify it under its most severe design conditions, corresponding to re-entry at 

the bottom of the corridor with deceleration of 20 g. The peak aerodynamic load 

occurred at Mach 16.4 and 102,000 feet. Observed ablation rates proved to be 

much higher than expected. In fact, the ablative heat shield eroded away completely! 

This caused serious concern, for if that were to happen during a manned mission, 

the spacecraft would burn up in the atmosphere and would kill its astronauts.68

The relatively high air pressure had subjected the heat shield to dynamic pressures three times higher than those of an Apollo re-entry. Those intense dynamic 

pressures corresponded to a hypersonic wind that had blown away the ablative char 

soon after it had formed. This char was important; it protected the underlying virgin 

ablator, and when it was severely thinned or removed, the erosion rate on the test 

heat shield increased markedly.

Much the same happened in October 1965, when another subscale heat shield 

underwent flight test atop another multistage solid rocket, the Pacemaker, that 

accelerated its test vehicle to Mach 10.6 at 67,500 feet. These results showed that 

failure to duplicate the true re-entry environment in flight test could introduce 

unwarranted concern, causing what analysts James Pavlosky and Leslie St. Leger 

described as “unnecessary anxiety and work.”69

An additional Project Fire flight could indeed have qualified the heat shield 

under fully realistic re-entry conditions, but NASA officials had gained confidence 

through their ability to understand the quasi-failure of the R-4. Rather than conduct further ad hoc heat-shield flight tests, they chose to merge its qualification with 

unmanned flights of complete Apollo spacecraft. Following three shots aboard the 
160

Facing the Heat Barrier: A History of Hypersonics

161

Widening Prospects for Re-entry

1 Quotes: DTIC ADA-303832, p. 78.

2 Hallion, Hypersonic, p. 189; Spaceflight, July-August 1980, pp. 270-272. 

3 Hallion, Hypersonic, pp. 189-91; DTIC ADA-303832, pp. 82-87.

4 Hallion, Hypersonic, p. 192; Spaceflight, July-August 1980, p. 270; Bell Aircraft D143-981-009.

5 DTIC ADA-303832, pp. 93-100; Hallion, Hypersonic, pp. 193-95; Bell Aircraft D143-945-

055; Sloop, Astronautics & Aeronautics, October 1972, pp. 52-57.

6 Hallion, Hypersonic, pp. 195-97; DTIC ADA-303832, pp. 106-08.

7 Hallion, Hypersonic, pp. 195, 388-90 (quote, p. 390); DTIC ADA-303832, pp. 122-23.

8 Hallion, Hypersonic, p. 198; DTIC ADA-303832, pp. 126-28.

9 Hallion, Hypersonic, pp. 199-201; DTIC ADA-303832, pp. 128-30. Quote: Memo, Graham 

to Associate Director, 3 July 1957.

10 Hallion, Hypersonic, pp. 201-204; Houchin, Air Power History, Fall 1999, pp. 15-16; DTIC 

ADA-303832, pp. 134-40.

11 Jenkins, Space Shuttle, pp. 22-25; Sunday and London, “X-20,” pp. 259, 262.

12 NACA TN 4046 (quotes, pp. 1, 29).

13 Quote: Memo, Graham to Associate Director, 3 July 1957.

14 NACA RM A55L05.

15 NASA SP-4305, pp. 367-69; Hallion, Hypersonic, pp. 388-91 (quote, p. 390).

16 NASA SP-4305, pp. 369-73; Hallion, Hypersonic, pp. 392-400.

17 Bell Aircraft D143-945-055.

18 Hallion, Hypersonic, pp. 406-07.

19 Sunday and London, “X-20,” pp. 260-61, 264-69 (quote, p. 266).

20 NASA SP-4221, p. 203; Jenkins, Space Shuttle, p. 26.

21 Geiger, “Termination”; Sunday and London, “X-20,” pp. 274-75 (includes quote). Cancellation 

documents: McNamara news briefing, 10 December 1963; telex, AFCVC-1918/63, HQ USAF 

to All Commands, 10 December 1963.

22 Sunday and London, “X-20,” p. 275; Hallion, Hypersonic, pp. 294-310. Technical issues: DTIC 

AD-449685, p. 61.

23 Hallion, Hypersonic, p. II-xvi.

24 DTIC AD-346912, pp. III-3-1-2 to -5, III-3-1-18 to -23, III-4-2-2 to -8. Rene 41: AD449685, pp. 28, 41.

25 DTIC AD-609169; AD-449685, pp. 3-5, 20; AD-346912, pp. III-3-1-7, III-4-1-2 to -3.

26 DTIC AD-346912, p. III-3-1-8; Hallion, Hypersonic, p. II-xv; NACA RM A55E26.

27 DTIC AD-609169; AD-449685, pp. 3-5, 20; AD-346912, pp. III-3-1-7, III-4-1-2 to -3.

28 DTIC AD-346912, p. III-3-1-8; AD-449685, p. iii.

29 DTIC AD-441740; AD-449685, pp. 5-7, 20 (quote, p. 7).

30 “Scientific Advisory Board Memo Report of the Aerospace Vehicles Panel on Dyna Soar Panel 

Flutter,” 20 April 1962; DTIC AD-346912, pp. III-3-6-2 to -15 (quotes, pp. -8, -11).

31 DTIC AD-449685, pp. 49-50; AD-609169, pp. 13, 29; AD-346912, pp. III-3-1-6 to -7, IIIfreedom during maneuver. Indeed, by 1970 there was only one major frontier in 

re-entry: the development of a lightweight heat shield that was simpler than the hot 

structure of Dyna-Soar and was reusable. This topic was held over for the following 

decade, amid the development of the space shuttle.
162

Facing the Heat Barrier: A History of Hypersonics

163

Widening Prospects for Re-entry

66 NASA TN D-7564, TM X-1182; Langley Working Paper LWP-54.

67 AIAA Paper 69-98; NASA TN D-7564, TM X-490, TM X-1407.

68 NASA TN D-7564, TM X-1182; Langley Working Paper LWP-54.

69 AIAA Paper 68-1142; NASA TN D-4713, TN D-7564 (quote, p. 16).

70 Time, 17 November 1967, 84-85 (includes quote).

71 NASA TN D-7564 (quote, p. 19).

72 AIAA Paper 68-1142 (quotes, p. 11).

3-6-2, III-3-6-13 to -14.

32 DTIC AD-346912, pp. III-3-1-8, III-3-4-4; Aviation Week, 21 September 1959, p. 53.

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.

36 Ibid., p. III-3-4-4; Hallion, Hypersonic, p. 357.

37 DTIC AD-346912, pp. III-4-6-5 to -7.

38 Ibid., pp. III-3-1-10 to -12; Hallion, Hypersonic, 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.

41 DTIC AD-346912, pp. III-3-1-10 to -11; Jenkins, Space Shuttle, p. 28.

42 DTIC AD-346912, pp. III-3-1-14 to -15; Hallion, Hypersonic, pp. 347-49, 360-61.

43 Hallion, Hypersonic, pp. 344-46; Sunday and London, “X-20,” p. 278. X-15: Schleicher, Journal of the Royal Aeronautical Society, October 1963, pp. 618-36.

44 Wolfe, Right Stuff, p. 56.

45 Faget: NASA SP-4201, p. 86. Quotes: NACA RM L58E07a, p. 1.

46 NACA RM L58E07a, Table I.

47 Aviation Week, 21 September 1959, pp. 52-59; NASA SP-4201, pp. 94-95.

48 Riddell and Teare, “Differences,” p. 184.

49 Steg, ARS Journal, September 1960, pp. 815-22.

50 Day et al., Eye, pp. 112-14; Greer, “Corona,” pp. 5-6; DTIC AD-328815.

51 Time, 27 April 1959, pp. 16, 65; Richelson: Espionage, p. 184; Secret Eyes, pp. 34-35. Twining 

quote: Day et al., Eye, p. 53.

52 NASA SP-4201, pp. 64-65, 95, 127-28.

53 NASA SP-4201, pp. 123, 125-26, 207.

54 Bond, “Big Joe” (quote, pp. 8-9).

55 Bond, “Big Joe”; memo with attachment, Grimwood to Everline, 3 July 1963; NASA TM X490.

56 Ibid; pp. 16-17; NASA SP-4201, pp. 203-205 (quote, p. 203).

57 NASA TM X-490.

58 Bond, “Big Joe”; Erb and Stephens, “Project Mercury.”

59 NASA TN D-7564; AIAA Paper 68-1142.

60 Lukasiewicz, Experimental, pp. 2-3; NASA RM L58E07a.

61 Lukasiewicz, Experimental, pp. 3-4.

62 Ibid., pp. 4-6 (quote, p. 6); AIAA Paper 67-166 (Griffith quote, p. 10).

63 NASA SP-4206, p. 340; Ley, Rockets, p. 495; Lukasiewicz, Experimental, p. 6.

64 AIAA Paper 68-1142; Lukasiewicz, Experimental, pp. 6-11; Journal of Spacecraft and Rockets, 

July 1968, pp. 843-48.

65 NASA TN D-2996, TN D-7564, TM X-1053, TM X-1120, TM X-1222, TM X-1305.

During the mid-1960s, two advanced flight projects sought to lay technical 

groundwork for an eventual reusable space shuttle. ASSET, which flew first, progressed beyond Dyna-Soar by operating as a flight vehicle that used a hot structure, 

placing particular emphasis on studies of aerodynamic flutter. PRIME, which followed, had a wingless and teardrop-shaped configuration known as a lifting body. 

Its flight tests exercised this craft in maneuvering entries. Separate flights, using 

piloted lifting bodies, were conducted for landings and to give insight into their 

handling qualities.

From the perspective of ASSET and PRIME then, one would have readily concluded that the eventual shuttle would be built as a hot structure and would have 

the aerodynamic configuration of a lifting body. Indeed, initial shuttle design studies, late in the 1960s, followed these choices. However, they were not adopted in 

the final design.

The advent of a highly innovative type of thermal protection, Lockheed’s reusable “tiles,” completely changed the game in both the design and the thermal areas. 

Now, instead of building the shuttle with the complexities of a hot structure, it 

could be assembled as an aluminum airplane of conventional type, protected by 

the tiles. Lifting bodies also fell by the wayside, with the shuttle having wings. The 

Air Force insisted that these be delta wings that would allow the shuttle to fly long 

distances to the side of a trajectory. While NASA at first preferred simple straight 

wings, in time it agreed.

The shuttle relied on carbon-carbon for thermal protection in the hottest areas. 

It was structurally weak, but this caused no problem for more than 100 missions. 

Then in 2003, damage to a wing leading edge led to the loss of Columbia. It was 

the first space disaster to bring the death of astronauts due to failure of a thermal 

protection system.

Preludes: Asset and Lifting Bodies

At the end of the 1950s, ablatives stood out both for the ICBM and for return 

from space. Insulated hot structures, as on Dyna-Soar, promised reusability and 

165

Hypersonics and

the Space Shuttle

6
166

Facing the Heat Barrier: A History of Hypersonics

167

Hypersonics and the Space Shuttle

the Thor-Able that had conducted nose-cone tests at intercontinental range as early 

as 1958.2

The program conducted six flights, which had the following planned values of 

range and of altitude and velocity at release:

Asset Flight Tests

Date Vehicle Booster Velocity, 

feet/second

Altitude, 

feet

Range, 

nautical miles

18 September 1963 ASV-1 Thor 16,000 205,000 987

24 March 1964 ASV-2 Thor-Delta 18,000 195,000 1800

22 July 1964 ASV-3 Thor-Delta 19,500 225,000 1830

27 October 1964 AEV-1 Thor 13,000 168,000 830

8 December 1964 AEV-2 Thor 13,000 187,000 620

23 February 1965 ASV-4 Thor-Delta 19,500 206,000 2300

Source: Hallion, Hypersonic, pp. 505, 510-519.

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. This 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. In the words of historian Richard Hallion, 

“for the first time in aerospace history, a lifting reentry spacecraft had successfully 

returned from space.”3

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 was successfully recovered. It carried a liquid-cooled 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 new nose cap. The standard ASSET type used zirconia dowels, 1.5 inches long by 

0.5 inch in diameter, that 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 re-entry so well that it “could 

have been used again,” and this was true for the craft as a whole. Hallion writes 

lighter weight but were less developed. 

As early as August 1959, the Flight 

Dynamics Laboratory at Wright-Patterson Air Force Base launched an in-house 

study of a small recoverable boost-glide 

vehicle that was to test hot structures 

during re-entry. From the outset there 

was strong interest in problems of aerodynamic flutter. This was reflected in 

the concept name: ASSET or Aerothermodynamic/elastic Structural Systems 

Environmental Tests.

ASSET won approval as a program 

late in January 1961. In April of that 

year the firm of McDonnell Aircraft, 

which was already building Mercury 

capsules, won a contract to develop the 

ASSET flight vehicles. Initial thought 

had called for use of the solid-fuel Scout 

as the booster. Soon, however, it became 

clear that the program could use the 

Thor for greater power. The Air Force 

had deployed these missiles in England. When they came home, during 1963, they 

became available for use as launch vehicles.

ASSET took shape as a flat-bottomed wing-body craft that used the low-wing 

configuration recommended by NASA-Langley. It had a length of 59 inches and 

a span of 55 inches. Its bill of materials closely resembled that of Dyna-Soar, for 

it used TZM to withstand 3,000oF on the forward lower heat shield, graphite for 

similar temperatures on the leading edges, and zirconia rods for the nose cap, which 

was rated at 4,000°F. But ASSET avoided the use of Rene 41, with cobalt and 

columbium alloys being employed instead.1

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 hence 

flew with single-stage Thors as the boosters. But the ASVs were built to study materials and structures in the re-entry environment, while taking data on temperatures, 

pressures, and heat fluxes. Such missions demanded higher speeds. These boostglide craft therefore used the two-stage Thor-Delta launch vehicle, which resembled 

ASSET, showing peak temperatures. 

(U.S. Air Force)
168

Facing the Heat Barrier: A History of Hypersonics

169

Hypersonics and the Space Shuttle

speeds. An early configuration, the M-1 of 1957, featured a blunt-nosed cone with 

a flattened top. It showed some capacity for hypersonic maneuver but could not 

glide subsonically or land on a runway. A new shape, the M-2, appeared as a slender 

half-cone with its flat side up. Its hypersonic L/D of 1.4 was nearly triple that of the 

M-1. Fitted with two large vertical fins for stability, it emerged as a basic configuration that was suitable for further research.6

Dale Reed, an engineer at NASA’s Flight Research Center, developed a strong 

interest in the bathtub-like shape of the M-2. He was a sailplane enthusiast and a 

builder of radio-controlled model aircraft. With support from the local community 

of airplane model builders, he proceeded to craft the M-2 as a piloted glider. Designating it as the M2-F1, he built it of plywood over a tubular steel frame. Completed 

early in 1963, it was 20 feet long and 13 feet across.

It needed a vehicle that could tow it into the air for initial tests. However, it 

produced too much drag for NASA’s usual vans and trucks, and Reed needed a tow 

car with more power. He and his friends bought a stripped-down Pontiac with a big 

engine and a four-barrel carburetor that reached speeds of 110 miles per hour. They 

took it to a funny-car shop in Long Beach for modification. Like any other flightline 

vehicle, it was painted yellow with “National Aeronautics and Space Administration” on its side. Early tow tests showed enough success to allow the project to 

use a C-47, called the Gooney Bird, for true aerial flights. During these tests the 

Gooney Bird towed the M2-F1 above 10,000 feet and then set it loose to glide to 

an Edwards AFB lakebed. Beginning in August 1963, the test pilot Milt Thompson 

did this repeatedly. Reed thus showed that although the basic M-2 shape had been 

crafted for hypersonic re-entry, it could glide to a safe landing.

As he pursued this work, he won support from Paul Bikle, the director of NASA 

Flight Research Center. As early as April 1963, Bikle alerted NASA Headquarters 

that “the lifting-body concept looks even better to us as we get more into it.” The 

success of the M2-F1 sparked interest within the Air Force as well. Some of its officials, along with their NASA counterparts, went on to pursue lifting-body programs 

that called for more than plywood and funny cars. An initial effort went beyond the 

M2-F1 by broadening the range of lifting-body shapes while working to develop 

satisfactory landing qualities.7

NASA contracted with the firm of Northrop to build two such aircraft: the M2-

F2 and HL-10. The M2-F2 amounted to an M2-F1 built to NASA standards; the 

HL-10 drew on an alternate lifting-body design by Eugene Love of NASA-Langley. 

This meant that both Langley and Ames now had a project. The Air Force effort, 

the X-24A, went to the Martin Company. It used a design of Frederick Raymes at 

the Aerospace Corporation that resembled a teardrop fitted with two large fins.

All three flew initially as gliders, with a B-52 rather than a C-47 as the mother 

ship. The lifting bodies mounted small rocket engines for acceleration to supersonic 

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 re-entry 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 ASV and AEV spacecraft.4

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 feet per second, 

Mach 18.4, was the highest in the series and was well above the design speed of 

18,000 feet per second. 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 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.5

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 re-entry 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.

A separate prelude to the space shuttle took form during the 1960s as NASA 

and the Air Force pursued a burgeoning interest in lifting bodies. The initial concept represented one more legacy of the blunt-body principle of H. Julian Allen 

and Alfred Eggers at NACA’s Ames Aeronautical Laboratory. After developing this 

principle, they considered that a re-entering body, while remaining blunt to reduce 

its heat load, might produce lift and thus gain the ability to maneuver at hypersonic 
170

Facing the Heat Barrier: A History of Hypersonics

171

Hypersonics and the Space Shuttle

speeds, thereby enabling tests of stability and handling qualities in transonic flight. 

The HL-10 set records for lifting bodies by making safe approaches and landings at 

Edwards from speeds up to Mach 1.86 and altitudes of 90,000 feet.8

Acceptable handling qualities were not easy to achieve. Under the best of circumstances, a lifting body flew like a brick at low speeds. Lowering the landing gear 

made the problem worse by adding drag, and test pilots delayed this deployment as 

long as possible. In May 1967 the pilot Bruce Peterson, flying the M2-F2, failed to 

get his gear down in time. The aircraft hit the lakebed at more than 250 mph, rolled 

over six times, and then came to rest on its back minus its cockpit canopy, main 

landing gear, and right vertical fin. Peterson, who might have died in the crash, got 

away with a skull fracture, a mangled face, and the loss of an eye. While surgeons 

reconstructed his face and returned him to active duty, the M2-F2 underwent surgery as well. Back at Northrop, engineers installed a center fin and a roll-control 

system that used reaction jets, while redistributing the internal weights. Gerauld 

Gentry, an Air Force test pilot, said that these changes turned “something I really 

did not enjoy flying at all into something that was quite pleasant to fly.”9

The manned lifting-body program sought to turn these hypersonic shapes into 

aircraft that could land on runways, but the Air Force was not about to overlook the 

need for tests of their hypersonic performance during re-entry. The program that 

addressed this issue took shape with the name PRIME, Precision Recovery Including Maneuvering Entry. Martin Marietta, builder of the X-24A, also developed the 

PRIME flight vehicle, the SV-5D that later was referred to as the X-23. Although it 

was only seven feet in length, it faithfully duplicated the shape of the X-24A, even 

including a small bubble-like protrusion near the front that represented the cockpit 

canopy.

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 AFB to locations in the western Pacific near Kwajalein Atoll. The SV5D had higher L/D than Gemini or Apollo, and as with those NASA programs, 

it was to demonstrate precision re-entry. The plans called for crossrange, 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.10

The X-24A was built of aluminum. The SV-5D used this material as well, for 

both the skin and primary structure. It mounted both aerodynamic and reaction 

controls, with the former taking shape as right and left body-mounted flaps set well 

aft. Used together, they controlled pitch; used individually, they produced yaw and 

roll. These flaps were beryllium plates that provided thermal heat sink. The fins 

were of steel honeycomb with surfaces of beryllium sheet.

Lifting bodies. Left to right: the X-24A, the M2-F3 which was modified from the M2-F2, and the 

HL-10. (NASA)

Landing a lifting body. The wingless X-24B required a particularly high angle of attack. (NASA)
172

Facing the Heat Barrier: A History of Hypersonics

173

Hypersonics and the Space Shuttle

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 re-entry. 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: a carbon-phenolic composite. At the nose, its thickness reached 3.5 inches.11

The PRIME program made three flights, which 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 feet per second and flew 4,300 

miles downrange, missing its target by only 900 feet. The vehicle executed pitch 

maneuvers but made no attempt at crossrange. The next two flights indeed achieved 

crossrange, of 500 and 800 nautical miles, and the precision again was impressive. 

Flight 2 missed its aim point by less than two miles. Flight 3 missed by more than 

four 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.12

By demonstrating both crossrange and high accuracy during maneuvering entry, 

PRIME broadened the range of hypersonic aircraft configurations and completed 

a line of development that dated to 1953. In December of that year the test pilot 

Chuck Yeager had nearly been killed when his X-1A fell out of the sky at Mach 

2.44 because it lacked tail surfaces that could produce aerodynamic stability. The 

X-15 was to fly to Mach 6, and Charles McLellan of NACA-Langley showed that it 

could use vertical fins of reasonable size if they were wedge-shaped in cross section. 

Meanwhile, Allen and Eggers were introducing their blunt-body principle. This 

led to missile nose cones with rounded tips, designed both as cones and as blunted 

cylinders that had stabilizing afterbodies in the shape of conic frustums.

For manned flight, Langley’s Maxime Faget introduced the general shape of a 

cone with its base forward, protected by an ablative heat shield. Langley’s John 

Becker entered the realm of winged re-entry configurations with his low-wing 

flat-bottom shapes that showed advantage over the high-wing flat-top concepts of 

NACA-Ames. The advent of the lifting body then raised the prospect of a structurally efficient shape that lacked wings, demanded thermal protection and added 

weight, and yet could land on a runway. Faget’s designs had found application in 

Mercury, Gemini, and Apollo, while Becker’s winged vehicle had provided a basis 

for Dyna-Soar. As NASA looked to the future, both winged designs and lifting 

bodies were in the forefront.13

Martin SV-5D, which became the X-23.

(U.S. Air Force)

Mission of the SV-5D. (U.S. Air Force)

Trajectory of the SV-5D, showing crossrange. (U.S. Air Force)
174

Facing the Heat Barrier: A History of Hypersonics

175

Hypersonics and the Space Shuttle

sintered into a composite. This resulted in a non-ablative 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 re-entry 

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, although the filler did not char. Instead it evaporated or volatilized, producing an outward flow of cool gas that protected the heat shield at high 

heat-transfer 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 pounds per cubic foot, the higher number 

being close to the density of water. Thicknesses of no more than an inch resulted in 

acceptably low back-face temperatures during simulations of re-entry.

This work with silica-fiber ceramics was well under way 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 pounds 

per cubic foot, 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 re-entry 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 the trade journal 

Astronautics & Aeronautics presents the thinking of the day—and in 1965 a Lockheed specialist, Max Hunter, introduced an influential configuration called Star 

Clipper. His design called for LI-1500 as the thermal protection.

Like other shuttle concepts, Star Clipper was to fly repeatedly, but the need for 

an impregnant in LI-1500 

compromised its reusability. In contrast to earlier 

entry vehicle concepts, Star 

Clipper was large, offering 

exposed surfaces that were 

sufficiently blunt to benefit 

from the Allen-Eggers principle. They had lower temperatures and heating rates, 

Reusable Surface Insulation

As PRIME and the lifting bodies broadened the choices of hypersonic shape, 

work at Lockheed made similar contributions in the field of thermal protection. 

Ablatives were unrivaled for once-only use, but during the 1960s the hot structure 

continued to stand out as the preferred approach for reusable craft such as DynaSoar. 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. 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 or even of titanium, 

which had greater heat resistance. On Dyna-Soar, only Rene 41 would do.14

Ablatives avoided such heat leaks, while being sufficiently capable as insulators 

to permit the use of aluminum, as on the SV-5D of PRIME. In principle, a third 

approach combined the best features of hot structure and ablatives. It called for the 

use of temperature-resistant tiles, made perhaps of ceramic, that 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, known as reusable surface insulation (RSI), gave rise in time to 

the thermal protection of the shuttle.

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 they 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. 

This thermal shock resulted from rapid temperature changes during re-entry.15

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 ceramic fiber in lieu of 

the solid blocks. The background to the shuttle’s tiles lay in work with such mats 

that dated to the early 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 re-entry, it had a filamentwound 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 Star Clipper concept. (Art by Dan Gautier)
176

Facing the Heat Barrier: A History of Hypersonics

177

Hypersonics and the Space Shuttle

and degradation of its properties. These problems threatened to grow worse during 

subsequent cycles of re-entry 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 six parts per million. Lockheed did these 

things not only at the laboratory level 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 thermal protection, in lieu of a hot-structure approach.19

General Electric also became actively involved, studying types of RSI made from 

zirconia and from mullite, 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 Dyna-Soar, 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 1000 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 

and higher. 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 pounds per cubic foot, midway between that of LI-900 

and LI-1500.20

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 to a lunar science 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. It specified 

mullite RSI for the undersurface and forward fuselage, a design feature that had 

been held over from the fully-reusable orbiter of the previous year.

Most of the primary structure was aluminum, but that of the nose was titanium, 

with insulation of zirconia lining the nose cap. The wing and fuselage upper surfaces, which had been titanium hot structure, now went over to an elastomeric RSI 

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, that 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 chromium oxide and later with silicon carbide, which further raised the emissivity. The 

glass coating and silica substrate were both silicon dioxide; this assured a match 

of their coefficients of thermal expansion, to prevent the coating from developing 

cracks under the temperature changes of re-entry. 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 re-entry test vehicle, 

along with several candidate ablators. This vehicle carried instruments, and it 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 proof-of-concept test gave further support to the concept of 

high-emittance reradiative tiles of coated silica for thermal protection.16

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 nine pounds per cubic 

foot, one-seventh the density of water.17 There also was much fundamental work 

on materials. Silica exists in three crystalline forms: quartz, cristobalite, tridymite. 

These not only have high coefficients of thermal expansion but also show sudden 

expansion or contraction with temperature due to 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 crystalline state having a very low coefficient of thermal 

expansion and 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.18

The raw silica fiber came from Johns-Manville, which produced it from highpurity 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 fibers broke, giving rise to internal cracking within the RSI 
178

Facing the Heat Barrier: A History of Hypersonics

179

Hypersonics and the Space Shuttle

Douglas Aircraft, for instance, Melvin Root had his two-stage Astro, a fully-reusable 

rocket-powered concept with both stages shaped as lifting bodies. It was to carry a 

payload of 37,150 pounds, and Root expected that a fleet of such craft would fly 

240 times per year. The contemporary Astrorocket of Martin Marietta, in turn, 

looked like two flat-bottom Dyna-Soar craft set belly to belly, foreshadowing fullyreusable space shuttle concepts of several years later.23

These concepts definitely belonged to the Aerospaceplane era. Astro dated to 

1963, whereas Martin’s Astrorocket studies went forward from 1961 to 1965. By 

mid-decade, though, the name “Aerospaceplane” was in bad odor within the Air 

Force. The new concepts were rocket-powered, whereas Aerospaceplanes generally 

had called for scramjets or LACE, and officials referred to these rocket craft as Integrated Launch and Re-entry Vehicles (ILRVs).24

Early contractor studies showed a definite preference for lifting bodies, generally 

with small foldout wings for use when landing. At Lockheed, Hunter’s Star Clipper 

introduced the stage-and-a-half configuration that mounted expendable propellant 

tanks to a reusable core vehicle. The core carried the flight crew and payload along 

with the engines and onboard systems. It had a triangular planform and fitted neatly 

into a large inverted V formed by the tanks. The McDonnell Tip Tank concept was 

broadly similar; it also mounted expendable tanks to a lifting-body core.25

At Convair, people took the view that a single airframe could serve both as a 

core and, when fitted with internal tankage, as a reusable carrier of propellants. 

This led to the Triamese concept, whereby a triplet of such vehicles was to form a 

single ILRV that could rise into the sky. All three were to have thermal protection 

and would re-enter, flying to a runway and deploying their extendable wings. The 

concept was excessively hopeful; the differing requirements of core and tankage 

vehicles proved to militate strongly against a one-size-fits-all approach to airframe 

design. Still, the Triamese approach showed anew that designers were ready to use 

their imaginations.26

NASA became actively involved in the ongoing ILRV studies during 1968. 

George Mueller, the Associate Administrator for Manned Space Flight, took a particular interest and introduced the term “space shuttle” by which such craft came to 

be known. He had an in-house design leader, Maxime Faget of the Manner Spacecraft Center, who was quite strong-willed and had definite ideas of his own as to 

how a shuttle should look. Faget particularly saw lifting bodies as offering no more 

than mixed blessings: “You avoid wing-body interference,” which brings problems 

of aerodynamics. “You have a simple structure. And you avoid the weight of wings.” 

He nevertheless saw difficulties that appeared severe enough to rule out lifting 

bodies for a practical design.

They had low lift and high drag, which meant a dangerously high landing speed. 

As he put it, “I don’t think it’s charming to come in at 250 knots.” Deployable 

wings could not be trusted; they might fail to extend. Lifting bodies also posed 

consisting of a foamed methylphenyl silicone, bonded to the orbiter in panel sizes 

as large as 36 inches. This RSI gave protection to 650oF.21

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 a hundred re-entries, as simulated in ground test. NASA conducted these tests during the 

last five 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 temperatures 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 re-entry. 

LI-1500 continued to show promise. By mid-November it demonstrated the equivalent of 20 cycles to 160 decibels, the acoustic level of a large launch vehicle, and 

2,300oF. 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 2,500oF and survived a thermal overshoot to 3,000°F as 

well as an acoustic overshoot to 174 decibels. Clearly, this was the material NASA 

wanted.22

As insulation, they were astonishing. You 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 using your 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 full-scale production 

facility in Sunnyvale, California. With this, tiles entered the mainstream of thermal 

protection.

Designing the Shuttle

In its overall technologies, the space shuttle demanded advances in a host of 

areas: rocket propulsion, fuel cells and other onboard systems, electronics and computers, and astronaut life support. As an exercise in hypersonics, two issues stood 

out: configuration and thermal protection. The Air Force supported some of the 

early studies, which grew seamlessly out of earlier work on Aerospaceplane. At 
180

Facing the Heat Barrier: A History of Hypersonics

181

Hypersonics and the Space Shuttle

bodies made poor containers, Faget’s wing-body concept was an excellent one. Its 

fuselage could readily be cylindrical, being given over almost entirely to propellant 

tankage.31

The design exercises of 1969 covered thermal protection as well as configuration. 

McDonnell Douglas introduced orbiter designs derived from the HL-10 lifting 

body, examining 13 candidate configurations for the complete two-stage vehicle. 

The orbiter had a titanium primary structure, obtaining thermal protection from a 

hot-structure approach that used external panels of columbium, nickel-chromium, 

and Rene 41. This study also considered the use of tiles made of a “hardened compacted fiber,” which was unrelated to Lockheed’s RSI. However, the company did 

not recommend this. Those tiles were heavier than panels or shingles of refractory 

alloy and less durable.32

North American Rockwell took Faget’s two-stage airplane as its preferred 

approach. It also used a titanium primary structure, with a titanium hot structure 

protecting the top of the orbiter, which faced a relatively mild thermal environment. For the thermally-protected bottom, North American adopted the work of 

Lockheed and specified LI-1500 tiles. The design also called for copious use of 

fiberglass insulation, which gave internal protection to the crew compartment and 

the cryogenic propellant tanks.33

Lockheed turned the Star Clipper core into a reusable second stage that retained 

its shape as a lifting body. Its structure was aluminum, as in a conventional airplane. 

The company was home to LI-1500, and designers considered its use for thermal 

protection. They concluded, though, that this carried high risk. They recommended 

instead a hot-structure approach that used corrugated Rene 41 along with shingles of 

nickel-chromium and columbium. The Lockheed work was independent of that at 

McDonnell Douglas, but engineers at the two firms reached similar conclusions.34

Convair, home of the Triamese concept, came in with new variants. These 

included a triplet launch vehicle with a core vehicle that was noticeably smaller than 

the two propellant carriers that flanked it. Another configuration placed the orbiter 

on the back of a single booster that continued to mount retractable wings. The 

orbiter had a primary structure of aluminum, with titanium for the heat-shield supports on the vehicle underside. Again this craft used a hot structure, with shingles of 

cobalt superalloy on the bottom and of titanium alloy on the top and side surfaces.

Significantly, these concepts were not designs that the companies were prepared 

to send to the shop floor and build immediately. They were paper vehicles that 

would take years to develop and prepare for flight. Yet despite this emphasis on 

the future, and notwithstanding the optimism that often pervades such preliminary design exercises, only North American was willing to recommend RSI as the 

baseline. Even Lockheed, its center of development, gave it no more than a highly 

equivocal recommendation. It lacked maturity, with hot structures standing as the 

approach that held greater promise.35

serious difficulties in development, for they required a fuselage that could do the 

work of a wing. This ruled out straightforward solutions to aerodynamic problems; 

the attempted solutions would ramify throughout the entire design. “They are very 

difficult to develop,” he added, “because when you’re trying to solve one problem, 

you’re creating another problem somewhere else.”27 His colleague Milton Silveira, 

who went on to head the Shuttle Engineering Office at MSC, held a similar view:

“If we had a problem with the aerodynamics on the vehicle, where the body 

was so tightly coupled to the aerodynamics, you couldn’t simply go out 

and change the wing. You had to change the whole damn vehicle, so if you 

make a mistake, being able to correct it was a very difficult thing to do.”28

Faget proposed instead to design his shuttle as a two-stage fully-reusable vehicle, 

with each stage being a winged airplane having low wings and a thermally-protected 

flat bottom. The configuration broadly resembled the X-15, and like that craft, it 

was to re-enter with its nose high and with its underside acting as the heat shield.

Faget wrote that “the vehicle would remain in this flight attitude throughout the 

entire descent to approximately 40,000 feet, where the velocity will have dropped 

to less than 300 feet per second. At this point, the nose gets pushed down, and the 

vehicle dives until it reaches adequate velocity for level flight.” The craft then was 

to approach a runway and land at a moderate 130 knots, half the landing speed of 

a lifting body.29

During 1969 NASA sponsored a round of contractor studies that examined 

anew the range of alternatives. In June the agency issued a directive that ruled out 

the use of expendable boosters such as the Saturn V first stage, which was quite 

costly. Then in August, a new order called for the contractors to consider only twostage fully reusable concepts and to eliminate partially-reusable designs such as Star 

Clipper and Tip Tank. This decision also was based in economics, for a fully-reusable shuttle could offer the lowest cost per flight. But it also delivered a new blow 

to the lifting bodies.30

There was a strong mismatch between lifting-body shapes, which were dictated 

by aerodynamics, and the cylindrical shapes of propellant tanks. Such tanks had to 

be cylindrical, both for ease in manufacturing and to hold internal pressure. This 

pressure was unavoidable; it resulted from boiloff of cryogenic propellants, and it 

served such useful purposes as stiffening the tanks’ structures and delivering propellants to the turbopumps. However, tanks did not fit well within the internal volume 

of a lifting body; in Faget’s words, “the lifting body is a damn poor container.” The 

Lockheed and McDonnell designers had bypassed that problem by mounting their 

tanks externally, with no provision for reuse, but the new requirement of full reusability meant that internal installation now was mandatory. Yet although lifting 
182

Facing the Heat Barrier: A History of Hypersonics

183

Hypersonics and the Space Shuttle

center is located midway between the wing’s leading and trailing edges. At subsonic 

speeds, this center moves forward and is much closer to the leading edge. To keep 

an airplane in proper balance, it requires an aerodynamic force that can compensate 

for this shift.

The Air Force had extensive experience with supersonic fighters and bombers 

that had successfully addressed this problem, maintaining good control and satisfactory handling qualities from Mach 3 to touchdown. Particularly for large aircraft—

the B-58 and XB-70 bombers and the SR-71 spy plane—the preferred solution was 

a delta wing, triangular in shape. Delta wings typically ran along much of the length 

of the fuselage, extending nearly to the tail. Such aircraft dispensed with horizontal 

stabilizers at the tail and relied instead on elevons, control surfaces resembling ailerons that were set at the wing’s trailing edge. Small deflections of these elevons then 

compensated for the shift in the center of lift, maintaining proper trim and balance 

without imposing excessive drag. Draper therefore proposed that both stages of 

Faget’s shuttle have delta wings.37

Faget would have none of this. He wrote that because the only real flying was 

to take place during the landing approach, a wing design “can be selected solely on 

the basis of optimization for subsonic cruise and landing.” The wing best suited 

to this limited purpose would be straight and unswept, like those of fighter planes 

in World War II. A tail would provide directional stability, as on a conventional 

airplane, enabling the shuttle to land in standard fashion. He was well aware of the 

center-of-lift shift but expected to avoid it by avoiding reliance on his wings until 

the craft was well below the speed of sound. He also believed that the delta would 

lose on its design merits. To achieve a suitably low landing speed, he argued that 

the delta would need a large wingspan. A straight wing, narrow in distance between 

its leading and trailing edges, would be light and would offer relatively little area 

demanding thermal protection. A delta of the same span, necessary for a moderate 

landing speed, would have a much larger area than the straight wing. This would 

add a great deal of weight, while substantially increasing the area that needed thermal protection.38

Draper responded with his own view. He believed that Faget’s straight-wing 

design would be barred on grounds of safety from executing its maneuver of stall, 

dive, and recovery. Hence, it would have to glide from supersonic speeds through 

the transonic zone and could not avoid the center-of-lift problem. To deal with it, a 

good engineering solution called for installation of canards, small wings set well forward on the fuselage that would deflect to give the desired control. Canards produce 

lift and would tend to push the main wings farther to the back. They would be well 

aft from the outset, for they were to support an airplane that was empty of fuel but 

that had heavy rocket engines at the tail, placing the craft’s center of gravity far to 

the rear. The wings’ center of lift was to coincide closely with this center of gravity. 

In the wake of the 1969 studies, NASA officials turned away from lifting bodies. 

Lockheed continued to study new versions of the Star Clipper, but the lifting body 

now was merely an alternative. The mainstream lay with Faget’s two-stage fully-reusable approach, showing rocket-powered stages that looked like airplanes. Very soon, 

though, the shape of the wings changed anew, as a result of problems in Congress.

The space shuttle was a political program, funded by federal appropriations, and 

it had to make its way within the environment of Washington. On Capitol Hill, 

an influential viewpoint held that the shuttle was to go forward only if it was a 

national program, capable of meeting the needs of military as well as civilian users. 

NASA’s shuttle studies had addressed the agency’s requirements, but this proved not 

to be the way to proceed. Matters came to a head in the mid-1970 as Congressman 

Joseph Karth, a longtime NASA supporter, declared that the shuttle was merely the 

first step on a very costly road to Mars. He opposed funding for the shuttle in committee, and when he did not prevail, he made a motion from the floor of the House 

to strike the funds from NASA’s budget. Other congressmen assured their colleagues 

that the shuttle had nothing to do with Mars, and Karth’s measure went down to 

defeat—but by the narrowest possible margin: a tie vote of 53 to 53. In the Senate, 

NASA’s support was only slightly greater.36

Such victories were likely to leave NASA undone, and the agency responded 

by seeking support for the shuttle from the Air Force. That service had tried and 

failed to build Dyna-Soar only a few years earlier; now it found NASA offering a 

much larger and more capable space shuttle on a silver platter. However, the Air 

Force was quite happy with its Titan III launch vehicles and made clear that it 

would work with NASA only if the shuttle was redesigned to meet the needs of the 

Pentagon. In particular, NASA was urged to take note of the views of the Air Force 

Flight Dynamics Laboratory (FDL), where specialists had been engaged in a running debate with Faget since early 1969.

The FDL had sponsored ILRV studies in parallel with the shuttle studies of 

NASA and had investigated such concepts as Lockheed’s Star Clipper. One of its 

managers, Charles Cosenza, had directed the ASSET program. Another FDL scientist, Alfred Draper, had taken the lead in questioning Faget’s approach. Faget 

wanted his shuttle stages to come in nose-high and then dive through 15,000 feet 

to pick up flying speed. With the nose so high, these airplanes would be fully stalled, 

and the Air Force disliked both stalls and dives, regarding them as preludes to an 

out-of-control crash. Draper wanted the shuttle to enter its glide while still supersonic, thereby maintaining much better control.

If the shuttle was to glide across a broad Mach range, from supersonic to subsonic, then it would face an important aerodynamic problem: a shift in the wing’s 

center of lift. A wing generates lift across its entire lower surface, but one may regard 

this lift as concentrated at a point, the center of lift. At supersonic speeds, this 
184

Facing the Heat Barrier: A History of Hypersonics

185

Hypersonics and the Space Shuttle

more robust and therefore heavier still. In turn, the extra weight ramified throughout the entire two-stage shuttle vehicle, making it larger and more costly.42

NASA’s key officials included the acting administrator, George Low, and the 

Associate Administrator for Manned Space Flight, Dale Myers. They would willingly have embraced Faget’s shuttle. But on the military side, the Undersecretary 

of the Air Force for Research and Development, Michael Yarymovich, had close 

knowledge of the requirements of the National Reconnaissance Office. He played a 

key role in emphasizing that only a delta would do.

The denouement came at a meeting in Williamsburg, Virginia, in January 1971. 

At nearby Yorktown, in 1781, Britain’s Lord Charles Cornwallis had surrendered 

to General George Washington, thereby ending America’s war of independence. 

One hundred and ninety years later NASA surrendered to the Air Force, agreeing particularly to build a delta-wing shuttle with full military crossrange of 1,100 

miles. In return, though, NASA indeed won the support from the Pentagon that it 

needed. Opposition faded on Capitol Hill, and the shuttle program went forward 

on a much stronger political foundation.43

The design studies of 1969 had counted as Phase A and were preliminary in 

character. In 1970 the agency launched Phase B, conducting studies in greater 

depth, with North American Rockwell and McDonnell Douglas as the contractors. 

Initially they considered both straight-wing and delta designs, but the Williamsburg 

decision meant that during 1971 they were to emphasize the deltas. These remained 

as two-stage fully-reusable configurations, which were openly presented at an AIAA 

meeting in July of that year.

In the primary structure and outer skin of the wings and fuselage, both contractors proposed to use titanium freely. They differed, however, in their approaches to 

thermal protection. McDonnell Douglas continued to favor hot structures. Most of 

the underside of the orbiter was covered with shingles of Hastelloy-X nickel superalloy. The wing leading edges called for a load-bearing structure of columbium, with 

shingles of coated columbium protecting these leading edges as well as other areas 

that were too hot for the Hastelloy. A nose cap of carbon-carbon completed the 

orbiter’s ensemble.44

North American had its own interest in titanium hot structures, specifying them 

as well for the upper wing surfaces and the upper fuselage. Everywhere else possible, 

the design called for applying mullite RSI directly to a skin of aluminum. Such tiles 

were to cover the entire underside of the wings and fuselage, along with much of the 

fuselage forward of the wings. The nose and leading edges, both of the wings, and 

the vertical fin used carbon-carbon. In turn, the fin was designed as a hot structure 

with a skin of Inconel 718 nickel alloy.45

By mid-1971, though, hot structures were in trouble. The new Office of Management and Budget had made clear that it expected to impose stringent limits on 

funding for the shuttle, which brought a demand for new configurations that could 

Draper wrote that the addition of canards “will move the wings aft and tend to close 

the gap between the tail and the wing.” The wing shape that fills this gap is the delta, 

and Draper added that “the swept delta would most likely evolve.”39

Faget had other critics, while Draper had supporters within NASA. Faget’s 

thoughts indeed faced considerable resistance within NASA, particularly among the 

highly skilled test and research pilots at the Flight Research Center. Their spokesman, Milton Thompson, was certainly a man who knew how to fly airplanes, for 

he was an X-15 veteran and had been slated to fly Dyna-Soar as well. But in addition, these aerodynamic issues involved matters of policy, which drove the Air Force 

strongly toward the delta. The reason was that a delta could achieve high crossrange, 

whereas Faget’s straight wing could not.

Crossrange was essential for single-orbit missions, launched from Vandenberg 

AFB on the California coast, which were to fly in polar orbit. The orbit of a spacecraft is essentially fixed with respect to distant stars, but the Earth rotates. In the 

course of a 90-minute shuttle orbit, this rotation carries the Vandenberg site eastward by 1,100 nautical miles. The shuttle therefore needed enough crossrange to 

cover that distance.

The Air Force had operational reasons for wanting once-around missions. A key 

example was rapid-response satellite reconnaissance. In addition, the Air Force was 

well aware that problems following launch could force a shuttle to come down at the 

earliest opportunity, executing a “once-around abort.” NASA’s Leroy Day, a senior 

shuttle manager, emphasized this point: “If you were making a polar-type launch 

out of Vandenberg, and you had Max’s straight-wing vehicle, there was no place you 

could go. You’d be in the water when you came back. You’ve got to go crossrange 

quite a few hundred miles in order to make land.”40

By contrast, NASA had little need for crossrange. It too had to be ready for oncearound abort, but it expected to launch the shuttle from Florida’s Kennedy Space 

Center on trajectories that ran almost due east. Near the end of its single orbit, the 

shuttle was to fly across the United States and could easily land at an emergency 

base. A 1969 baseline program document, “Desirable System Characteristics,” 

stated that the agency needed only 250 to 400 nautical miles of crossrange, which 

Faget’s straight wing could deliver with straightforward modifications.41

Faget’s shuttle had a hypersonic L/D of about 0.5. Draper’s delta-wing design 

was to achieve an L/D of 1.7, and the difference in the associated re-entry trajectories increased the weight penalty for the delta. A delta orbiter in any case needed 

a heavier wing and a larger area of thermal protection, and there was more. The 

straight-wing craft was to have a relatively brief re-entry and a modest heating rate. 

The delta orbiter was to achieve its crossrange by gliding hypersonically, executing a 

hypersonic glide that was to produce more lift and less drag. It also would increase 

both the rate of heating and its duration. Hence, its thermal protection had to be 
186

Facing the Heat Barrier: A History of Hypersonics

187

Hypersonics and the Space Shuttle

Other refractory metals offered alternatives to columbium, but even when proposing to use them, the complexity of a hot structure also militated against their 

selection. As a mechanical installation, it called for large numbers of clips, brackets, 

stand-offs, 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 overtemperatures. They also faced the continuing issue of sealing panel edges 

against ingestion of hot gas during re-entry.51

In this fashion, having taken a long look at hot structures, NASA did an aboutface as it turned toward the RSI that Lockheed’s Max Hunter had recommended as 

early as 1965. The choice of aluminum for the primary structure reflected the excellent insulating properties of RSI, but there was more. Titanium offered a potential advantage because of its temperature resistance; hence, its thermal protection 

might be lighter. However, the apparent weight saving was largely lost due to 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 

thermal conductivity than titanium. Hence, it 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 approximately 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.52

The move toward RSI and aluminum simplified the design and cut the development cost. Substantially larger cost savings came into view as well, as NASA moved 

away from full reusability of its two-stage concepts. The emphasis now was on partial reusability, which the prescient Max Hunter had advocated as far back as 1965 

when he placed the liquid hydrogen of Star Clipper in expendable external tanks. 

The new designs kept propellants within the booster, but they too called for the use 

of external tankage for the orbiter. This led to reduced sizes for both stages and had 

a dramatic impact on the problem of providing thermal protection for the shuttle’s 

booster.

On paper, a shuttle booster amounted to the world’s largest airplane, combining 

the size of a Boeing 747 or C-5A with performance exceeding that of the X-15. It 

was to re-enter at speeds well below orbital velocity, but still it needed thermal protection, and the reduced entry velocities did not simplify the design. North American, for one, specified RSI for its Phase B orbiter, but the company also had to show 

cut the cost of development. Within weeks, the contractors did a major turnabout. 

They went over to primary structures of aluminum. They also 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, switching to RSI once it was ready.46

What brought this dramatic change? The advent of RSI production at Lockheed 

was critical. This drew attention from Faget, who had kept his hand in the field of 

shuttle design, offering a succession of conceptual configurations that had helped to 

guide the work of the contractors. His most important concept, designated MSC040, came out in September 1971 and served as a point of reference. It used aluminum and RSI.47

“My history has always been to take the most conservative approach,” Faget 

declared. Everyone knew how to work with aluminum, for it was the most familiar 

of materials, but titanium was literally a black art. Much of the pertinent shopfloor experience had been gained within the SR-71 program and was classified. Few 

machine shops had the 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 because these metals had been used mostly in 

turbine blades. Lockheed had encountered serious difficulties as its machinists and 

metallurgists wrestled with the use of titanium. With the shuttle facing cost constraints, no one wanted to risk an overrun while machinists struggled with the problems of other new materials.48

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. In addition, coatings posed serious 

problems. Silicides showed promise of reusability and long life, but they were fragile 

and damaged easily. 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.49 “The least little 

scratch in the coating, the shingle would be destroyed during re-entry,” said Faget. 

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.”50
188

Facing the Heat Barrier: A History of Hypersonics

189

Hypersonics and the Space Shuttle

that it understood hot structures. These went into the booster, which protected its 

hot areas with titanium, Inconel 718, carbon-carbon, Rene 41, Haynes 188 steel—

and coated columbium.

The move toward external tankage brought several advantages, the most important of which was a reduction in staging velocity. When designing a two-stage 

rocket, standard methods exist for dividing the work load between the two stages so 

as to achieve the lowest total weight. These methods give an optimum staging velocity. A higher value makes the first stage excessively large and heavy; a lower velocity 

means more size and weight for the orbiter. Ground rules set at NASA’s Marshall 

Space Flight Center, based on such optimization, placed this staging velocity close 

to 10,000 feet per second.

But by offloading propellants into external tankage, the orbiter could shrink 

considerably in size and weight. The tanks did not need thermal protection or heavy 

internal structure; they might be simple aluminum shells stiffened with internal 

pressure. With the orbiter being lighter, and being little affected by a change in staging velocity, a recalculation of the optimum value showed advantage in making the 

tanks larger so that they could carry more propellant. This meant that the orbiter 

was to gain more speed in flight—and the booster would gain less. Hence, the 

booster also could shrink in size. Better yet, the reduction in staging velocity eased 

the problem of booster thermal protection.

Grumman was the first company to pursue this line of thought, as it studied 

alternative concepts alongside the mainstream fully reusable designs of McDonnell 

Douglas and North American. Grumman gave a fully-reusable concept of its own, 

for purposes of comparison, but emphasized a partially-reusable orbiter that put 

the liquid hydrogen in two external tanks. The liquid oxygen, which was dense and 

compact, remained aboard this vehicle, but the low density of the hydrogen meant 

that its tanks could be bulky while remaining light in weight.

The fully-reusable design followed the NASA Marshall Space Flight Center 

ground rules and showed a staging velocity of 9,750 feet per second. The externaltank configuration cut this to 7,000 feet per second. Boeing, which was teamed with 

Grumman, found that this substantially reduced the need for thermal protection as 

such. The booster now needed neither tiles nor exotic metals. Instead, like the X15, it was to use its structure as a heat sink. During re-entry, it would experience a 

sharp but brief pulse of heat, which a conventional aircraft structure could absorb 

without exceeding temperature limits. Hot areas continued to demand a titanium 

hot structure, which was to cover some one-eighth of the booster. However, the rest 

of this vehicle could make considerable use of aluminum.

How could bare aluminum, without protection, serve in a shuttle booster? It was 

common understanding that aluminum airframes lost strength due to aerodynamic 

heating at speeds beyond Mach 2, with titanium being necessary at higher speeds. 

However, this held true for aircraft in cruise, which faced their temperatures conThermal-protection tiles for the space shuttle. (NASA)

Thermal-protection system for the space shuttle. (NASA)
190

Facing the Heat Barrier: A History of Hypersonics

191

Hypersonics and the Space Shuttle

The Loss of Columbia

Thermal protection was delicate. The tiles lacked structural strength and were 

brittle. It was not possible even to bond them directly to the underlying skin, for 

they would fracture and break due to their inability to follow the flexing of the skin 

under its loads. Designers therefore placed an intermediate layer between tiles and 

skin that had some elasticity and could stretch in response to shuttle skin flexing 

without transmitting excessive strain to the tiles. It worked; there never was a serious 

accident due to fracturing of tiles.57

The same was not true of another piece of delicate work: a thermal-protection 

panel made of carbon that had the purpose of protecting one of the wing leading 

edges. It failed during re-entry in virtually the final minutes of a flight of Columbia, 

on 1 February 2003. For want of this panel, that spacecraft broke up over Texas in 

a shower of brilliant fireballs. All aboard were killed.

The background to this accident lay in the fact that for the nose and leading 

edges of the shuttle, silica RSI was not enough. These areas needed thermal protection with greater temperature resistance, and carbon was the obvious candidate. It 

was lighter than aluminum and could be protected against oxidation with a coating. 

It also had a track record, having formed the primary structure of the Dyna-Soar 

nose cap and the leading edge of ASSET. Graphite was the standard form, but in 

contrast to ablative materials, it failed to enter the aerospace mainstream. It was brittle and easily damaged and 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 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: light weight, temperature resistance, and resistance to oxidation when coated. It had the useful property 

of actually gaining strength with temperature, being up to 50 percent stronger at 

3,000oF than at room temperature. It had 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, 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 due to 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.58

tinually. The Boeing booster was to re-enter at Mach 7, matching the top speed of 

the X-15. Even so, its thermal environment resembled a fire that does not burn your 

hand when you whisk it through quickly. Across part of the underside, the vehicle 

was to protect itself by the simple method of metal with more than usual thickness 

to cope with the heat. Even these areas were limited in extent, with the contractors 

noting that “the material gauges [thicknesses] required for strength exceed the minimum heat-sink gauges over the majority of the vehicle.”53

McDonnell Douglas went further. In mid-1971 it introduced its own externaltank orbiter that lowered the staging velocity to 6,200 feet per second. Its winged 

booster was 82 percent aluminum heat sink. Their selected configuration was optimized from a thermal standpoint, bringing the largest savings in the weight of thermal protection.54 Then in March 1972 NASA selected solid-propellant rockets for 

the boosters. The issue of their thermal protection now went away entirely, for these 

big solids used steel casings that were 0.5 inch thick and that provided heat sink 

very effectively.55

Amid the design changes, NASA went over to the Air Force view and embraced 

the delta wing. Faget himself accepted it, making it a feature of his MSC-040 concept. Then the Office of Management and Budget asked whether NASA might 

return to Faget’s straight wing after all, abandoning the delta and thereby saving 

money. Nearly a year after the Williamsburg meeting, Charles Donlan, acting director of the shuttle program office at Headquarters, ruled this out. In a memo to 

George Low, he wrote that high crossrange was “fundamental to the operation of 

the orbiter.” It would enhance its maneuverability, greatly broadening the opportunities to abort a mission and perhaps save the lives of astronauts. High crossrange 

would also provide more frequent opportunities to return to Kennedy Space Center 

in the course of a normal mission.

Delta wings also held advantages that were entirely separate from crossrange. A 

delta orbiter would be stable in flight from hypersonic to subsonic speeds, throughout a wide range of nose-high attitudes. The aerodynamic flow over such an orbiter 

would be smooth and predictable, thereby permitting accurate forecasts of heating 

during re-entry and giving confidence in the design of the shuttle’s thermal protection. In addition, the delta vehicle would experience relatively low temperatures of 

600 to 800oF over its sides and upper surfaces.

By contrast, straight-wing configurations produced complicated hypersonic 

flow fields, with high local temperatures and severe temperature changes on the 

wing, body, and tail. Temperatures on the sides of the fuselage would run from 900 

to 1,300oF, making the design and analysis of thermal protection more complex. 

During transition from supersonic to subsonic speeds, the straight-wing orbiter 

would experience unsteady flow and buffeting, making it harder to fly. This combination of aerodynamic and operational advantages led Donlan to favor the delta for 

reasons that were entirely separate from those of the Air Force.56
192

Facing the Heat Barrier: A History of Hypersonics

193

Hypersonics and the Space Shuttle

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 carboncarbon panel and delivered what proved to be a fatal wound.

Ground controllers became aware of this the following day, during a review of 

high-resolution film taken at the time of launch. The mission continued for two 

weeks, and in the words of the accident report, investigators concluded that “some 

localized heating damage would most likely occur during re-entry, but they could 

not definitively state that structural damage would result.”62

Yet the damage was mortal. Again, in words of the accident report,

Columbia re-entered Earth’s atmosphere with a pre-existing 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°F) 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.63

It was not feasible to go over to a form of thermal protection, for the wing leading edges, that would use a material other than carbon-carbon and that would be 

substantially more robust. Even so, three years of effort succeeded in securing the 

foam and the shuttle returned to flight in July 2006 with foam that stayed put. 

In addition, people took advantage of the fact that most such missions had already 

been intended to dock with the International Space Station. It now became a rule 

that the shuttle could fly only if it were to go there, where it could be inspected 

minutely prior to re-entry and where astronauts could stay, if necessary, until a damaged shuttle was repaired or a new one brought up. In this fashion, rather than the 

thermal protection being shaped to fit the needs of the missions, the missions were 

shaped to fit the requirements of having safe thermal protection.64

Researchers at Vought conducted exploratory studies during the early 1960s, 

investigating resins, fibers, weaves, and coatings. In 1964 they fabricated a DynaSoar nose cap of carbon-carbon, with this exercise permitting comparison of the 

new nose cap with the standard versions that used graphite and zirconia tiles. In 

1966 this firm crafted a heat shield for the Apollo afterbody, which lay leeward of 

the curved ablative front face. A year and a half later the company constructed a 

wind-tunnel model of a Mars probe that was designed to enter the atmosphere of 

that planet.59

These exercises did not approach the full-scale development that Dyna-Soar and 

ASSET had brought to hot structures. They definitely were in the realm of the preliminary. Still, as they went forward along with Lockheed’s work on silica RSI and 

GE’s studies of mullite, the work at Vought made it clear that carbon-carbon was 

likely to take its place amid the new generation of thermal-protection materials.

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 RSI.60

Other 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 five by eight 

feet. It had two leading-edge 

panels of carbon-carbon set side 

by side, a section of wing structure that included its main spars, 

and aluminum skin covered with 

RSI. It had insulated attachments, 

internal insulation, and interface 

seals between the carbon-carbon 

and the RSI. It withstood simulated air loads, launch acoustics, 

and mission temperature-pressure environments, not once but 

many times.61

There was no doubt that left to themselves, the panels of carbon-carbon that 

protected the leading edges would have continued to do so. Unfortunately, they 

were not left to themselves. During the ascent of Columbia, on 16 January 2003, a 

Improvements in strength of carbon-carbon after 1981. 

(AIAA)
194

Facing the Heat Barrier: A History of Hypersonics

195

Hypersonics and the Space Shuttle

torical Series, Shuttle Interviews Box 2. JSC History Collection, University of Houston-Clear 

Lake.

29 Faget, Astronautics & Aeronautics, January 1970, pp. 52-61 (quote, p. 53).

30 NASA SP-4221, pp. 218-219.

31 Author interview, Max Faget, 4 March 1997 (includes quote). Folder 18649, NASA Historical 

Reference Collection, NASA History Division, Washington, D.C. 20546. 

32 McDonnell Douglas Report MDC E0056; Jenkins, Space Shuttle, pp. 84-86.

33 North American Rockwell Report SD 69-573-1; Jenkins, Space Shuttle, pp. 87-89.

34 Lockheed Report LMSC-A959837; Jenkins, Space Shuttle, pp. 90-91.

35 General Dynamics Report GDC-DCB-69-046; Jenkins, Space Shuttle, pp. 92-95.

36 NASA SP-4221, pp. 179-86, 219-21.

37 Ibid., pp. 215-16; Draper et al., Astronautics & Aeronautics, January 1971, pp. 26-35; Hallion, 

Hypersonic, pp. 459, 1032; Jenkins, Space Shuttle, pp. 66-68.

38 Faget, Astronautics & Aeronautics, January 1970, pp.52-61 (quote, p. 57).

39 AIAA Paper 70-1249; Draper et al., Astronautics & Aeronautics, January 1971, pp. 26-35 

(quotes, p. 29).

40 Personal discussions with John Pike, Federation of American Scientists, July 1997; Pace, 

Engineering, pp. 146-49. Quote: John Mauer interview with Leroy Day, 17 October 1983, 

p. 41. Oral Historical Series, Shuttle Interviews Box 1, JSC History Collection, University of 

Houston-Clear Lake. 

41 Day, Task Group Report, Vol. II, pp. 40-42; Faget, Astronautics & Aeronautics, January 1970,

p. 59.

42 NASA SP-4221, p. 217; Faget and Silveira, “Fundamental.”

43 Pace, Engineering, p. 116; NASA SP-4221, pp. 230-31, 233-34.

44 NASA SP-4221, pp. 223-25; AIAA Papers 71-804, 71-805; McDonnell Douglas Report MDC 

E0308.

45 North American Rockwell SV 71-28, SD 71-114-1.

46 Grumman Reports B35-43 RP-33, p. 19; North American Rockwell SV 71-50, p. 9 and SV 

71-59, p. 3; McDonnell Douglas “Interim Report,” p. 36.

47 NASA SP-4221, pp. 341-346; Jenkins, Space Shuttle, pp. 141-50.

48 Heppenheimer, Turbulent, pp. 209-210; author interview, Max Faget, 4 March 1997. Folder 

18649, NASA Historical Reference Collection, NASA History Division, Washington, D.C. 

20546.

49 CASI 81A-44344; Korb et al., “Shuttle,” p. 1189.

50 John Mauer interview, Charles Donlan, 19 October 1983. Oral Historical Series, Shuttle 

Interviews Box 1, JSC History Collection, University of Houston-Clear Lake. Also, author 

interview, Max Faget, 4 March 1997. Folder 18649, NASA Historical Reference Collection, 

NASA History Division, Washington, D.C. 20546.

51 CASI 81A-44344.

52 Pace, Engineering, pp. 179-88.

53 NASA SP-4221, pp. 335-40, 348-49; Grumman Report B35-43 RP-11; Aviation Week, 12 July 

1971, pp. 36-39.

1 “Advanced Technology Program: Technical Development Plan for Aerothermodynamic/Elastic 

Structural System Environmental Tests (ASSET).” Air Force Systems Command, 9 September 

1963.

2 Ibid., pp. 4, 11-13; Hallion, Hypersonic, pp. 451, 464-65.

3 Hallion, Hypersonic, pp. 510-12 (quote, p. 512).

4 Ibid., pp. 512-16 (quote, p. 515); DTIC AD-357523; advertisement, Aviation Week, 24 May 

1965, p. 62.

5 Aviation Week, 2 November 1964, pp. 25-26; DTIC AD-366546 (quote, p. 156); Hallion, 

Hypersonic, pp. 516-19.

6 Hallion, Hypersonic, pp. 524, 529, 535, 864-66.

7 NASA: SP-4220, pp. 33-39; SP-4303, pp. 148-52 (quote, p. 151).

8 NASA: SP-4220, index references; SP-4303, ch. 8; Jenkins, Space Shuttle, p. 34.

9 NASA SP 4220, pp. 106-09, 111-16 (quote, p. 116).

10 Miller, X-Planes, pp. 256-61; Martin Marietta Report ER 14465.

11 Hallion, Hypersonic, pp. 641, 648-49; Martin Marietta Report ER 14465, ch. VI; Aviation 

Week, 16 May 1966, pp. 64-75.

12 Martin Marietta Report ER 14465, p. I-1; Hallion, Hypersonic, pp. 694-702; Aviation Week, 10 

July 1967, pp. 99-101.

13 See index references in the present volume.

14 AIAA Paper 71-443.

15 Schramm et al., “Space Shuttle” (Lockheed Horizons).

16 Ibid.

17 Korb et al., “Shuttle,” p. 1189; Schramm, “HRSI,” p. 1194.

18 NASA TM X-2273, pp. 57-58; Korb et al., “Shuttle,” pp. 1189-90; CASI 72A-10764.

19 Schramm et al., “Space Shuttle,” p. 11 (Lockheed Horizons); Schramm, “HRSI,” p. 1195.

20 CASI 72A-10764; NASA TM X-2273, pp. 39-93.

21 Aviation Week, 17 January 1972, p. 17; North American Rockwell Report SV 72-19.

22 Schramm et al., “Space Shuttle,” pp. 11-14 (Lockheed Horizons); Aviation Week, 27 March 

1972, p. 48.

23 Root and Fuller, Astronautics & Aeronautics, January 1964, pp. 42-51; Hallion, Hypersonic, pp. 

952-54.

24 Jenkins, Space Shuttle, pp. 60-62, 67; Hallion, Hypersonic, pp. 951, 995.

25 Hunter, “Origins”; Schnyer and Voss, “Review,” pp. 17-22, 35-39; Jenkins, Space Shuttle, pp. 

67-70.

26 NASA SP-4221, pp. 89-91; General Dynamics Reports GDC-DCB-67-031, GDC-DCB-68-

017.

27 NASA SP-4221, pp. 92-94; author interview, Max Faget,4 March 1997 (includes quote). 

Folder 18649, NASA Historical Reference Collection, NASA History Division, Washington, 

D.C. 20546. 

28 Milton Silveira interview by Joe Guilmartin and John Mauer, 14 November 1984. Oral His-
196

Facing the Heat Barrier: A History of Hypersonics

54 McDonnell Douglas Report MDC E0376-1.

55 NASA SP-4221, pp. 420-22; Heppenheimer, Development, p. 188.

56 Jenkins, Space Shuttle, p. 147; memo, Donlan to Low, 5 December 1971 (includes quote).

57 Space World, June-July 1979, p. 23; CASI 77A-35304; Korb et al., “Shuttle.”

58 Becker, “Leading”; Korb et al., “Shuttle.”

59 Becker, “Leading”; “Technical” (Vought).

60 Astronautics & Aeronautics, January 1976, pp. 60, 63-64.

61 Aviation Week, 31 March 1975, p. 52; AIAA Paper 78-485.

62 Columbia, p. 38.

63 Ibid., p. 12.

64 Ibid, pp. 225-227.
During the 1960s and 1970s, work in re-entry went from strength to strength. 

The same was certainly not true of scramjets, which reached a peak of activity in 

the Aerospaceplane era and then quickly faded. Partly it was their sheer difficulty, 

along with an appreciation that whatever scramjets might do tomorrow, rockets 

were already doing today. Yet the issues went deeper.

The 1950s saw the advent of antiaircraft missiles. Until then, the history of air 

power had been one of faster speeds and higher altitudes. At a stroke, though, it 

became clear that missiles held the advantage. A hot fighter plane, literally hot from 

aerodynamic heating, now was no longer a world-class dogfighter; instead it was a 

target for a heat-seeking missile.

When antiaircraft no longer could outrace defenders, they ceased to aim at speed 

records. They still needed speed but not beyond a point at which this requirement 

would compromise other fighting qualities. Instead, aircraft were developed with an 

enhanced ability to fly low, where missiles could lose themselves in ground clutter, 

and became stealthy. In 1952, late in the dogfight era, Clarence “Kelly” Johnson 

designed the F-104 as the “missile with a man in it,” the ultimate interceptor. No 

one did this again, not after the real missiles came in.

This was bad news for ramjets. The ramjet had come to the fore around 1950, in 

projects such as Navaho, Bomarc, and the XF-103, because it offered Mach 3 at a 

time when turbojets could barely reach Mach 1. But Mach 3, when actually achieved 

in craft such as the XB-70 and SR-71, proved to be a highly specialized achievement 

that had little to do with practical air power. No one ever sent an SR-71 to conduct 

close air support at subsonic speed, while the XB-70 gave way to its predecessor, the 

B-52, because the latter could fly low whereas the XB-70 could not.

Ramjets also faltered on their merits. The ramjet was one of two new airbreathers that came forth after the war, with the other being the turbojet. Inevitably this 

set up a Darwinian competition in which one was likely to render the other extinct. 

Ramjets from the start were vulnerable, for while they had the advantage of speed, 

they needed an auxiliary boost from a rocket or turbojet. Nor was it small; the 

Navaho booster was fully as large as the winged missile itself.

197

The Fading,

the Comeback

7
198

Facing the Heat Barrier: A History of Hypersonics

199

The Fading, the Comeback 

Scramjets Pass Their Peak

From the outset, scramjets received attention for the propulsion of tactical missiles. In 1959 APL’s Gordon Dugger and Frederick Billig disclosed a concept that 

took the name SCRAM, Supersonic Combustion Ramjet Missile. Boosted by a 

solid-fuel rocket, SCRAM was to cruise at Mach 8.5 and an altitude of 100,000 

feet, with range of more than 400 miles. This cruise speed resulted in a temperature 

of 3,800oF at the nose, which was viewed as the limit attainable with coated materials.1

The APL researchers had a strong interest in fuels other than liquid hydrogen, 

which could not be stored. The standard fuel, a boron-rich blend, used ethyl decaborane. It ignited easily and gave some 25 percent more energy per pound than 

gasoline. Other tests used blends of pentaborane with heavy hydrocarbons, with the 

pentaborane promoting their ignition. The APL group went on to construct and 

test a complete scramjet of 10-inch diameter.2

Paralleling this Navy-sponsored work, the Air Force strengthened its own efforts 

in scramjets. In 1963 Weldon Worth, chief scientist at the Aero Propulsion Laboratory, joined with Antonio Ferri and recommended scramjets as a topic meriting 

attention. Worth proceeded by funding new scramjet initiatives at General Electric 

and Pratt & Whitney. This was significant; these firms were the nation’s leading 

builders of turbojet and turbofan engines.

GE’s complete scramjet was axisymmetric, with a movable centerbody that 

included the nose spike. It was water-cooled and had a diameter of nine inches, 

with this size being suited to the company’s test facility. It burned hydrogen, which 

was quite energetic. Yet the engine failed to deliver net thrust, with this force being 

more than canceled out by drag.3

The Pratt & Whitney effort drew on management and facilities at nearby United 

Aircraft Research Laboratories. Its engine also was axisymmetric and used a long 

cowl that extended well to the rear, forming the outer wall of the nozzle duct. This 

entire cowl moved as a unit, thereby achieving variable geometry for all three major 

components: inlet, combustor, and nozzle. The effort culminated in fabrication of 

a complete water-cooled test unit of 18-inch diameter.4

A separate Aero Propulsion Lab initiative, the Incremental Flight Test Vehicle 

(IFTV), also went forward for a time. It indeed had the status of a flight vehicle, 

with Marquardt holding the prime contract and taking responsibility for the engine. 

Lockheed designed and built the vehicle and conducted wind-tunnel tests at its Rye 

Canyon facility, close to Marquardt’s plant in Van Nuys, California.

The concept called for this craft to ride atop a solid-fuel Castor rocket, which 

was the second stage of the Scout launch vehicle. Castor was to accelerate the IFTV 

to 5,400 feet per second, with this missile then separating and entering free flight. 

Burning hydrogen, its engines were to operate for at least five seconds, adding an 

The problem of compressor stall limited turbojet performance for a time. But 

from 1950 onward, several innovations brought means of dealing with it. They led 

to speedsters such as the F-104 and F-105, operational aircraft that topped Mach 

2, along with the B-58 which also did this. The SR-71, in turn, exceeded Mach 3. 

This meant that there was no further demand for ramjets, which were not selected 

for new aircraft.

The ramjet thus died not only because its market was lost to advanced turbojets, 

but because the advent of missiles made it clear that there no longer was a demand 

for really fast aircraft. This, in turn, was bad news for scramjets. The scramjet was 

an advanced ramjet, likely to enter the aerospace mainstream only while ramjets 

remained there. The decline of the ramjet trade meant that there was no industry 

that might build scramjets, no powerful advocates that might press for them.

The scramjet still held the prospective advantage of being able to fly to orbit as 

a single stage. With Aerospaceplane, the Air Force took a long look as to whether 

this was plausible, and the answer was no, at least not soon. With this the scramjet 

lost both its rationale in the continuing pursuit of high speed and the prospect of an 

alternate mission—ascent to orbit—that might allow it to bypass this difficulty.

In its heyday the scramjet had stood on the threshold of mainstream research and 

development, with significant projects under way at General Electric and United Aircraft Research Laboratories, which was affiliated with Pratt & Whitney. As scramjets faded, though, even General Applied Science Laboratories (GASL), a scramjet 

center that had been founded by Antonio Ferri himself, had to find other activities. 

For a time the only complete scramjet lab in business was at NASA-Langley.

And then—lightning struck. President Ronald Reagan announced the Strategic 

Defense Initiative (SDI), which brought the prospect of a massive new demand for 

access to space. The Air Force already was turning away from the space shuttle, while 

General Lawrence Skantze, head of the Air Force Systems Command, was strongly 

interested in alternatives. He had no background in scramjets, but he embraced the 

concept as his own. The result was the National Aerospace Plane (NASP) effort, 

which aimed at airplane-like flight to orbit.

In time SDI faded as well, while lessons learned by researchers showed that NASP 

offered no easy path to space flight. NASP faded in turn and with it went hopes for 

a new day for hypersonics. Final performance estimates for the prime NASP vehicle, 

the X-30, were not terribly far removed from the early and optimistic estimates that 

had made the project appear feasible. Still, the X-30 design was so sensitive that 

even modest initial errors could drive its size and cost beyond what the Pentagon 

was willing to accept.
200

Facing the Heat Barrier: A History of Hypersonics

201

The Fading, the Comeback 

that NASA was to accept all budgetary responsibility for the X-15 as of 1 January 

1968. This meant that NASA was to pay for further flights—which it refused to do. 

This brought an end to the prospect of using this research airplane for flight testing 

of hypersonic engines.8

The second decision, in August 1967, terminated IFTV. Arthur Thomas, the 

Marquardt program manager, later stated that it had been a major error to embark 

on a flight program before ground test had established attainable performance levels. 

When asked why this systematic approach had not been pursued, Thomas pointed 

to the pressure of a fast-paced schedule that ruled out sequential development. He 

added that Marquardt would have been judged “nonresponsive” if its proposal had 

called for sequential development at the outset. In turn, this tight schedule reflected 

the basic attitude of the Aero Propulsion Lab: to develop a successful scramjet 

quickly and inexpensively, or not to develop one at all.9

Then in September 1968 the Navy elected to close its Ordnance Aerophysics 

Laboratory (OAL). This facility had stood out because it could accommodate test 

engines of realistic size. In turn, its demise brought a premature end to the P & W 

scramjet effort. That project succeeded in testing its engine at OAL at Mach 5, but 

only about 20 runs were conducted before OAL shut down, which was far too few 

for serious development. Nor could this engine readily find a new home; its 18-inch 

diameter had been sized to fit the capabilities of OAL. This project therefore died 

both from withdrawal of Air Force support and from loss of its principal test facility.10

As dusk fell on the Air Force hypersonics program, Antonio Ferri was among the 

first to face up to the consequences. After 1966 he became aware that no major new 

contracts would be coming from the Aero Propulsion Lab, and he decided to leave 

GASL, where he had been president. New York University gave him strong encouragement, offering him the endowed Astor Professorship. He took this appointment 

during the spring of 1967.11

He proceeded to build new research facilities in the Bronx, as New York University bought a parcel of land for his new lab. A landmark was a vacuum sphere for his 

wind tunnel, which his friend Louis Nucci called “the hallmark of hypersonic flow” 

as it sucks high-pressure air from a stored supply. Ferri had left a trail of such spheres 

at his previous appointments: NACA-Langley, Brooklyn Polytechnic, GASL. But 

his new facilities were far less capable than those of GASL, and his opportunities 

were correspondingly reduced. He set up a consulting practice within an existing 

firm, Advanced Technology Labs, and conducted analytical studies. Still, Nucci 

recalls that “Ferri’s love was to do experiments. To have only [Advanced Technology 

Labs] was like having half a body.”

GASL took a significant blow in August 1967, as the Air Force canceled IFTV. 

The company had been giving strong support to the developmental testing of its 

“increment” of velocity of at least 600 feet per second. Following launch over the 

Pacific from Vandenberg AFB, it was to telemeter its data to the ground.

This was the first attempt to develop a scramjet as the centerpiece of a flight program, and much of what could go wrong did go wrong. The vehicle grew in weight 

during development. It also increased its drag and found itself plagued for a time 

with inlets that failed to start. The scramjets themselves gave genuine net thrust but 

still fell short in performance.

The flight vehicle mounted four scramjets. The target thrust was 597 pounds. 

The best value was 477 pounds. However, the engines needed several hundred 

pounds of thrust merely to overcome drag on the vehicle and accelerate, and this 

reduction in performance meant that the vehicle could attain not quite half of the 

desired velocity increase of 600 feet per second.5

Just then, around 1967, the troubles of the IFTV were mirrored by troubles in 

the overall scramjet program. Scramjets had held their promise for a time, with a 

NASA/Air Force Ad Hoc Working Group, in a May 1965 report, calling for an 

expanded program that was to culminate in a piloted hypersonic airplane. The SAB 

had offered its own favorable words, while General Bernard Schriever, head of the 

Air Force Systems Command—the ARDC, its name having changed in 1961—

attempted to secure $50 million in new funding.6

He did not get it, and the most important reason was that conventional ramjets, 

their predecessors, had failed to win a secure role. The ramjet-powered programs of 

the 1950s, including Navaho and Bomarc, now appeared as mere sidelines within 

a grand transformation that took the Air Force in only 15 years from piston-powered B-36 and B-50 bombers to the solid-fuel Minuteman ICBM and the powerful 

Titan III launch vehicle. The Air Force was happy with both and saw no reason for 

scramjet craft as alternatives. This was particularly true because Aerospaceplane had 

come up with nothing compelling.

The Aero Propulsion Laboratory had funded the IFTV and the GE and Pratt 

scramjets, but it had shown that it would support this engine only if it could be 

developed quickly and inexpensively. Neither had proved to be the case. The IFTV 

effort, for one, had escalated in cost from $3.5 million to $12 million, with its 

engine being short on power and its airframe having excessive drag and weight.7

After Schriever’s $50-million program failed to win support, Air Force scramjet 

efforts withered and died. More generally, between 1966 and 1968, three actions 

ended Air Force involvement in broad-based hypersonic research and brought an 

end to a succession of halcyon years. The Vietnam War gave an important reason for 

these actions, for the war placed great pressure on budgets and led to cancellation of 

many programs that lacked urgency.

The first decision ended Air Force support for the X-15. In July 1966 the joint 

NASA-Air Force Aeronautics and Astronautics Coordinating Board determined 
202

Facing the Heat Barrier: A History of Hypersonics

203

The Fading, the Comeback 

This suggestion drew little initial interest, but in November a serious accident 

reopened the question. Though badly damaged, the aircraft, Tail Number 66671, 

proved to be repairable. It returned to flight in June 1964, with modifications that 

indeed gave it the option for engine testing.

The X-15 program thus had this flight-capable testbed in prospect during 1963, 

at a time when engines for test did not even exist on paper. It was not long, though, 

before NASA responded to its opportunity, as Hugh Dryden, the Agency’s Deputy 

Administrator, joined with Robert Seamans, the Associate Administrator, in approving a new program that indeed sought to build a test engine. It took the name of 

Hypersonic Research Engine (HRE).

Three companies conducted initial studies: General Electric, Marquardt, and 

Garrett AiResearch. All eyes soon were on Garrett, as it proposed an axisymmetric configuration that was considerably shorter than the others. John Becker later 

wrote that it “was the smallest, simplest, easiest to cool, and had the best structural approach of the three designs.” Moreover, Garrett had shown strong initiative 

through the leadership of its study manager, Anthony duPont.15

He was a member of the famous duPont family in the chemical industry. Casual 

and easygoing, he had already shown a keen eye for the technologies of the future. 

As early as 1954, as a student, he had applied for a patent on a wing made of 

engine, and in Nucci’s words, “we had to use our know-how in flow and combustion.” Having taken over from Ferri as company president, he won a contract from 

the Department of Transportation to study the aerodynamics of high-speed trains 

running in tubes.

“We had to retread everybody,” Nucci adds. Boeing held a federal contract to 

develop a supersonic transport; GASL studied its sonic boom. GASL also investigated the “parasol wing,” a low-drag design that rode atop its fuselage at the end of 

a pylon. There also was work on pollution for the local utility, Long Island Lighting Company, which hoped to reduce its smog-forming emissions. The company 

stayed alive, but its employment dropped from 80 people in 1967 to only 45 five 

years later.12

Marquardt made its own compromises. It now was building small rocket engines, 

including attitude-control thrusters for Apollo and later for the space shuttle. But it 

too left the field of hypersonics. Arthur Thomas had managed the company’s work 

on IFTV, and as he recalls, “I was chief engineer and assistant general manager. I got 

laid off. We laid off two-thirds of our people in one day.” He knew that there was no 

scramjet group he might join, but he hoped for the next-best thing: conventional 

ramjets, powering high-speed missiles. “I went all over the country,” he continues. 

“Everything in ramjet missiles had collapsed.” He had to settle for a job working 

with turbojets, at McDonnell Douglas in St. Louis.13

Did these people ever doubt the value of their work? “Never,” says Billig. Nucci, 

Ferri’s old friend, gives the same answer: “Never. He always had faith.” The problem 

they faced was not to allay any doubts of their own, but to overcome the misgivings 

of others and to find backers who would give them new funding. From time to time 

a small opportunity appeared. Then, as Billig recalls, “we were highly competitive. 

Who was going to get the last bits of money? As money got tighter, competition got 

stronger. I hope it was a friendly competition, but each of us thought he could do 

the job best.”14

Amid this dark night of hypersonic research, two candles still flickered. There 

was APL, where a small group continued to work on missiles powered by scramjets 

that were to burn conventional fuels. More significantly, there was the Hypersonic 

Propulsion Branch at NASA-Langley, which maintained itself as the one place where 

important work on hydrogen-fueled scramjets still could go forward. As scramjets 

died within the Air Force, the Langley group went ahead, first with its Hypersonic 

Research Engine (HRE) and then with more advanced airframe-integrated designs.

Scramjets at NASA-Langley

The road to a Langley scramjet project had its start at North American Aviation, 

builder of the X-15. During 1962 manager Edwin Johnston crafted a proposal to 

modify one of the three flight vehicles to serve as a testbed for hypersonic engines. The HRE concepts of three competing contractors. (NASA)
204

Facing the Heat Barrier: A History of Hypersonics

205

The Fading, the Comeback 

the X-15 could not be allowed to fly with them. Indeed, it soon stopped flying altogether. Thus, during 1968, it became clear that the HRE could survive only through 

a complete shift in focus to ground test.

Earlier plans had called for a hydrogen-cooled flightweight engine. Now the 

program’s research objectives were to be addressed using two separate wind-tunnel 

versions. Each was to have a diameter of 18 inches, with a configuration and flow 

path matching those of the earlier flight-rated concept. The test objectives then were 

divided between them.

A water-cooled Aerothermodynamic Integration Model (AIM) was to serve for 

hot-fire testing. Lacking provision for hydrogen cooling, it stood at the technical level of the General Electric and Pratt & Whitney test scramjets. In addition, 

continuing interest in flightweight hydrogen-cooled engine structures brought a 

requirement for the Structures Assembly Model (SAM), which did not burn fuel. 

It operated at high temperature in Langley’s eight-foot diameter High Temperature 

Structures Tunnel, which reached Mach 7.20

SAM arrived at NASA-Langley in August 1970. Under test, its inlet lip showed 

robustness for it stood up to the impact of small particles, some of which blocked 

thin hydrogen flow passages. Other impacts produced actual holes as large as 1/16 

composite materials. He flew 

as a co-pilot with Pan American, commemorating those 

days with a framed picture of 

a Stratocruiser airliner in his 

office. He went on to Douglas 

Aircraft, where he managed 

studies of Aerospaceplane. 

Then Clifford Garrett, who 

had a strong interest in scramjets, recruited him to direct 

his company’s efforts.16

NASA’s managers soon 

offered an opportunity to the 

HRE competitors. The Ordnance Aerophysics Laboratory was still in business, and any of them could spend a 

month there testing hardware—if they could build scramjet components on short 

notice. Drawing on $250,000 in company funds, DuPont crafted a full-scale HRE 

combustor in only sixty days. At OAL, it yielded more than five hours of test data. 

Neither GE nor Marquardt showed similar adroitness, while DuPont’s initiative 

suggested that the final HRE combustor would be easy to build. With this plus the 

advantages noted by Becker, Garrett won the contract. In July 1966 the program 

then moved into a phase of engine development and test.17

Number 66671 was flying routinely, and it proved possible to build a dummy 

HRE that could be mounted to the lower fin of that X-15. This led to a flighttest program that approached disaster in October 1967, when the test pilot Pete 

Knight flew to Mach 6.72. “We burned the engine off,” Knight recalls. “I was on 

my way back to Edwards; my concern was to get the airplane back in one piece.” 

He landed safely, but historian Richard Hallion writes that the airplane “resembled 

burnt firewood.... It was the closest any X-15 came to structural failure induced by 

heating.”18

Once again it went back to the shops, marked for extensive repair. Then in midNovember another X-15 was lost outright in the accident that killed its test pilot, 

Mike Adams. Suddenly the X-15 was down from three flight-rated airplanes to 

only one, and while Number 66671 returned to the flight line the following June, 

it never flew again. Nor would it fly again with the HRE. This dummy engine had 

set up the patterns of airflow that had caused the shock-impingement heating that 

had nearly destroyed it.19

In a trice then the HRE program was completely turned on its head. It had 

begun with the expectation of using the X-15 for flight test of advanced engines, 

at a moment when no such engines existed. Now Garrett was building them—but 

The selected HRE design. (NASA)

Test pilot William “Pete” Knight initiates his record flight, which reached Mach 6.72. (NASA)
206

Facing the Heat Barrier: A History of Hypersonics

207

The Fading, the Comeback 

inch in diameter. The lip nevertheless rode through the subsequent shock-impingement heating without coolant starvation or damage from overheating. This represented an important advance in scramjet technology, for it demonstrated the feasibility of crafting a flightweight fuel-cooled structure that could withstand foreign 

object damage along with very severe heating.21

AIM was also on the agenda. It reached its test center at Plum Brook, Ohio, in 

August 1971, but the facility was not ready. It took a year before the program undertook data runs, and then most of another year before the first run that was successful. 

Indeed, of 63 test runs conducted across 18 months, 42 returned little or no useful 

data. Moreover, while scramjet advocates had hoped to achieve shock-free flow, it 

certainly did not do this. In addition, only about half of the injected fuel actually 

burned. But shocks in the subsonic-combustion zone heated the downstream flow 

and unexpectedly enabled the rest of the fuel to burn. In Becker’s words, “without 

this bonanza, AIM performance would have been far below its design values.”22

The HRE was axisymmetric. A practical engine of this type would have been 

mounted in a pod, like a turbojet in an airliner. An airliner’s jet engines use only a 

small portion of the air that flows past the wings and fuselage, but scramjets have 

far less effectiveness. Therefore, to give enough thrust for acceleration at high Mach, 

they must capture and process as much as possible of the air flowing along the 

vehicle.

Podded engines like the HRE cannot do this. The axisymmetry of the HRE made 

it easy to study because it had a two-dimensional layout, but it was not suitable for 

an operational engine. The scramjet that indeed could capture and process most of 

the airflow is known as an airframe-integrated engine, in which much of the aircraft serves as part of the 

propulsion system. Its 

layout is three-dimensional and hence is more 

complex, but only an 

airframe-integrated concept has the additional 

power that can make it 

practical for propulsion.

Paper studies of airframe-integrated concepts began at Langley in 1968, breaking 

completely with those 

of HRE. These investigations considered the 

entire undersurface of a hypersonic aircraft as an element of the propulsion system. 

The forebody produced a strong oblique shock that precompressed the airflow prior 

to its entry into the inlet. The afterbody was curved and swept upward to form a 

half-nozzle. This concept gave a useful shape for the airplane while retaining the 

advantages of airframe-integrated scramjet operation.

Within the Hypersonic Propulsion Branch, John Henry 

and Shimer Pinckney developed the initial concept. Their 

basic installation was a module, 

rectangular in shape, with a 

number of them set side by 

side to encircle the lower fuselage and achieve the required 

high capture of airflow. Their 

inlet had a swept opening that 

angled backward at 48 degrees. 

This provided a cutaway that 

readily permitted spillage of 

airflow, which otherwise could 

choke the inlet when starting.

The bow shock gave greater 

compression of the flow at high Mach, thereby reducing the height of the cowl 

and the required size of the engine. At Mach 10 this reduction was by a factor of 

three. While this shock compressed the flow vertically, wedge-shaped sidewalls compressed it horizontally. This two-plane compression diminished changes in the inlet 

flow field with increasing Mach, making it possible to cover a broad Mach range in 

fixed geometry.

Like the inlet, the combustor was to cover a large Mach range in fixed geometry. 

This called for thermal compression, and Langley contracted with Antonio Ferri at 

New York University to conduct analyses. This brought Ferri back into the world of 

scramjets. The design called for struts as fuel injectors, swept at 48 degrees to parallel the inlet and set within the combustor flow path. They promised more effective 

fuel injection than the wall-mounted injectors of earlier designs.

The basic elements of the Langley concept thus included fixed geometry, airframe integration, a swept inlet, thermal compression, and use of struts for fuel 

injection. These elements showed strong synergism, for in addition to the aircraft 

undersurface contributing to the work of the inlet and nozzle, the struts also served 

as part of the inlet and thereby made it shorter. This happened because the flow 

from the inlet underwent further compression as it passed between the struts.23 Contributions to scramjet thrust from airframe integration. 

(NASA)

Airframe-integrated scramjet concept. (Garrett Corp.)
208

Facing the Heat Barrier: A History of Hypersonics

209

The Fading, the Comeback 

Experimental work paced the Langley effort as it went forward during the 1970s 

and much of the 1980s. Early observations, published in 1970, showed that struts 

were practical for a large supersonic combustor in flight at Mach 8. This work supported the selection of strut injection as the preferred mode.24

Initial investigations involved inlets and combustors that were treated as separate components. These represented preludes to studies made with complete engine 

modules at two critical simulated flight speeds: Mach 4 and Mach 7. At Mach 4 the 

inlet was particularly sensitive to unstarts. The inlet alone had worked well, as had 

the strut, but now it was necessary to test them together and to look for unpleasant surprises. The Langley researchers therefore built a heavily instrumented engine 

of nickel and tested it at GASL, thereby bringing new work in hypersonics to that 

center as well.

Mach 7 brought a different set of problems. Unstarts now were expected to 

be less of a difficulty, but it was necessary to show that the fuel indeed could mix 

and burn within the limited length of the combustor. Mach 7 also approached the 

limitations of available wind tunnels. A new Langley installation, the Arc-Heated 

Scramjet Test Facility, reached temperatures of 3,500oF and provided the appropriate flows.

Integration of scramjets with an aircraft. (NASA)

Airflow within an airframe-integrated scramjet. (NASA)

Fuel-injecting strut. Arrows show how hydrogen is injected either parallel or perpendicular to the 

flow. (Garrett Corporation)
210

Facing the Heat Barrier: A History of Hypersonics

211

The Fading, the Comeback 

inlet portion of the engine, but the remainder of the engine is unswept. In 

fact, the hardware is designed in sections so that inlet sweep can be changed 

(by substituting new inlet sidewalls) without removing the engine from the 

wind tunnel.

The Parametric Scramjet explored techniques for alleviating combustor-inlet 

interactions at Mach 4. The Step-Strut design also addressed this issue, mounting a single long internal strut fitted with fuel injectors, with a swept leading edge 

that resembled a staircase. 

Northam and Anderson wrote 

that it “was also tested at Mach 

4 and demonstrated good performance without combustorinlet interaction.”27

How, specifically, did Langley develop a workable scramjet? Answers remain classified, 

with Northam and Anderson 

noting that “several of the figures have no dimension on the 

axes and a discussion of the figures omits much of the detail.” 

A 1998 review was no more 

helpful. However, as early as 

1986 the Langley researchers 

openly published a plot showing data taken at Mach 4 and at 

Mach 7. Curves showed values 

of thrust and showed that the 

scramjets of the mid-1980s 

indeed could produce net thrust. Even at Mach 7, at which the thrust was less, these 

engines could overcome the drag of a complete vehicle and produce acceleration. 

In the words of Northam and Anderson, “at both Mach 4 and Mach 7 flight conditions, there is ample thrust for acceleration and cruise.”28

The Advent of NASP

With test engines well on their way in development, there was the prospect of 

experimental aircraft that might exercise them in flight test. Such a vehicle might 

come forth as a successor to Number 66671, the X-15 that had been slated to fly the 

Separate engines operated at GASL and Langley. Both used heat sink, with the 

run times being correspondingly short. Because both engines were designed for use 

in research, they were built for easy substitution of components. An inlet, combustor, nozzle, or set of fuel-injecting struts could be installed without having to modify 

the rest of the engine. This encouraged rapid prototyping without having to construct entirely new scramjets.

More than 70 runs at Mach 4 were made during 1978, first with no fuel injection to verify earlier results from inlet tests, and then with use of hydrogen. Simple 

theoretical calculations showed that “thermal choking” was likely, with heat addition in the combustor limiting the achievable flow rate, and indeed it appeared. 

Other problems arose from fuel injection. The engine used three struts, a main one 

on the centerline flanked by two longer ones, and fuel from these side struts showed 

poor combustion when injected parallel to the flow. Some unwanted inlet-combustor interactions sharply reduced the measured thrust. These occurred because the 

engine ingested boundary-layer flow from the top inner surface of the wind-tunnel 

duct. This simulated the ingestion of an aircraft boundary layer by a flight engine.

The thermal choking and the other interactions were absent when the engine ran 

very fuel-lean, and the goal of the researchers was to eliminate them while burning 

as much fuel as possible. They eased the problem of thermal choking by returning 

to a fuel-injection method that had been used on the HRE, with some fuel being 

injected downstream as the wall. However, the combustor-inlet interactions proved 

to be more recalcitrant. They showed up when the struts were injecting only about 

half as much fuel as could burn in the available airflow, which was not the formula 

for a high-thrust engine.25

Mach 7 brought its own difficulties, as the Langley group ran off 90 tests between 

April 1977 and February 1979. Here too there were inlet-combustor interactions, 

ranging from increased inlet spillage that added drag and reduced the thrust, to 

complete engine unstarts. When the latter occurred, the engine would put out good 

thrust when running lean; when the fuel flow increased, so did the measured force. 

In less than a second, though, the inlet would unstart and the measured thrust 

would fall to zero.26

No simple solution appeared capable of addressing these issues. This meant that 

in the wake of those tests, as had been true for more than a decade, the Langley 

group did not have a working scramjet. Rather, they had a research problem. They 

addressed it after 1980 with two new engines, the Parametric Scramjet and the 

Step-Strut Engine. The Parametric engine lacked a strut but was built for ease of 

modification. In 1986 the analysts Burton Northam and Griffin Anderson wrote:

This engine allows for easy variation of inlet contraction ratio, internal area 

ratio and axial fuel injection location. Sweep may be incorporated in the 

Performance of scramjets. Note that figures are missing 

from the axes. (NASA)
212

Facing the Heat Barrier: A History of Hypersonics

213

The Fading, the Comeback 

the 11 February meeting, as he drew extensively on the views of Watkins. Reagan 

showed strong interest and told the Chiefs that he wanted a written proposal. Robert 

McFarlane, Deputy to the National Security Advisor, already had begun to explore 

concepts for missile defense. During the next several weeks his associates took the 

lead in developing plans for a program and budget.32

On 23 March 1983 Reagan spoke to the nation in a televised address. He dealt 

broadly with issues of nuclear weaponry. Toward the end of the speech, he offered 

new thoughts:

“Let me share with you a vision of the future which offers hope. It is that 

we embark on a program to counter the awesome Soviet missile threat with 

measures that are defensive. Let us turn to the very strengths in technology 

that spawned our great industrial base and that have given us the quality of 

life we enjoy today.

What if free people could live secure in the knowledge that their 

security did not rest upon the threat of instant U.S. retaliation to deter a 

Soviet attack, that we could intercept and destroy strategic ballistic missiles 

before they reached our own soil or that of our allies?...

I call upon the scientific community in our country, those who gave us 

nuclear weapons, to turn their great talents now to the cause of mankind 

and world peace, to give us the means of rendering these nuclear weapons 

impotent and obsolete.”33

The ensuing Strategic Defense Initiative never deployed weapons that could 

shoot down a missile. Yet from the outset it proved highly effective in shooting 

down the nuclear freeze. That movement reached its high-water mark in May 1983, 

as a strengthened Democratic majority in the House indeed passed Markey’s resolution. But the Senate was still held by Republicans, and the freeze went no further. 

The SDI gave everyone something new to talk about. Reagan’s speech helped him 

to regain the initiative, and in 1984 he swept to re-election with an overwhelming 

majority.34

The SDI brought the prospect of a major upsurge in traffic to orbit, raising the 

prospect of a flood of new military payloads. SDI supporters asserted that some one 

hundred orbiting satellites could provide an effective strategic defense, although 

the Union of Concerned Scientists, a center of criticism, declared that the number 

would be as large as 2,400. Certainly, though, an operational missile defense was 

likely to place new and extensive demands on means for access to space.

Within the Air Force Systems Command, there already was interest in a nextgeneration single-stage-to-orbit launch vehicle that was to use the existing Space 

Shuttle Main Engine. Lieutenant General Lawrence Skantze, Commander of the 

HRE. An aircraft of this type indeed took shape before long, with the designation 

X-30. However, it did not originate purely as a technical exercise. Its background 

lay in presidential politics.

The 1980 election took place less than a year after the Soviets invaded Afghanistan. President Jimmy Carter had placed strong hope in arms control and had 

negotiated a major treaty with his Soviet counterpart, Leonid Brezhnev. But the 

incursion into Afghanistan took Carter by surprise and destroyed the climate of 

international trust that was essential for Senate ratification of this treaty. Reagan 

thus came to the White House with arms-control prospects on hold and with the 

Cold War once more in a deep freeze. He responded by launching an arms buildup 

that particularly included new missiles for Europe.29

Peace activist Randall Forsberg replied by taking the lead in calling for a nuclear 

freeze, urging the superpowers to halt the “testing, production and deployment of 

nuclear weapons” as an important step toward “lessening the risk of nuclear war.” 

His arguments touched a nerve within the general public, for within two years, 

support for a freeze topped 70 percent. Congressman Edward Markey introduced 

a nuclear-freeze resolution in the House of Representatives. It failed by a margin of 

only one vote, with Democratic gains in the 1982 mid-term elections making passage a near certainty. By the end of that year half the states in the Union adopted 

their own freeze resolutions, as did more than 800 cities, counties, and towns.30

To Reagan, a freeze was anathema. He declared that it “would be largely unverifiable.... It would reward the Soviets for their massive military buildup while preventing us from modernizing our aging and increasingly vulnerable forces.” He asserted 

that Moscow held a “present margin of superiority” and that a freeze would leave 

America “prohibited from catching up.”31

With the freeze ascendant, Admiral James Watkins, the Chief of Naval Operations, took a central role in seeking an approach that might counter its political 

appeal. Exchanges with Robert McFarlane and John Poindexter, deputies within the 

National Security Council, drew his thoughts toward missile defense. Then in January 1983 he learned that the Joint Chiefs were to meet with Reagan on 11 February. 

As preparation, he met with a group of advisors that included the physicist Edward 

Teller.

Trembling with passion, Teller declared that there was enormous promise in a 

new concept: the x-ray laser. This was a nuclear bomb that was to produce intense 

beams of x-rays that might be aimed to destroy enemy missiles. Watkins agreed that 

the broad concept of missile defense indeed was attractive. It could introduce a new 

prospect: that America might counter the Soviet buildup, not with a buildup of its 

own but by turning to its strength in advanced technology.

Watkins succeeded in winning support from his fellow Joint Chiefs, including 

the chairman, General John Vessey. Vessey then gave Reagan a half-hour briefing at 
214

Facing the Heat Barrier: A History of Hypersonics

215

The Fading, the Comeback 

1982 as a group of ramjet specialists met with Richard De Lauer, the Undersecretary 

of Defense Research and Engineering. They urged him to keep the field alive with 

enough new funds to prevent them from having to break up their groups. De Lauer 

responded with letters that he sent to the Navy, Air Force, and DARPA, asking them 

to help.38

This provided an opening for Tony duPont, who had designed the HRE. He 

had taken a strong interest in combined-cycle concepts and decided that the scramlace was the one he preferred. It was to eliminate the big booster that every ramjet 

needed, by using an ejector, but experimental versions weren’t very powerful. 

DuPont thought he could do better by using the HRE as a point of departure, as he 

added an auxiliary inlet for LACE and a set of ejector nozzles upstream of the combustor. He filed for a patent on his engine in 1970 and won it two years later.39

In 1982 he still believed in it, and he learned that Anthony Tether was the 

DARPA man who had been attending TAV meetings. The two men met several 

times, with Tether finally sending him up to talk with Cooper. Cooper listened to 

duPont and sent him over to Robert Williams, one of DARPA’s best aerodynamicists. Cooper declares that Williams “was the right guy; he knew the most in this 

area. This wasn’t his specialty, but he was an imaginative fellow.”40

Williams had come up within the Navy, working at its David Taylor research 

center. His specialty was helicopters; he had initiated studies of the X-wing, which 

was to stop its rotor in midair and fly as a fixed-wing aircraft. He also was interested in high-speed flight. He had studied a missile that was to fight what the Navy 

Anthony duPont’s engine. (GASL)

Air Force Systems Command’s Aeronautical Systems Division (ASD), 

launched work in this area early in 

1982 by directing the ASD planning 

staff to conduct an in-house study 

of post-shuttle launch vehicles. It 

then went forward under the leadership of Stanley Tremaine, the ASD’s 

Deputy for Development Planning, 

who christened these craft as Transatmospheric Vehicles. In December 

1984 Tremaine set up a TAV Program 

Office, directed by Lieutenant Colonel Vince Rausch.35

Moreover, General Skantze was 

advancing into high-level realms of 

command, where he could make his 

voice heard. In August 1982 he went 

to Air Force Headquarters, where 

he took the post of Deputy Chief of 

Staff for Research, Development, and 

Acquisition. This gave him responsibility for all Air Force programs in these areas. In October 1983 he pinned on his 

fourth star as he took an appointment as Air Force Vice Chief of Staff. In August 

1984 he became Commander of the Air Force Systems Command.36

He accepted these Washington positions amid growing military disenchantment 

with the space shuttle. Experience was showing that it was costly and required a 

long time to prepare for launch. There also was increasing concern for its safety, 

with a 1982 Rand Corporation study flatly predicting that as many as three shuttle 

orbiters would be lost to accidents during the life of the program. The Air Force was 

unwilling to place all its eggs in such a basket. In February 1984 Defense Secretary 

Caspar Weinberger approved a document stating that total reliance on the shuttle 

“represents an unacceptable national security risk.” Air Force Secretary Edward 

Aldridge responded by announcing that he would remove 10 payloads from the 

shuttle beginning in 1988 and would fly them on expendables.37

Just then the Defense Advanced Research Projects Agency was coming to the 

forefront as an important new center for studies of TAV-like vehicles. DARPA was 

already reviving the field of flight research with its X-29, which featured a forwardswept wing along with an innovative array of control systems and advanced materials. Robert Cooper, DARPA’s director, held a strong interest in such projects and 

saw them as a way to widen his agency’s portfolio. He found encouragement during 

Transatmospheric Vehicle concepts, 1984. 

(U.S. Air Force)
216

Facing the Heat Barrier: A History of Hypersonics

217

The Fading, the Comeback 

Council. Keyworth recalls that “here were people who normally would ask questions for hours. But after only about a half-hour, David Packard said, ‘What’s keeping us? Let’s do it!’” Packard was Deputy Secretary of Defense.45

During 1985, as Copper Canyon neared conclusion, the question arose of 

expanding the effort with support from NASA and the Air Force. Cooper attended 

a classified review and as he recalls, “I went into that meeting with a high degree 

of skepticism.” But technical presentations brought him around: “For each major 

problem, there were three or four plausible ways to deal with it. That’s extraordinary. Usually it’s—‘Well, we don’t know exactly how we’ll do it, but we’ll do it.’ Or, 

‘We have a way to do it, which may work.’ It was really a surprise to me; I couldn’t 

pick any obvious holes in what they had done. I could find no reason why they 

couldn’t go forward.”46

Further briefings followed. Williams gave one to Admiral Watkins, whom 

Cooper describes as “very supportive, said he would commit the Navy to support of 

the program.” Then in July, Cooper accompanied Williams as they gave a presentation to General Skantze.

They displayed their viewgraphs and in Cooper’s words, “He took one look at 

our concept and said, ‘Yeah, that’s what I meant. I invented that idea.’” Not even 

the stars on his shoulders could give him that achievement, but his endorsement 

reflected the fact that he was dissatisfied with the TAV studies. He had come away 

appreciating that he needed something better than rocket engines—and here it was. 

“His enthusiasm came from the fact that this was all he had anticipated,” Cooper 

continues. “He felt as if he owned it.”

Skantze wanted more than viewgraphs. He wanted to see duPont’s engine in 

operation. A small version was under test at GASL, without LACE but definitely 

with its ejector, and one technician had said, “This engine really does put out static 

thrust, which isn’t obvious for a ramjet.” Skantze saw the demonstration and came 

away impressed. Then, Williams adds, “the Air Force system began to move with 

the speed of a spaceplane. In 

literally a week and a half, the 

entire Air Force senior command was briefed.”

Later that year the Secretary 

of Defense, Caspar Weinberger, 

granted a briefing. With him 

were members of his staff, 

along with senior people from 

NASA and the military service. 

After giving the presentation, 

Williams recalls that “there was 

silence in the room. The Seccalled the “outer air battle,” which might use a scramjet. This had brought him 

into discussions with Fred Billig, who also worked for the Navy and helped him to 

learn his hypersonic propulsion. He came to DARPA in 1981 and joined its Tactical Technologies Office, where he became known as the man to see if anyone was 

interested in scramjets.41

Williams now phoned duPont and gave him a test: “I’ve got a very ambitious 

problem for you. If you think the airplane can do this, perhaps we can promote 

a program. Cooper has asked me to check you out.” The problem was to achieve 

single-stage-to-orbit flight with a scramjet and a suite of heat-resistant materials, and duPont recalls his response: “I stayed up all night; I was more and more 

intrigued with this. Finally I called him back: ‘Okay, Bob, it’s not impossible. Now 

what?’”42

DuPont had been using a desktop computer, and Williams and Tether responded 

to his impromptu calculations by giving him $30,000 to prepare a report. Soon 

Williams was broadening his circle of scramjet specialists by talking with old-timers 

such as Arthur Thomas, who had been conducting similar studies a quarter-century 

earlier, and who quickly became skeptical. DuPont had patented his propulsion 

concept, but Thomas saw it differently: “I recognized it as a Marquardt engine. 

Tony called it the duPont cycle, which threw me off, but I recognized it as our 

engine. He claimed he’d improved it.” In fact, “he’d made a mistake in calculating 

the heat capacity of air. So his engine looked so much better than ours.”

Thomas nevertheless signed on to contribute to the missionary work, joining 

Williams and duPont in giving presentations to other conceptual-design groups. 

At Lockheed and Boeing, they found themselves talking to other people who knew 

scramjets. As Thomas recalls, “The people were amazed at the component efficiencies that had been assumed in the study. They got me aside and asked if I really 

believed it. Were these things achievable? Tony was optimistic everywhere: on 

mass fraction, on air drag of the vehicle, on inlet performance, on nozzle performance, on combustor performance. The whole thing, across the board. But what 

salved our conscience was that even if these weren’t all achieved, we still could have 

something worth while. Whatever we got would still be exciting.”43

Williams recalls that in April 1984, “I put together a presentation for Cooper 

called ‘Resurrection of the Aerospaceplane.’ He had one hour; I had 150 slides. 

He came in, sat down, and said Go. We blasted through those slides. Then there 

was silence. Cooper said, ‘I want to spend a day on this.’” After hearing additional briefings, he approved a $5.5-million effort known as Copper Canyon, which 

brought an expanded program of studies and analyses.44

Copper Canyon represented an attempt to show how the SDI could achieve 

its access to space, and a number of high-level people responded favorably when 

Cooper asked to give a briefing. He and Williams made a presentation to George 

Keyworth, Reagan’s science advisor. They then briefed the White House Science Initial version of the duPont engine under test at GASL. 

(GASL)
218

Facing the Heat Barrier: A History of Hypersonics

219

The Fading, the Comeback 

used it to rescue NASP. He led the Space Council to recommend proceeding with 

the program under a reduced but stable budget, and with a schedule slip. This plan 

won acceptance, giving the program leeway to face a new issue: excessive technical 

optimism.49

During 1984, amid the Copper Canyon activities, Tony duPont devised a conceptual configuration that evolved into the program’s baseline. It had a gross weight 

of 52,650 pounds, which included a 2,500-pound payload that it was to carry to 

polar orbit. Its weight of fuel was 28,450 pounds. The propellant mass fraction, the 

ratio of these quantities, then was 0.54.50

The fuel had low density and was bulky, demanding high weight for the tankage and airframe. To save weight, duPont’s concept had no landing gear. It lacked 

reserves of fuel; it was to reach orbit by burning its last drops. Once there it could 

not execute a controlled deorbit, for it lacked maneuvering rockets as well as fuel 

and oxidizer for them. DuPont also made no provision for a reserve of weight to 

accommodate normal increases during development.51

Williams’s colleagues addressed these deficiencies, although they continued to 

accept duPont’s optimism in the areas of vehicle drag and engine performance. 

The new concept had a gross weight of 80,000 pounds. Its engines gave a specific 

impulse of 1,400 seconds, averaged over the trajectory, which corresponded to a 

mean exhaust velocity of 45,000 feet per second. (That of the SSME was 453.5 seconds in vacuum, or 14,590 feet per second.) The effective velocity increase for the 

X-30 was calculated at 47,000 feet per second, with orbital velocity being 25,000 feet 

X-30 concept of 1985. (NASA)

retary said, ‘Interesting,’ and turned to his staff. Of course, all the groundwork had 

been laid. All of the people there had been briefed, and we could go for a yes-or-no 

decision. We had essentially total unanimity around the table, and he decided that 

the program would proceed as a major Defense Department initiative. With this, 

we moved immediately to issue requests for proposal to industry.”47

In January 1986 the TAV effort was formally terminated. At Wright-Patterson 

AFB, the staff of its program office went over to a new Joint Program Office that now 

supported what was called the National Aerospace Plane. It brought together representatives from the Air Force, Navy, and NASA. Program management remained 

at DARPA, where Williams retained his post as the overall manager.48

In this fashion, NASP became a significant federal initiative. It benefited from 

a rare alignment of the political stars, for Reagan’s SDI cried out for better launch 

vehicles and Skantze was ready to offer them. Nor did funding appear to be a problem, at least initially. Reagan had shown favor to aerospace through such acts as 

approving NASA’s space station in 1984. Pentagon spending had surged, and DARPA’s Cooper was asserting that an X-30 might be built for an affordable cost.

Yet NASP was a leap into the unknown. Its scramjets now were in the forefront 

but not because the Langley research had shown that they were ready. Instead they 

were a focus of hope because Reagan wanted SDI, SDI needed better access to 

space, and Skantze wanted something better than rockets.

The people who were making Air Force decisions, such as Skantze, did not know 

much about these engines. The people who did know them, such as Thomas, were 

well aware of duPont’s optimism. There thus was abundant opportunity for high 

hope to give way to hard experience.

The Decline of NASP

NASP was one of Reagan’s programs, and for a time it seemed likely that it 

would not long survive the change in administrations after he left office in 1989. 

That fiscal year brought a high-water mark for the program, as its budget peaked at 

$320 million. During the spring of that year officials prepared budgets for FY 1991, 

which President George H. W. Bush would send to Congress early in 1990. Military 

spending was already trending downward, and within the Pentagon, analyst David 

Chu recommended canceling all Defense Department spending for NASP. The new 

Secretary of Defense, Richard Cheney, accepted this proposal. With this, NASP 

appeared dead.

NASP had a new program manager, Robert Barthelemy, who had replaced Williams. Working through channels, he found support in the White House from Vice 

President Dan Quayle. Quayle chaired the National Space Council, which had been 

created by law in 1958 and that just then was active for the first time in a decade. He 
220

Facing the Heat Barrier: A History of Hypersonics

221

The Fading, the Comeback 

rocket stages of NASA and calculating their values of propellant mass fraction if 

both their hydrogen and oxygen tanks were filled with NASP fuel. This was slush 

hydrogen, a slurry of the solid and liquid. The stages are the S-II and S-IVB of 

Apollo and the space shuttle’s external tank. Liquid hydrogen has 1/16 the density 

of liquid oxygen. With NASP slush having 1.16 times the density of liquid hydrogen,55 the propellant mass fractions are as follows:56

S-IVB, third stage of the Saturn V 0.722

S-II, second stage of the Saturn V 0.753

External Tank 0.868

The S-II, which comes close to Kerrebrock’s value of 0.75, was an insulated shell 

that mounted five rocket engines. It withstood compressive loads along its length 

that resulted from the weight of the S-IVB and the Apollo moonship but did not 

require reinforcement to cope with major bending loads. It was constructed of aluminum alloy and lacked landing gear, thermal protection, wings, and a flight deck.

How then did NASP offer an X-30 concept that constituted a true hypersonic 

airplane rather than a mere rocket stage? The answer lay in adding weight to the 

fuel, which boosted the propellant mass fraction. The 

vehicle was not to reach 

orbit entirely on slushfueled scramjets but was to 

use a rocket for final ascent. 

It used tanked oxygen—

with nearly 14 times the 

density of slush hydrogen. 

In addition, design requirements specified a tripropellant system that was to 

burn liquid methane during 

the early part of the flight. 

This fuel had less energy 

than hydrogen, but it too 

added weight because it was 

relatively dense. The recommended mix called for 69 

percent hydrogen, 20 percent oxygen, and 11 percent 

methane.57

per second; the difference represented loss due to drag. This version of the X-30 was 

designated the “government baseline” and went to the contractors for further study.52

The initial round of contract awards was announced in April 1986. Five airframe 

firms developed new conceptual designs, introducing their own estimates of drag 

and engine performance along with their own choices of materials. They gave the 

following weight estimates for the X-30:

Rockwell International 175,000 pounds

McDonnell Douglas 245,000

General Dynamics 280,000

Boeing 340,000

Lockheed 375,000

A subsequent downselection, in October 1987, eliminated the two heaviest concepts while retaining Rockwell, McDonnell Douglas, and General Dynamics for 

further work.53

What brought these weight increases? Much of the reason lay in a falloff in 

estimated engine performance, which fell as low as 1,070 seconds of averaged specific impulse. New estimates of drag pushed the required effective velocity increase 

during ascent to as much as 

52,000 feet per second.

A 1989 technical review, 

sponsored by the National 

Research Council, showed 

what this meant. The chairman, Jack Kerrebrock, was an 

experienced propulsion specialist from MIT. His panel 

included other men of similar 

background: Seymour Bogdonoff of Princeton, Artur 

Mager of Marquardt, Frank 

Marble from Caltech. Their 

report stated that for the X-30 

to reach orbit as a single stage, 

“a fuel fraction of approximately 0.75 is required.”54

One gains insight by considering three hydrogen-fueled 

Evolution of the X-30. The government baseline of 1986 had 

Isp of 1,400 seconds, delta-V to reach orbit of 47,000 feet per 

second, and propellant mass fraction of 0.54. Its 1992 counterpart had less Isp, more drag, propellant mass fraction of 0.75, 

and could not reach orbit. (NASP National Program Office)

X-30 concept of 1990, which had grown considerably.

(U.S. Air Force)
222

Facing the Heat Barrier: A History of Hypersonics

223

The Fading, the Comeback 

than half—were to be achieved by September 1993. The situation was particularly 

worrisome in the critical area of structures and materials, for which only six of 19 

milestones were slated for completion. The GAO therefore recommended delaying 

a commitment to mainstream development “until critical technologies are developed and demonstrated.”61

The DSB concurred, highlighting specific technical deficiencies. The most 

important involved the prediction of scramjet performance and of boundary-layer 

transition. In the latter, an initially laminar or smoothly flowing boundary layer 

becomes turbulent. This brings large increases in heat transfer and skin friction, a 

major source of drag. The locations of transition thus had to be known.

The scramjet-performance problem arose because of basic limitations in the 

capabilities of ground-test facilities. The best of them could accommodate a complete engine, with inlet, combustor, and nozzle, but could conduct tests only below 

Mach 8. “Even at Mach 8,” the DSB declared, “the scramjet cycle is just beginning 

to be established and consequently, there is uncertainty associated with extrapolating the results into the higher Mach regime. At speeds above Mach 8, only small 

components of the scramjet can be tested.” This brought further uncertainty when 

predicting the performance of complete engines.

Boundary-layer transition to turbulence also demanded attention: “It is essential 

to understand the boundary-layer behavior at hypersonic speeds in order to ensure 

thermal survival of the airplane structure as designed, as well as to accurately predict 

the propulsion system performance and airplane drag. Excessive conservatism in 

boundary-layer predictions will lead to an overweight design incapable of achieving 

[single stage to orbit], while excessive optimism will lead to an airplane unable to 

survive in the hypersonic flight environment.”

The DSB also showed strong concern over issues of control in flight of the X30 and its engines. These were not simple matters of using ailerons or pushing 

throttles. The report stated that “controllability issues for NASP are so complex, 

so widely ranging in dynamics and frequency, and so interactive between technical 

disciplines as to have no parallels in aeronautical history...the most fundamental 

initial requirements for elementary aircraft control are not yet fully comprehended.” 

An onboard computer was to manage the vehicle and its engines in flight, but an 

understanding of the pertinent forces and moments “is still in an embryonic state.” 

Active cooling of the vehicle demanded a close understanding of boundary-layer 

transition. Active cooling of the engine called for resolution of “major uncertainties...connected with supersonic burning.” In approaching these issues, “very great 

uncertainties exist at a fundamental level.”

The DSB echoed the GAO in calling for extensive additional research before 

proceeding into mainstream development of the X-30:

In 1984, with optimism at its height, Cooper had asserted that the X-30 would 

be the size of an SR-71 and could be ready in three years. DuPont argued that his 

concept could lead to a “5-5-50” program by building a 50,000-pound vehicle in 

five years for $5 billion.58 Eight years later, in October 1990, the program had a 

new chosen configuration. It was rectangular in cross section, with flat sides. Three 

scramjet engines were to provide propulsion. Two small vertical stabilizers were at 

the rear, giving better stability than a single large one. A single rocket engine of 

approximately 60,000 pounds of thrust, integrated into the airframe, completed the 

layout. Other decisions selected the hot structure as the basic approach to thermal 

protection. The primary structure was to be of titanium-matrix composite, with 

insulated panels of carbon to radiate away the heat.59

This 1990 baseline design showed little resemblance to its 1984 ancestor. As 

revised in 1992, it no longer was to fly to a polar orbit but would take off on a 

due-east launch from Kennedy Space Center, thereby gaining some 1,340 feet per 

second of launch velocity. Its gross weight was quoted at 400,000 pounds, some 

40 percent heavier than the General Dynamics weight that had been the heaviest 

acceptable in the 1987 downselect. Yet even then the 1992 concept was expected to 

fall short of orbit by some 3,000 feet per second. An uprated version, with a gross 

weight of at least 450,000 pounds, appeared necessary to reach orbital velocity. The 

prospective program budget came to $15 billion or more, with the time to first 

flight being eight to ten years.60

During 1992 both the Defense Science Board (DSB) and Congress’s General 

Accounting Office (GAO) conducted major program reviews. The immediate issue 

was whether to proceed as planned by making a commitment that would actually 

build and fly the X-30. Such a decision would take the program from its ongoing 

phase of research and study into a new phase of mainstream engineering development.

Both reviews focused on technology, but international issues were in the background, for the Cold War had just ended. The Soviet Union had collapsed in 1991, 

with communists falling from power while that nation dissolved into 15 constituent 

states. Germany had already reunified; the Berlin Wall had fallen, and the whole of 

Eastern Europe had won independence from Moscow. The western border of Russia 

now approximated that of 1648, at the end of the Thirty Years’ War. Two complete 

tiers of nominally independent nations now stood between Russia and the West.

These developments greatly diminished the military urgency of NASP, while the 

reviews’ conclusions gave further reason to reduce its priority. The GAO noted that 

program managers had established 38 technical milestones that were to be satisfied 

before proceeding to mainstream development. These covered the specific topics of 

X-30 design, propulsion, structures and materials, and use of slush hydrogen as a 

fuel. According to the contractors themselves, only 17 of those milestones—fewer 
224

Facing the Heat Barrier: A History of Hypersonics

225

The Fading, the Comeback 

1 AIAA Paper 93-2329.

2 Johns Hopkins APL Technical Digest, Vol. 13, No. 1 (1992), pp. 63-65.

3 Hallion, Hypersonic, pp. 754-55; Harshman, “Design and Test.”

4 Waltrup et al., “Supersonic,” pp. 42-18 to 42-19; DTIC AD-386653.

5 Hallion, Hypersonic, pp. VI-xvii to VI-xx; “Scramjet Flight Test Program” (Marquardt brochure, 

September 1965); DTIC AD-388239.

6 Hallion, Hypersonic, pp. VI-xiv, 780; “Report of the USAF Scientific Advisory Board Aerospace 

Vehicles Panel,” February 1966.

7 Hallion, Hypersonic, p. 780.

8 NASA SP-2000-4518, p. 63; NASA SP-4303, pp. 125-26.

9 AIAA Paper 93-2328; interoffice memo (GASL), E. Sanlorenzo to L. Nucci, 24 October 1967.

10 Mackley, “Historical”; DTIC AD-393374.

11 Journal of Aircraft, January-February 1968, p. 3.

12 Author interviews, Louis Nucci, 13 November 1987 and 24 June 1988 (includes quotes). 

Folder 18649, NASA Historical Reference Collection, NASA History Division, Washington, 

D.C. 20546. 

13 Author interviews, Arthur Thomas, 24 September 1987 and 24 June 1988 (includes quotes). 

Folder 18649, NASA Historical Reference Collection, NASA History Division, Washington, 

D.C. 20546.

14 Author interviews: Fred Billig, 27 June 1987 and Louis Nucci, 24 June 1988. Folder 18649, 

NASA Historical Reference Collection, NASA History Division, Washington, D.C. 20546.

15 Hallion, Hypersonic, pp. 756-78; Miller, X-Planes, pp. 189-90.

16 Author interview, Anthony duPont, 23 November 1987. Folder 18649, NASA Historical 

Reference Collection, NASA History Division, Washington, D.C. 20546.

17 Hallion, Hypersonic, pp. 774-78; “Handbook of Texas Online: Ordnance Aerophysics Laboratory,” internet website. 

18 Hallion, Hypersonic, pp. 142-45; Miller, X-Planes, pp. 190-91, 194-95; NASA SP-4303, pp. 

121-22; author interview, William “Pete” Knight, 24 September 1987. Folder 18649, NASA 

Historical Reference Collection, NASA History Division, Washington, D.C. 20546.

19 NASA SP-2000-4518, pp. 59-60.

20 Hallion, Hypersonic, pp. 792-96; AIAA Paper 93-2323; NASA TM X-2572.

21 Hallion, Hypersonic, pp. 798-802; AIAA Paper 93-2323; NASA TM X-2572.

22 Hallion, Hypersonic, pp. 802-22 (quote, p. 818); Mackley, “Historical.”

23 Waltrup et al., “Supersonic,” pp. 42-13 to 42-14; NASA SP-292, pp. 157-77; NASA TM X2895; Astronautics & Aeronautics, February 1978, pp. 38-48; AIAA Paper 86-0159.

24 AIAA Papers 70-715, 75-1212.

25 AIAA Papers 79-7045, 98-2506.

26 AIAA Paper 79-7045.

27 AIAA Paper 86-0159 (quotes, p. 7).

We have concluded [that] fundamental uncertainties will continue to 

exist in at least four critical areas: boundary-layer transition; stability and 

controllability; propulsion performance; and structural and subsystem 

weight. Boundary-layer transition and scramjet performance cannot 

be validated in existing ground-test facilities, and the weight estimates 

have insufficient reserves for the inevitable growth attendant to material 

allowables, fastening and joining, and detailed configuration issues.... 

Using optimistic assumptions on transition and scramjet performance, and 

the present weight estimates on material performance and active cooling, 

the vehicle design does not yet close; the velocity achieved is short of orbital 

requirements.62

Faced with the prospect that the flight trajectory of the X-30 would merely 

amount to a parabola, budget makers turned the curve of program funding into 

a parabola as well. The total budget had held at close to $250 million during FY 

1990 and 1991, falling to $205 million in 1992. But in 1993 it took a sharp dip 

to $140 million. The NASP National Program Office tried to rescue the situation 

by proposing a six-year program with a budget of $2 billion, called Hyflite, that 

was to conduct a series of unmanned flight tests. The Air Force responded with a 

new technical group, the Independent Review Team, that turned thumbs down on 

Hyflite and called instead for a “minimum” flight test program. Such an effort was 

to address the key problem of reducing uncertainties in scramjet performance at 

high Mach.

The National Program Office came back with a proposal for a new program 

called HySTP. Its budget request came to $400 million over five years, which would 

have continued the NASP effort at a level only slightly higher than its allocation 

of $60 million for FY 1994. Yet even this minimal program budget proved to be 

unavailable. In January 1995 the Air Force declined to approve the HySTP budget 

and initiated the formal termination of the NASP program.63

In this fashion, NASP lived and died. Like SDI and the space station, one could 

view it as another in a series of exercises in Reaganesque optimism that fell short. Yet 

from the outset, supporters of NASP had emphasized that it was to make important 

contributions in such areas as propulsion, hypersonic aerodynamics, computational 

fluid dynamics, and materials. The program indeed did these things and thereby 

laid groundwork for further developments.
226

Facing the Heat Barrier: A History of Hypersonics

227

The Fading, the Comeback 

53 Schweikart, Quest, pp. 51, 199-200; Miller, X-Planes, p. 310.

54 Air Force Studies Board, Hypersonic (quote, p. 5).

55 Sutton, Rocket, pp. 170-71; AIAA Paper 89-5014.

56 NASA SP-2000-4029, p. 284; Jenkins, Space Shuttle, pp. 421, 424.

57 DTIC ADB-197189, Item 063.

58 Aerospace America, June 1984, p. 1.

59 Aviation Week: 29 October 1990, pp. 36-37, 46; 1 April 1991, p. 80.

60 AIAA Paper 95-6031.

61 General Accounting Office Report NSIAD-93-71.

62 DTIC ADA-274530.

63 Schweikart, Quest, table opposite p. 182; AIAA Paper 95-6031.

28 Ibid. (quotes, pp. 1, 7); AIAA Paper 98-2506.

29 Malia, Soviet, pp. 378-80; Walker, Cold War, pp. 250-57.

30 Fitzgerald, Way Out, pp. 179-82, 190-91 (quotes, p. 180).

31 “Address by the President to the Nation.” Washington, DC: The White House, 23 March 1983.

32 Baucom, Origins, pp. 184-92; Broad, Teller’s War, pp. 121-24; Fitzgerald, Way Out, pp. 195-98, 

200-02.

33 “Address by the President to the Nation.” Washington, DC: The White House, 23 March 1983.

34 Fitzgerald, Way Out, p. 225.

35 Schweikart, Quest, pp. 20-22; Hallion, Hypersonic, pp. 1336-46; AIAA Paper 84-2414; Air 

Power History, Spring 1994, 39.

36 “Biography: General Lawrence A. Skantze.” Washington, DC: Office of Public Affairs, Secretary of the Air Force, August 1984.

37 Science, 29 June 1984, pp. 1407-09 (includes quote).

38 Miller, X-Planes, ch. 33; author interview, Robert Cooper, 12 May 1986. Folder 18649, NASA 

Historical Reference Collection, NASA History Division, Washington, D.C. 20546.

39 U.S. Patent 3,690,102 (duPont). Also author interview, Anthony duPont, 23 November 1987. 

Folder 18649, NASA Historical Collection, NASA History Division, Washington, D.C. 

20546. 

40 Author interviews, Robert Cooper, 12 May 1986 (includes quote); Anthony duPont, 23 November 1987. Folder 18649, NASA Historical Reference Collection, NASA History Division, 

Washington, D.C. 20546.

41 Schweikart, Quest, p. 20; author interview, Robert Williams, 1 May 1986 and 23 November 

1987. Folder 18649, NASA Historical Reference Collection, NASA History Division, Washington, D.C. 20546.

42 Author interviews, Robert Williams, 1 May 1986 and 23 November 1987. Folder 18649, 

NASA Historical Reference Collection, NASA History Division, Washington, D.C. 20546. 

43 Author interviews, Arthur Thomas, 24 September 1987 and 22 July 1988. Folder 18649, 

NASA Historical Reference Collection, NASA History Divison, Washington, D.C. 20546.

44 Author interview, Robert Williams, 1 May 1986. Folder 18649, NASA Historical Reference 

Collection, NASA History Division, Washington, D.C. 20546.

45 Author interview, George Keyworth, 23 May 1986. Folder 18469, NASA Historical Reference 

Collection, NASA History Division, Washington, D.C. 20546.

46 Author interview, Robert Cooper,12 May 1986. Folder 18469, NASA Historical Reference 

Collection, NASA History Division, Washington, D.C. 20546.

47 Ibid.; author interview, Robert Williams, 1 May 1986. Folder 18469, NASA Historical Reference Collection, NASA History Division, Washington, D.C. 20546. 

48 Schweikart, Quest, pp. 47-51; Hallion, Hypersonic, pp. 1345-46, 1351, 1364-65.

49 Schweikart, Quest, pp. 161-62.

50 AIAA Paper 95-6031.

51 Schweikart, Quest, pp. 29, 199.

52 AIAA Paper 95-6031. SSME: Rocketdyne Report RI/RD87-142.

NASP was founded on optimism, but it involved a good deal more than blind 

faith. Key technical areas had not been properly explored and offered significant 

prospects of advance. These included new forms of titanium, along with the use of 

an ejector to eliminate the need for an auxiliary engine as a separate installation, for 

initial boost of a scramjet. There also was the highly promising field of computational fluid dynamics (CFD), which held the prospect of supplementing flight test 

and work in wind tunnels with sophisticated mathematical simulation.

Still NASP fell short, and there were reasons. CFD proved not to be an exact science, particularly at high Mach. Investigators worked with the complete equations 

of fluid mechanics, which were exact, but were unable to give precise treatments 

in such crucial areas as transition to turbulence and the simulation or modeling of 

turbulence. Their discussions introduced approximations that took away the accuracy and left NASP with more drag and less engine performance than people had 

sought.

In the field of propulsion, ejectors had not been well studied and stood as a topic 

that was ripe for deeper investigation. Even so, the ejectors offered poor performance 

at the outset, and subsequent studies did not bring substantial improvements. This 

was unfortunate, for use of a highly capable ejector was a key feature of Anthony 

duPont’s patented engine cycle, which had provided technical basis for NASP.

With drag increasing and engine performance falling off, metallurgists might 

have saved the day by offering new materials. They indeed introduced Beta-21S 

titanium, which approached the heat resistance of Rene 41, the primary structural 

material of Dyna-Soar, but had only half the density. Yet even this achievement was 

not enough. Structural designers needed still more weight saving, and while they 

experimented with new types of beryllium and carbon-carbon, they came up with 

no significant contributions to the state of the art.

Aerodynamics

In March 1984, with the Copper Canyon studies showing promise, a classified 

program review was held near San Diego. In the words of George Baum, a close 

229

Why NASP

Fell Short

8
230

Facing the Heat Barrier: A History of Hypersonics

231

Why NASP Fell Short

CD CL L/D

Experimental data 0.03676 0.03173 1.158

Numerical results 0.03503 0.02960 1.183

Percent error 4.71 6.71 2.16

(Source: AIAA Paper 85-1509)

In that year the state of the art permitted extensive treatments of scramjets. 

Complete three-dimensional simulations of inlets were available, along with twodimensional discussions of scramjet flow fields that covered the inlet, combustor, 

and nozzle. In 1984 Fred Billig noted that simulation of flow through an inlet using 

complete Navier-Stokes equations typically demanded a grid of 80,000 points and 

up to 12,000 time steps, with each run demanding four hours on a Control Data 

Cyber 203 supercomputer. A code adapted for supersonic flow was up to a hundred 

times faster. This made it useful for rapid surveys of a number of candidate inlets, 

with full Navier-Stokes treatments being reserved for a few selected choices.4

Availability of test facilities. Continuous-flow wind tunnels are far below the requirements of realistic simulation of full-size aircraft in flight. Impulse facilities, such as shock tunnels, come close to 

the requirements but are limited by their very short run times. (NASA)

associate of Robert Williams, “We had to put together all the technology pieces to 

make it credible to the DARPA management, to get them to come out to a meeting 

in La Jolla and be willing to sit down for three full days. It wasn’t hard to get people 

out to the West Coast in March; the problem was to get them off the beach.”

One of the attendees, Robert Whitehead of the Office of Naval Research, gave a 

talk on CFD. Was the mathematics ready; were computers at hand? Williams recalls 

that “he explained, in about 15 minutes, the equations of fluid mechanics, in a 

memorable way. With a few simple slides, he could describe their nature in almost 

an offhand manner, laying out these equations so the computer could solve them, 

then showing that the computer technology was also there. We realized that we 

could compute our way to Mach 25, with high confidence. That was a high point 

of the presentations.”1

Whitehead’s point of departure lay in the fundamental equations of fluid flow: 

the Navier-Stokes equations, named for the nineteenth-century physicists ClaudeLouis-Marie Navier and Sir George Stokes. They form a set of nonlinear partial 

differential equations that contain 60 partial derivative terms. Their physical content is simple, comprising 

the basic laws of conservation of mass, momentum, 

and energy, along with an 

equation of state. Yet their 

solutions, when available, 

cover the entire realm of 

fluid mechanics.2

An example of an 

important development, 

contemporaneous with 

Whitehead’s presentation, 

was a 1985 treatment of 

flow over a complete X24C vehicle at Mach 5.95. 

The authors, Joseph Shang 

and S. J. Scheer, were at 

the Air Force’s Wright 

Aeronautical Laboratories. 

They used a Cray X-MP 

supercomputer and gave 

lift and drag coefficients:3

Development of CFD prior to NASP. In addition to vast improvement in computers, there also was similar advance in the 

performance of codes. (NASA)
232

Facing the Heat Barrier: A History of Hypersonics

233

Why NASP Fell Short

NASP-era analysts 

fell back on the “eN

method,” which gave a 

greatly simplified summary of the pertinent 

physics but still gave 

results that were often 

viewed as useful. It used 

the Navier-Stokes equations to solve for the 

overall flow in the laminary boundary layer, 

upstream of transition. 

This method then introduced new and simple 

equations derived from 

the original NavierStokes. These were 

linear and traced the 

growth of a small disturbance as one followed the flow downstream. When it had 

grown by a factor of 22,000—e10, with N = 10—the analyst accepted that transition 

to turbulence had occurred.7

One can obtain a solution in this fashion, but transition results from local roughnesses along a surface, and these can lead to results that vary dramatically. Thus, 

the repeated re-entries of the space shuttle, during dozens of missions, might have 

given numerous nearly identical data sets. In fact, transition has occurred at Mach 

numbers from 6 to 19! A 1990 summary presented data from wind tunnels, ballistic 

ranges, and tests of re-entry vehicles in free flight. There was a spread of as much as 

30 to one in the measured locations of transition, with the free-flight data showing 

transition positions that typically were five times farther back from a nose or leading 

edge than positions observed using other methods. At Mach 7, observed locations 

covered a range of 20 to one.8

One may ask whether transition can be predicted accurately even in principle 

because it involves minute surface roughnesses whose details are not known a priori 

and may even change in the course of a re-entry. More broadly, the state of transition was summarized in a 1987 review of problems in NASP hypersonics that was 

written by three NASA leaders in CFD:

Almost nothing is known about the effects of heat transfer, pressure 

gradient, three-dimensionality, chemical reactions, shock waves, and other 

Experimentally determined locations of the onset of transition to 

turbulent flow. The strong scatter of the data points defeats attempts to find a predictive rule. (NASA)

CFD held particular promise because it had the potential of overcoming the 

limitations of available facilities. These limits remained in place all through the 

NASP era. A 1993 review found “adequate” test capability only for classical aerodynamic experiments in a perfect gas, namely helium, which could support such work 

to Mach 20. Between Mach 13 and 17 there was “limited” ability to conduct tests 

that exhibited real-gas effects, such as molecular excitation and dissociation. Still, 

available facilities were too small to capture effects associated with vehicle size, such 

as determining the location of boundary-layer transition to turbulence.

For scramjet studies, the situation was even worse. There was “limited” ability to test combustors out to Mach 7, but at higher Mach the capabilities were 

“inadequate.” Shock tunnels supported studies of flows in rarefied air from Mach 

16 upward, but the whole of the nation’s capacity for such tests was “inadequate.” 

Some facilities existed that could study complete engines, either by themselves or 

in airframe-integrated configurations, but again the whole of this capability was 

“inadequate.”5

Yet it was an exaggeration in 1984, and remains one to this day, to propose that 

CFD could remedy these deficiencies by computing one’s way to orbital speeds 

“with high confidence.” Experience has shown that CFD falls short in two areas: 

prediction of transition to turbulence, which sharply increases drag due to skin friction, and in the simulation of turbulence itself.

For NASP, it was vital not only to predict transition but to understand the properties of turbulence after it appeared. One could see this by noting that hypersonic 

propulsion differs substantially from propulsion of supersonic aircraft. In the latter, 

the art of engine design allows engineers to ensure that there is enough margin of 

thrust over drag to permit the vehicle to accelerate. A typical concept for a Mach 3 

supersonic airliner, for instance, calls for gross thrust from the engines of 123,000 

pounds, with ram drag at the inlets of 54,500. The difference, nearly 80,000 pounds 

of thrust, is available to overcome skin-friction drag during cruise, or to accelerate.

At Mach 6, a representative hypersonic-transport design shows gross thrust of 

330,000 pounds and ram drag of 220,000. Again there is plenty of margin for what, 

after all, is to be a cruise vehicle. But in hypersonic cruise at Mach 12, the numbers 

typically are 2.1 million pounds for gross thrust—and 1.95 million for ram drag! 

Here the margin comes to only 150,000 pounds of thrust, which is narrow indeed. 

It could vanish if skin-friction drag proves to be higher than estimated, perhaps 

because of a poor forecast of the location of transition. The margin also could vanish 

if the thrust is low, due to the use of optimistic turbulence models.6

Any high-Mach scramjet-powered craft must not only cruise but accelerate. In 

turn, the thrust driving this acceleration appears as a small difference between two 

quantities: total drag and net thrust, the latter being net of losses within the engines. 

Accordingly, valid predictions concerning transition and turbulence are matters of 

the first importance.
234

Facing the Heat Barrier: A History of Hypersonics

235

Why NASP Fell Short

information, end the regress, and give a set of equations for turbulent flow in which 

the number of equations again would match the number of unknowns.12

The standard means to address this issue has been a turbulence model. This takes 

the form of one or more auxiliary equations, either algebraic or partial-differential, 

which are solved simultaneously with the Navier-Stokes equations in Reynolds-averaged form. In turn, the turbulence model attempts to derive one or more quantities 

that describe the turbulence and to do so in a way that ends the regress.

Viscosity, a physical property of every liquid and gas, provides a widely used 

point of departure. It arises at the molecular level, and the physics of its origin is well 

understood. In a turbulent flow, one may speak of an “eddy viscosity” that arises 

by analogy, with the turbulent eddies playing the role of molecules. This quantity 

describes how rapidly an ink drop will mix into a stream—or a parcel of hydrogen 

into the turbulent flow of a scramjet combustor.13

Like the eN method in studies of transition, eddy viscosity presents a view of turbulence that is useful and can often be made to work, at least in well-studied cases. 

The widely used Baldwin-Lomax model is of this type, and it uses constants derived 

from experiment. Antony Jameson of Princeton University, a leading writer of flow 

codes, described it in 1990 as “the most popular turbulence model in the industry, 

primarily because it’s easy to program.”14

This approach indeed gives a set of equations that are solvable and avoid the 

regress, but the analyst pays a price: Eddy viscosity lacks standing as a concept 

supported by fundamental physics. Peter Bradshaw of Stanford University virtually rejects it out of hand, declaring, “Eddy viscosity does not even deserve to be 

described as a ‘theory’ of turbulence!” He adds more broadly, “The present state is 

that even the most sophisticated turbulence models are based on brutal simplification of the N-S equations and hence cannot be relied on to predict a large range of 

flows with a fixed set of empirical coefficients.”15

Other specialists gave similar comments throughout the NASP era. Thomas 

Coakley of NASA-Ames wrote in 1983 that “turbulence models that are now used 

for complex, compressible flows are not well advanced, being essentially the same 

models that were developed for incompressible attached boundary layers and shear 

flows. As a consequence, when applied to compressible flows they yield results that 

vary widely in terms of their agreement with experimental measurements.”16

A detailed critique of existing models, given in 1985 by Budugur Lakshminarayana of Pennsylvania State University, gave pointed comments on algebraic models, 

which included Baldwin-Lomax. This approach “provides poor predictions” for 

flows with “memory effects,” in which the physical character of the turbulence does 

not respond instantly to a change in flow conditions but continues to show the influence of upstream effects. Such a turbulence model “is not suitable for flows with 

curvature, rotation, and separation. The model is of little value in three-dimensional 

complex flows and in situations where turbulence transport effects are important.”

influences on hypersonic transition. This is caused by the difficulty of 

conducting meaningful hypersonic transition experiments in noisy groundbased facilities and the expense and difficulty of carrying out detailed and 

carefully controlled experiments in flight where it is quiet. Without an 

adequate, detailed database, development of effective transition models will 

be impossible.9

Matters did not improve in subsequent years. In 1990 Mujeeb Malik, a leader 

in studies of transition, noted “the long-held view that conventional, noisy ground 

facilities are simply not suitable for simulation of flight transition behavior.” A subsequent critique added that “we easily recognize that there is today no reasonably 

reliable predictive capability for engineering applications” and commented that “the 

reader...is left with some feeling of helplessness and discouragement.”10 A contemporary review from the Defense Science Board pulled no punches: “Boundary layer 

transition...cannot be validated in existing ground test facilities.”11

There was more. If transition could not be predicted, it also was not generally 

possible to obtain a valid simulation, from first principles, of a flow that was known 

to be turbulent. The Navier-Stokes equations carried the physics of turbulence at 

all scales. The problem was that in flows of practical interest, the largest turbulent 

eddies were up to 100,000 times bigger than the smallest ones of concern. This 

meant that complete numerical simulations were out of the question.

Late in the nineteenth century the physicist Osborne Reynolds tried to bypass 

this difficulty by rederiving these equations in averaged form. He considered the 

flow velocity at any point as comprising two elements: a steady-flow part and a 

turbulent part that contained all the motion due to the eddies. Using the NavierStokes equations, he obtained equations for averaged quantities, with these quantities being based on the turbulent velocities.

He found, though, that the new equations introduced additional unknowns. 

Other investigators, pursuing this approach, succeeded in deriving additional 

equations for these extra unknowns—only to find that these introduced still more 

unknowns. Reynolds’s averaging procedure thus led to an infinite regress, in which 

at every stage there were more unknown variables describing the turbulence than 

there were equations with which to solve for them. This contrasted with the NavierStokes equations themselves, which in principle could be solved because the number 

of these equations and the number of their variables was equal.

This infinite regress demonstrated that it was not sufficient to work from the 

Navier-Stokes equations alone—something more was needed. This situation arose 

because the averaging process did not preserve the complete physical content of 

the Navier-Stokes formulation. Information had been lost in the averaging. The 

problem of turbulence thus called for additional physics that could replace the lost 
236

Facing the Heat Barrier: A History of Hypersonics

237

Why NASP Fell Short

In this scenario, two flows that have different velocities proceed along opposite 

sides of a thin plate, which terminates within a channel. The mixing layer then 

forms and grows at the interface between these streams. In Roshko’s words, “a onepercent periodic disturbance in the free stream completely changes the mixing layer 

growth.” This has been seen in experiments and in highly detailed solutions of the 

Navier-Stokes equations that solve the complete equations using a very fine grid. It 

has not been seen in solutions of Reynolds-averaged equations that use turbulence 

models.22

And if simple flows of this type bring such difficulties, what can be said of hypersonics? Even in the free stream that lies at some distance from a vehicle, one finds 

strong aerodynamic heating along with shock waves and the dissociation, recombination, and chemical reaction of air molecules. Flow along the aircraft surface adds 

a viscous boundary layer that undergoes shock impingement, while flow within the 

engine adds the mixing and combustion of fuel.

As William Dannevik of Lawrence Livermore National Laboratory describes 

it, “There’s a fully nonlinear interaction among several fields: an entropy field, an 

acoustic field, a vortical field.” By contrast, in low-speed aerodynamics, “you can 

often reduce it down to one field interacting with itself.” Hypersonic turbulence also 

brings several channels for the flow and exchange of energy: internal energy, density, 

and vorticity. The experimental difficulties can be correspondingly severe.23

Roshko sees some similarity between turbulence modeling and the astronomy of 

Ptolemy, who flourished when the Roman Empire was at its height. Ptolemy represented the motions of the planets using epicycles and deferents in a purely empirical 

fashion and with no basis in physical theory. “Many of us have used that example,” 

Roshko declares. “It’s a good analogy. People were able to continually keep on fixing 

up their epicyclic theory, to keep on accounting for new observations, and they were 

completely wrong in knowing what was going on. I don’t think we’re that badly off, 

but it’s illustrative of another thing that bothers some people. Every time some new 

thing comes around, you’ve got to scurry and try to figure out how you’re going to 

incorporate it.”24

A 1987 review concluded, “In general, the state of turbulence modeling for 

supersonic, and by extension, hypersonic, flows involving complex physics is poor.” 

Five years later, late in the NASP era, little had changed, for a Defense Science 

Board program review pointed to scramjet development as the single most important issue that lay beyond the state of the art.25

Within NASP, these difficulties meant that there was no prospect of computing 

one’s way in orbit, or of using CFD to make valid forecasts of high-Mach engine 

performance. In turn, these deficiencies forced the program to fall back on its test 

facilities, which had their own limitations.

“Two-equation models,” which used two partial differential equations to give 

more detail, had their own faults. In the view of Lakshminarayana, they “fail to capture many of the features associated with complex flows.” This class of models “fails 

for flows with rotation, curvature, strong swirling flows, three-dimensional flows, 

shock-induced separation, etc.”17

Rather than work with eddy viscosity, some investigators used “Reynolds stress” 

models. Reynolds stresses were not true stresses, which contributed to drag. Rather, 

they were terms that appeared in the Reynolds-averaged Navier-Stokes equations 

alongside other terms that indeed represented stress. Models of this type offered 

greater physical realism, but again this came at the price of severe computational 

difficulty.18

A group at NASA-Langley, headed by Thomas Gatski, offered words of caution 

in 1990: “...even in the low-speed incompressible regime, it has not been possible 

to construct a turbulence closure model which can be applied over a wide class of 

flows.... In general, Reynolds stress closure models have not been very successful in 

handling the effects of rotation or three-dimensionality even in the incompressible 

regime; therefore, it is not likely that these effects can be treated successfully in the 

compressible regime with existing models.”19

Anatol Roshko of Caltech, widely viewed as a dean of aeronautics, has his own 

view: “History proves that each time you get into a new area, the existing models 

are found to be inadequate.” Such inadequacies have been seen even in simple flows, 

such as flow over a flat plate. The resulting skin friction is known to an accuracy of 

around one percent. Yet values calculated from turbulence models can be in error by 

up to 10 percent. “You can always take one of these models and fix it so it gives the 

right answer for a particular case,” says Bradshaw. “Most of us choose the flat plate. 

So if you can’t get the flat plate right, your case is indeed piteous.”20

Another simple case is flow within a channel that suddenly widens. Downstream 

of the point of widening, the flow shows a zone of strongly whirling circulation. It 

narrows until the main flow reattaches, flowing in a single zone all the way to the 

now wider wall. Can one predict the location of this reattachment point? “This is 

a very severe test,” says John Lumley of Cornell University. “Most of the simple 

models have trouble getting reattachment within a factor of two.” So-called “k-epsilon models,” he says, are off by that much. Even so, NASA’s Tom Coakley describes 

them as “the most popular two-equation model,” whereas Princeton University’s 

Jameson speaks of them as “probably the best engineering choice around” for such 

problems as...flow within a channel.21

Turbulence models have a strongly empirical character and therefore often fail 

to predict the existence of new physics within a flow. This has been seen to cause 

difficulties even in the elementary case of steady flow past a cylinder at rest, a case 

so simple that it is presented in undergraduate courses. Nor do turbulence models 

cope with another feature of some flows: their strong sensitivity to slight changes in 

conditions. A simple example is the growth of a mixing layer.
238

Facing the Heat Barrier: A History of Hypersonics

239

Why NASP Fell Short

For takeoff from a runway, the X-30 was to use a Low-Speed System (LSS). It 

comprised two principal elements: the Special System, an ejector ramjet; and the 

Low Speed Oxidizer System, which used LACE.28 The two were highly synergistic. 

The ejector used a rocket, which might have been suitable for the final ascent to 

orbit, with ejector action increasing its thrust during takeoff and acceleration. By 

giving an exhaust velocity that was closer to the vehicle velocity, the ejector also 

increased the fuel economy.

The LACE faced the standard problem of requiring far more hydrogen than 

could be burned in the air it liquefied. The ejector accomplished some derichening by providing a substantial flow of entrained air that burned some of the excess. 

Additional hydrogen, warmed in the LACE heat exchanger, went into the fuel 

tanks, which were full of slush hydrogen. By melting the slush into conventional 

liquid hydrogen (LH2

), some LACE coolant was recycled to stretch the vehicle’s fuel 

supply.29

There was good news in at least one area of LACE research: deicing. LACE 

systems have long been notorious for their tendency to clog with frozen moisture 

within the air that they liquefy. “The largest LACE ever built made around half a 

pound per second of liquid air,” Paul Czysz of McDonnell Douglas stated in 1986. 

“It froze up at six percent relative humidity in the Arizona desert, in 38 seconds.” 

Investigators went on to invent more than a dozen methods for water alleviation. 

The most feasible approach called for injecting antifreeze into the system, to enable 

the moisture to condense out as liquid water without freezing. A rotary separator 

eliminated the water, with the dehumidified air being so cold as to contain very little 

residual water vapor.30

The NASP program was not run by shrinking violets, and its managers stated 

that its LACE was not merely to operate during hot days in the desert near Phoenix. 

It was to function even on rainy days, for the X-30 was to be capable of flight from 

anywhere in the world. At NASA-Lewis, James Van Fossen built a water-alleviation 

system that used ethylene glycol as the antifreeze, spraying it directly onto the cold 

tubes of a heat exchanger. Water, condensing on those tubes, dissolved some of 

the glycol and remained liquid as it swept downstream with the flow. He reported 

that this arrangement protected the system against freezing at temperatures as low 

as −55oF, with the moisture content of the chilled air being reduced to 0.00018 

pounds in each pound of this air. This represented removal of at least 99 percent of 

the humidity initially present in the airflow.31

Pratt & Whitney conducted tests of a LACE precooler that used this arrangement. A company propulsion manager, Walt Lambdin, addressed a NASP technical 

review meeting in 1991 and reported that it completely eliminated problems of 

reduced performance of the precooler due to formation of ice. With this, the problem of ice in a LACE system appeared amenable to control.32

Propulsion

In the spring of 1992 the NASP Joint Program Office presented a final engine 

design called the E22A. It had a length of 60 feet and included an inlet ramp, 

cowled inlet, combustor, and nozzle. An isolator, located between the inlet and 

combustor, sought to prevent unstarts by processing flow from the inlet through a 

series of oblique shocks, which increased the backpressure from the combustor.

Program officials then constructed two accurately scaled test models. The Subscale Parametric Engine (SXPE) was built to one-eighth scale and had a length of 

eight feet. It was tested from April 1993 to March 1994. The Concept Demonstrator Engine (CDE), which followed, was built to a scale of 30 percent. Its length 

topped 16 feet, and it was described as “the largest airframe-integrated scramjet 

engine ever tested.”26

In working with the SXPE, researchers had an important goal in achieving combustion of hydrogen within its limited length. To promote rapid ignition, the engine 

used a continuous flow of a silane-hydrogen mixture as a pilot, with the silane igniting spontaneously on exposure to air. In addition, to promote mixing, the model 

incorporated an accurate replication of the spacing between the fuel-injecting struts 

and ramps, with this spacing being preserved at the model’s one-eighth scale. The 

combustor length required to achieve the desired level of mixing then scaled in this 

fashion as well.

The larger CDE was tested within the Eight-Foot High-Temperature Tunnel, 

which was Langley’s biggest hypersonic facility. The tests mapped the flowfield 

entering the engine, determined the performance of the inlet, and explored the 

potential performance of the design. Investigators varied the fuel flow rate, using the 

combustors to vary its distribution within the engine.

Boundary-layer effects are important in scramjets, and the tests might have replicated the boundary layers of a full-scale engine by operating at correspondingly 

higher flow densities. For the CDE, at 30 percent scale, the appropriate density 

would have been 1/0.3 or 3.3 times that of the atmospheric density at flight altitude. For the SXPE, at one-eighth scale, the test density would have shown an eightfold increase over atmospheric. However, the SXPE used an arc-heated test facility 

that was limited in the power that drove its arc, and it provided its engine with air 

at only one-fiftieth of that density. The High Temperature Tunnel faced limits on its 

flow rate and delivered its test gas at only one-sixth of the appropriate density.

Engineers sought to compensate by using analytical methods to determine the 

drag in a full-scale engine. Still, this inability to replicate boundary-layer effects 

meant that the wind-tunnel tests gave poor simulations of internal drag within the 

test engines. This could have led to erroneous estimates of true thrust, net of drag. 

In turn, this showed that even when working with large test models and with test 

facilities of impressive size, true simulations of the boundary layer were ruled out 

from the start.27
240

Facing the Heat Barrier: A History of Hypersonics

241

Why NASP Fell Short

Even so, this remained beyond the state of the art for NASP, a quarter-century later. Weight estimates for the X-30 LACE heat exchanger were based on the 

assumed use of 3-mil Weldalite tubing, but a 1992 Lockheed review stated, “At 

present, only small quantities of suitable, leak free, 3-mil tubing have been fabricated.” The plans of that year called for construction of test prototypes using 6-mil 

Weldalite tubing, for which “suppliers have been able to provide significant quantities.” Still, a doubled thickness of the tubing wall was not the way to achieve low 

weight.38

Other weight problems arose in seeking to apply an ingenious technique for 

derichening the product stream by increasing the heat capacity of the LH2

 coolant. 

Molecular hydrogen, H2

, has two atoms in its molecule and exists in two forms: 

para and ortho, which differ in the orientation of the spins of their electrons. The 

ortho form has parallel spin vectors, while the para form has spin vectors that are 

oppositely aligned. The ortho molecule amounts to a higher-energy form and loses 

energy as heat when it transforms into the para state. The reaction therefore is exothermic.

The two forms exist in different equilibrium concentrations, depending on the 

temperature of the bulk hydrogen. At room temperature the gas is about 25 percent 

para and 75 percent ortho. When liquefied, the equilibrium state is 100 percent 

para. Hence it is not feasible to prepare LH2

 simply by liquefying the room-temperature gas. The large component of ortho will relax to para over several hours, 

producing heat and causing the liquid to boil away. The gas thus must be exposed 

to a catalyst to convert it to the para form before it is liquefied.

These aspects of fundamental chemistry also open the door to a molecular shift 

that is endothermic and that absorbs heat. One achieves this again by using a catalyst to convert the LH2

 from para to ortho. This reaction requires heat, which is 

obtained from the liquefying airflow within the LACE. As a consequence, the air 

chills more readily when using a given flow of hydrogen refrigerant. This effect is 

sufficiently strong to increase the heat-sink capacity of the hydrogen by as much as 

25 percent.39

This concept also dates to the 1960s. Experiments showed that ruthenium metal 

deposited on aluminum oxide provided a suitable catalyst. For 90 percent para-toortho conversion, the LACE required a “beta,” a ratio of mass to flow rate, of five 

to seven pounds of this material for each pound per second of hydrogen flow. Data 

published in 1988 showed that a beta of five pounds could achieve 85 percent conversion, with this value showing improvement during 1992. However, X-30 weight 

estimates assumed a beta of two pounds, and this performance remained out of 

reach.40

During takeoff, the X-30 was to be capable of operating from existing runways 

and of becoming airborne at speeds similar to those of existing aircraft. The lowIt was also possible to gain insight into the LACE state of the art by considering 

contemporary work that was under way in Japan. The point of departure in that 

country was the H-2 launch vehicle, which first flew to orbit in February 1994. It 

was a two-stage expendable rocket, with a liquid-fueled core flanked by two solid 

boosters. LACE was pertinent because a long-range plan called for upgrades that 

could replace the solid strap-ons with new versions using LACE engines.33

Mitsubishi Heavy Industries was developing the H-2’s second-stage engine, designated LE-5. It burned hydrogen and oxygen to produce 22,000 pounds of thrust. 

As an initial step toward LACE, this company built heat exchangers to liquefy air 

for this engine. In tests conducted during 1987 and 1988, the Mitsubishi heat 

exchanger demonstrated liquefaction of more than three pounds of air for every 

pound of LH2

. This was close to four to one, the theoretical limit based on the thermal properties of LH2

 and of air. Still, it takes 34.6 pounds of air to burn a pound 

of hydrogen, and an all-LACE LE-5 was to run so fuel-rich that its thrust was to be 

only 6,000 pounds.

But the Mitsubishi group found their own path to prevention of ice buildup. 

They used a freeze-thaw process, melting ice by switching periodically to the use of 

ambient air within the cooler after its tubes had become clogged with ice from LH2

. 

The design also provided spaces between the tubes and allowed a high-speed airflow 

to blow ice from them.34

LACE nevertheless remained controversial, and even with the moisture problem 

solved, there remained the problem of weight. Czysz noted that an engine with 

100,000 pounds of thrust would need 600 pounds per second of liquid air: “The 

largest liquid-air plant in the world today is the AiResearch plant in Los Angeles, at 

150 pounds per second. It covers seven acres. It contains 288,000 tubes welded to 

headers and 59 miles of 3/32-inch tubing.”35

Still, no law required the use of so much tubing, and advocates of LACE have 

long been inventive. A 1963 Marquardt concept called for an engine with 10,000 

pounds of thrust, which might have been further increased by using an ejector. This 

appeared feasible because LACE used LH2

 as the refrigerant. This gave far greater 

effectiveness than the AiResearch plant, which produced its refrigerant on the spot 

by chilling air through successive stages.36

For LACE heat exchangers, thin-walled tubing was essential. The Japanese 

model, which was sized to accommodate the liquid-hydrogen flow rate of the LE5, used 5,400 tubes and weighed 304 pounds, which is certainly noticeable when 

the engine is to put out no more than 6,000 pounds of thrust. During the mid1960s investigators at Marquardt and AiResearch fabricated tubes with wall thicknesses as low as 0.001 inch, or one mil. Such tubes had not been used in any heat 

exchanger subassemblies, but 2-mil tubes of stainless steel had been crafted into a 

heat exchanger core module with a length of 18 inches.37
242

Facing the Heat Barrier: A History of Hypersonics

243

Why NASP Fell Short

“buzz” or unwanted vibration of the inlet structure. Even with no primary flow, 

the inlet failed to start. The main burner never achieved thermal choking, where 

the flow rate would rise to the maximum permitted by heat from burning fuel. 

Ingestion of the boundary layer significantly degraded engine performance. Thrust 

measurements were described as “no good” due to nonuniform thermal expansion 

across a break between zones of measurement. As a contrast to this litany of woe, 

operation of the primary gave a welcome improvement in the isolation of the inlet 

from the combustor.

Also at GASL, again during 1987, an ejector from Boeing underwent static test. 

It used a markedly different configuration that featured an axisymmetric duct and 

a fuel-air mixer. The primary flow was fuel-rich, with temperatures and pressures 

similar to those of NASA-Lewis. On the whole, the results of the Boeing tests were 

encouraging. Combustion efficiencies appeared to exceed 95 percent, while measured values of thrust, entrained airflow, and pressures were consistent with company predictions. However, the mixer performance was no more than marginal, and 

its length merited an increase for better performance.45

In 1989 Pratt & Whitney emerged as a major player, beginning with a subscale 

ejector that used a flow of helium as the primary. It underwent tests at company 

facilities within the United Technologies Research Center. These tests addressed the 

basic issue of attempting to increase the entrainment of secondary flow, for which 

non-combustible helium was useful. Then, between 1990 and 1992, Pratt built 

three versions of its Low Speed Component Integration Rig (LSCIR), testing them 

all within facilities of Marquardt.

LSCIR-1 used a design that included a half-scale X-30 flowpath. It included an 

inlet, front and main combustors, and nozzle, with the inlet cowl featuring fixed 

geometry. The tests operated using ambient air as well as heated air, with and without fuel in the main combustor, while the engine operated as a pure ramjet for 

several runs. Thermal choking was achieved, with measured combustion efficiencies 

lying within 2 percent of values suitable for the X-30. But the inlet was unstarted for 

nearly all the runs, which showed that it needed variable geometry. This refinement 

was added to LSCIR-2, which was put through its paces in July 1991, at Mach 2.7. 

The test sequence would have lasted longer but was terminated prematurely due to a 

burnthrough of the front combustor, which had been operating at 1,740oF. Thrust 

measurements showed only limited accuracy due to flow separation in the nozzle.

LSCIR-3 followed within months. The front combustor was rebuilt with a larger 

throat area to accommodate increased flow and received a new ignition system that 

used silane. This gas ignited spontaneously on contact with air. In tests, leaks developed between the main combustor, which was actively cooled, and the uncooled 

nozzle. A redesigned seal eliminated the leakage. The work also validated a method 

for calculating heat flux to the wall due to impingement of flow from primaries.

speed system, along with its accompanying LACE and ejector systems, therefore 

needed substantial levels of thrust. The ejector, again, called for a rocket exhaust to 

serve as a primary flow within a duct, entraining an airstream as the secondary flow. 

Ejectors gave good performance across a broad range of flight speeds, showing an 

effectiveness that increased with Mach. In the SR-71 at Mach 2.2, they accounted 

for 14 percent of the thrust in afterburner; at Mach 3.2 this was 28.4 percent. 

Nor did the SR-71 ejectors burn fuel. They functioned entirely as aerodynamic 

devices.41

It was easy to argue during the 1980s that their usefulness might be increased 

still further. The most important unclassified data had been published during the 

1950s. A good engine needed a high pressure increase, but during the mid-1960s 

studies at Marquardt recommended a pressure rise by a factor of only about 1.5, 

when turbojets were showing increases that were an order of magnitude higher.42

The best theoretical treatment of ejector action dated to 1974. Its author, NASA’s 

B. H. Anderson, also wrote a computer program called REJECT that predicted 

the performance of supersonic ejectors. However, he had done this in 1974, long 

before the tools of CFD were in hand. A 1989 review noted that since then “little 

attention has been directed toward a better understanding of the details of the flow 

mechanism and behavior.”43

Within the NASP program, then, the ejector ramjet stood as a classic example 

of a problem that was well suited to new research. Ejectors were known to have 

good effectiveness, which might be increased still further and which stood as a good 

topic for current research techniques. CFD offered an obvious approach, and NASP 

activities supplemented computational work with an extensive program of experiment.44

The effort began at GASL, where Tony duPont’s ejector ramjet went on a static 

test stand during 1985 and impressed General Skantze. DuPont’s engine design 

soon took the title of the Government Baseline Engine and remained a topic of 

active experimentation during 1986 and 1987. Some work went forward at NASALangley, where the Combustion Heated Scramjet Test Facility exercised ejectors 

over the range of Mach 1.2 to 3.5. NASA-Lewis hosted further tests, at Mach 0.06 

and from Mach 2 to 3.5 within its 10 by 10 foot Supersonic Wind Tunnel.

The Lewis engine was built to accommodate growth of boundary layers and 

placed a 17-degree wedge ramp upstream of the inlet. Three flowpaths were mounted 

side by side, but only the center duct was fueled; the others were “dummies” that 

gave data on unfueled operation for comparison. The primary flow had a pressure 

of 1,000 pounds per square inch and a temperature of 1,340oF, which simulated a 

fuel-rich rocket exhaust. The experiments studied the impact of fuel-to-air ratio on 

performance, although the emphasis was on development of controls.

Even so, the performance left much to be desired. Values of fuel-to-air ratio 

greater than 0.52, with unity representing complete combustion, at times brought 
244

Facing the Heat Barrier: A History of Hypersonics

245

Why NASP Fell Short

the X-30 was defined by aerodynamic heating and by the separate issue of flutter.48

A single concern dominated issues of structural design: The vehicle was to fly 

as low as possible in the atmosphere during ascent to orbit. Re-entry called for 

flight at higher altitudes, and the loads during ascent therefore were higher than 

those of re-entry. Ascent at lower altitude—200,000 feet, for instance, rather than 

250,000—increased the drag on the X-30. But it also increased the thrust, giving a 

greater margin between thrust and drag that led to increased acceleration. Considerations of ascent, not re-entry, therefore shaped the selection of temperature-resistant 

materials.

Yet the aircraft could not fly too low, or it would face limits set by aerodynamic 

flutter. This resulted from forces on the vehicle that were not steady but oscillated, 

at frequencies of oscillation that changed as the vehicle accelerated and lost weight. 

The wings tended to vibrate at characteristic frequencies, as when bent upward 

and released to flex up and down. If the frequency of an aerodynamic oscillation 

matched that at which the wings were prone to flex, the aerodynamic forces could 

tear the wings off. Stiffness in materials, not strength, was what resisted flutter, and 

the vehicle was to fly a “flutter-limited trajectory,” staying high enough to avoid the 

problem.

The mechanical 

properties of metals 

depend on their finegrained structure. An 

ingot of metal consists 

of a mass of interlaced 

grains or crystals, and 

small grains give higher 

strength. Quenching, 

plunging hot metal 

into water, yields small 

grains but often makes 

the metal brittle or 

hard to form. Alloying 

a metal, as by adding 

small quantities of 

carbon to make steel, 

is another traditional 

practice. However, 

some additives refuse to dissolve or separate out from the parent metal as it cools.

To overcome such restrictions, techniques of powder metallurgy were in the forefront. These methods gave direct control of the microstructure of metals by forming 

Other results were less successful. Ignition proceeded well enough using pure 

silane, but a mix of silane and hydrogen failed as an ignitant. Problems continued to 

recur due to inlet unstarts and nozzle flow separation. The system produced 10,000 

pounds of thrust at Mach 0.8 and 47,000 pounds at Mach 2.7, but this performance still was rated as low.

Within the overall LSS program, a Modified Government Baseline Engine went 

under test at NASA-Lewis during 1990, at Mach 3.5. The system now included 

hydraulically-operated cowl and nozzle flaps that provided variable geometry, along 

with an isolator with flow channels that amounted to a bypass around the combustor. This helped to prevent inlet unstarts.

Once more the emphasis was on development of controls, with many tests operating the system as a pure ramjet. Only limited data were taken with the primaries 

on. Ingestion of the boundary layer gave significant degradation in engine performance, but in other respects most of the work went well. The ramjet operations 

were successful. The use of variable geometry provided reliable starting of the inlet, 

while operation in the ejector mode, with primaries on, again improved the inlet 

isolation by diminishing the effect of disturbances propagating upstream from the 

combustor.46

Despite these achievements, a 1993 review at Rocketdyne gave a blunt conclusion: “The demonstrated performance of the X-30 special system is lower than the 

performance level used in the cycle deck...the performance shortfall is primarily 

associated with restrictions on the amount of secondary flow.” (Secondary flow is 

entrained by the ejector’s main flow.) The experimental program had taught much 

concerning the prevention of inlet unstarts and the enhancement of inlet-combustor isolation, but the main goal—enhanced performance of the ejector ramjet—still 

lay out of reach.

Simple enlargement of a basic design offered little promise; Pratt & Whitney 

had tried that, in LSCIR-3, and had found that this brought inlet flow separation 

along with reduced inlet efficiency. Then in March 1993, further work on the LSS 

was canceled due to budget cuts. NASP program managers took the view that they 

could accelerate an X-30 using rockets for takeoff, as an interim measure, with the 

LSS being added at a later date. Thus, although the LSS was initially the critical 

item in duPont’s design, in time it was put on hold and held off for another day.47

Materials

No aircraft has ever cruised at Mach 5, and an important reason involves structures and materials. “If I cruise in the atmosphere for two hours,” says 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.” The thermal environment of 

Ascent trajectory of an airbreather. (NASA)
246

Facing the Heat Barrier: A History of Hypersonics

247

Why NASP Fell Short

also could 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 mil when metal matrix composites were fabricated.50

Such titanium-matrix composites were used in representative X-30 structures. 

The Non-Integral Fuselage Tank Article (NIFTA) represented a section of X-30 

fuselage at one-fourth scale. It was oblong in shape, eight feet long and measuring 

four by seven feet in cross section, and it contained a splice. Its skin thickness was 

0.040 inches, about the same as for the X-30. It held an insulated tank that could 

hold either liquid nitrogen or LH2

 in tests, which stood as a substantial engineering 

item in its own right.

The tank had a capacity of 940 gallons and was fabricated of graphite-epoxy 

composite. No liner protected the tankage on the inside, for graphite-epoxy was 

impervious to damage by LH2

. However, the exterior was insulated with two halfinch thicknesses of Q-felt, a quartz-fiber batting with density of only 3.5 pounds per 

cubic foot. A thin layer of Astroquartz high-temperature cloth covered the Q-felt. 

This insulation filled space between the tank wall and the surrounding wall of the 

main structure, with both this space and the Q-felt being purged with helium.51

The test sequence for NIFTA duplicated the most severe temperatures and 

stresses of an ascent to orbit. These stresses began on the ground, with the vehicle 

being heavy with fuel and subject to a substantial bending load. There was also a 

Comparison of some matrix alloys. (NASA)

them from powder, with the grains of powder sintering or welding together by being 

pressed in a mold at high temperature. A manufacturer could control the grain 

size independently of any heat-treating process. Powder metallurgy also overcame 

restrictions on alloying by mixing in the desired additives as powdered ingredients.

Several techniques existed to produce the powders. Grinding a metal slab to sawdust was the simplest, yielding relatively coarse grains. “Splat-cooling” gave better 

control. It extruded molten metal onto the chilled rim of a rotating wheel, which 

cooled it instantly into a thin ribbon. This represented a quenching process that 

produced a fine-grained microstructure in the metal. The ribbon then was chemically treated with hydrogen, which made it brittle, so that it could be ground into a 

fine powder. Heating the powder then drove off the hydrogen.

The Plasma Rotating Electrode Process, developed by the firm of Nuclear Metals, 

showed particular promise. The parent metal was shaped into a cylinder that rotated 

at up to 30,000 revolutions per minute and served as an electrode. An electric arc 

melted the spinning metal, which threw off droplets within an atmosphere of cool 

inert helium. The droplets plummeted in temperature by thousands of degrees 

within milliseconds, and their microstructures were so fine as to approach an amorphous state. Their molecules did not form crystals, even tiny ones, but arranged 

themselves in formless patterns. This process, called “rapid solidification,” promised 

particular gains in high-temperature strength.

Standard titanium alloys, for instance, lost strength at temperatures above 700 to 

900oF. By using rapid solidification, McDonnell Douglas raised this limit to 1,100oF 

prior to 1986. Philip Parrish, the manager of powder metallurgy at DARPA, noted 

that his agency had spent some $30 million on rapid-solidification technology since 

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.”49

Nevertheless 1,100°F was not enough, for it appeared that the X-30 needed a 

material that was rated at 1,700oF. This stemmed from the fact that for several years, 

NASP design and trajectory studies indicated that a flight vehicle indeed would face 

such temperatures on its fuselage. But after 1990 the development of new baseline 

configurations led to an appreciation that the pertinent areas of the vehicle would 

face temperatures no higher than 1,500oF. At that temperature, advanced titanium 

alloys could serve in “metal matrix composites,” with thin-gauge metals being reinforced with fibers.

The new 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 titanium along with 15 percent molybdenum, 2.8 percent 

columbium, 3 percent aluminum, and 0.2 percent silicon. Resistance to oxidation 

proved to be its strong suit, with this alloy showing resistance that was two orders of 

magnitude greater than that of conventional aircraft titanium. Tests showed that it 
248

Facing the Heat Barrier: A History of Hypersonics

249

Why NASP Fell Short

composites. “We could get the cost down below a thousand dollars a pound if we 

had enough volume,” Bill Grant, a company manager, told Aerospace America. His 

colleague Jim Henshaw added, “We think SCS/titanium composites are fully developed for structural applications.”54

Such materials served to 1,500oF, but on the X-30 substantial areas were to withstand temperatures approaching 3,000oF, which is hotter than molten iron. If a 

steelworker were to plunge a hand into a ladle of this metal, the hand would explode 

from the sudden boiling of water in its tissues. In such areas, carbon-carbon was 

necessary. It had not been available for use in Dyna-Soar, but the Pentagon spent 

$200 million to fund its development between 1970 and 1985.55

Much of this supported the space shuttle, on which carbon-carbon protected 

such hot areas as the nose cap and wing leading edges. For the X-30, these areas 

expanded to cover the entire nose and much of the vehicle undersurface, along with 

the rudders and both the top and bottom surfaces of the wings. The X-30 was to 

execute 150 test flights, exposing its heat shield to prolonged thermal soaks while 

still in the atmosphere. This raised the problem of protection against oxidation.56

large shear load, with portions of the vehicle being pulled transversely in opposite 

directions. This happened because the landing gear pushed upward to support the 

entire weight of the craft, while the weight of the hydrogen tank pushed downward 

only a few feet away. Other major bending and shear loads arose during subsonic 

climbout, with the X-30 executing a pullup maneuver.

Significant stresses arose near Mach 6 and resulted from temperature differences 

across the thickness of the stiffened skin. Its outer temperature was to be 800oF, but 

the tops of the stiffeners, a few inches away, were to be 350oF. These stiffeners were 

spot-welded to the skin panels, which raised the issue of whether the welds would 

hold amid the different thermal expansions. Then between Mach 10 and 16, the 

vehicle was to reach peak temperatures of 1,300oF. The temperature differences 

between the top and bottom of the vehicle also would be at their maximum.

The tests combined both thermal and mechanical loads and were conducted 

within a vacuum chamber at Wyle Laboratories during 1991. Banks of quartz lamps 

applied up to 1.5 megawatts of heat, while jacks imposed bending or shear forces 

that reached 100 percent of the design limits. Most tests placed nonflammable 

liquid nitrogen in the tank for safety, but the last of them indeed used LH2

. With 

this supercold fuel at −423oF, the lamps raised the exterior temperature of NIFTA to 

the full 1,300oF, while the jacks applied the full bending load. A 1993 paper noted 

“100% successful completion of these tests,” including the one with LH2

 that had 

been particularly demanding.52

NIFTA, again, was at one-fourth scale. In a project that ran from 1991 through 

the summer of 1994, McDonnell Douglas engineers designed and fabricated the 

substantially larger Full Scale Assembly. Described as “the largest and most representative NASP fuselage structure built,” it took shape as a component measuring 

10 by 12 feet. It simulated a section of the upper mid-fuselage, just aft of the crew 

compartment.

A 1994 review declared that it “was developed to demonstrate manufacturing and 

assembly of a full scale fuselage panel incorporating all the essential structural details 

of a flight vehicle fuselage assembly.” Crafted in flightweight, it used individual 

panels of titanium-matrix composite that were as large as four by eight feet. These 

were stiffened with longitudinal members of the same material and were joined to 

circumferential frames and fittings of Ti-1100, a titanium alloy that used no fiber 

reinforcement. The complete assembly posed manufacturing challenges because the 

panels were of minimum thickness, having thinner gauges than had been used previously. The finished article was completed just as NASP was reaching its end, but it 

showed that the thin panels did not introduce significant problems.53

The firm of Textron manufactured the fibers, designated SCS-6 and -9, that 

reinforced the composites. As a final touch, in 1992 this company opened the 

world’s first manufacturing plant dedicated to the production of titanium-matrix 

Selection of NASP materials based on temperature. (General Accounting Office)
250

Facing the Heat Barrier: A History of Hypersonics

251

Why NASP Fell Short

Technologies gave the best results, suggesting 25 reuses for orbital missions of the 

X-30 and 50 reuses for the less-demanding missions of hypersonic cruise.60

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, and LTV’s box constituted the portion to the left of this line. Built 

at full scale, it represented a hot-structure wing proposed by General Dynamics. It 

measured five by eight 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 X30 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 994oF and the latter at 1,671oF. At 

that moment the load on the wing box corresponded to 34 percent of the Mach 2.2 

maximum. Farther along, the wing box was to reach its peak temperature, 1,925oF, 

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 key parts delaminated during their fabrication, seriously compromising the ability of the wing box to bear its specified load. Plans to impose the peak or 

Mach 2.2 load were abandoned, with the maximum planned load being reduced to 

the 34 percent 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 1,925oF. As the wing box approached 

this value, a problem arose due to the use of metal fasteners in its assembly. Some 

were made from coated columbium and were rated for 2,300oF, but most were of 

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 

Standard approaches called for mixing oxidation inhibitors into the carbon 

matrix and covering the surface with a coating of silicon carbide. However, there 

was a mismatch between the thermal expansions of the coating and the carboncarbon substrate, which led to cracks. An interlayer of glass-forming sealant, placed 

between them, produced an impervious barrier that softened at high temperatures to 

fill the cracks. But these glasses did not flow readily at temperatures below 1,500oF. 

This meant that air could penetrate the coating and reach the carbon through open 

cracks to cause loss by oxidation.57

The goal was to protect carbon-carbon against oxidation for all 150 of those test 

flights, or 250 hours. These missions included 75 to orbit and 75 in hypersonic 

cruise. The work proceeded initially by evaluating several dozen test samples that 

were provided by commercial vendors. Most of these materials proved to resist oxidation for only 10 to 20 hours, but one specimen from the firm of Hitco reached 

70 hours. Its surface had been grooved to promote adherence of the coating, and it 

gave hope that long operational life might be achieved.58

Complementing the study of vendors’ samples, researchers ordered new types 

of carbon-carbon and conducted additional tests. The most durable came from the 

firm of Rohr, with a coating by Science Applications International. It easily withstood 2,000oF for 200 hours and was still going strong at 2,500 oF when the tests 

stopped after 150 hours. This excellent performance stemmed from its use of large 

quantities of oxidation inhibitors, which promoted long life, and of multiple glass 

layers in the coating.

But even the best of these carbon-carbons showed far poorer performance when 

tested in arcjets at 2,500oF. The high-speed airflows forced oxygen into cracks and 

pores within the material, while promoting evaporation of the glass sealants. Powerful roars within the arcjets imposed acoustic loads that contributed to cracking, with 

other cracks arising from thermal shock as test specimens were suddenly plunged 

into a hot flow stream. The best results indicated lifetimes of less than two hours.

Fortunately, actual X-30 missions were to impose 2,500°F temperatures for 

only a few minutes during each launch and reentry. Even a single hour of lifetime 

therefore could permit panels of carbon-carbon to serve for a number of flights. A 

1992 review concluded that “maximum service temperatures should be limited to 

2,800oF; above this temperature the silicon-based coating systems afford little practical durability,” due to active oxidation. In addition, “periodic replacement of parts 

may be inevitable.”59

New work on carbon-carbon, reported in 1993, gave greater encouragement as 

it raised the prospect of longer lifetimes. The effort evaluated small samples rather 

than fabricated panels and again used the arcjet installations of NASA-Johnson and 

Ames. Once again there was an orders-of-magnitude difference in the observed 

lifetimes of the carbon-carbon, but now the measured lifetimes extended into the 

hundreds of minutes. A formulation from the firm of Carbon-Carbon Advanced 
252

Facing the Heat Barrier: A History of Hypersonics

253

Why NASP Fell Short

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—structural carbon-carbon, high-temperature beryllium—that might 

have saved the situation. With this, NASP spiraled to its end. It also left its supporters with renewed appreciation for rockets as launch vehicles, which had been flying 

to orbit for decades.

counted as a success because it had come within 8°F of the specified temperature.61

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, successfully achieved the Mach 2.2 

peak load, incorporated torsion, and 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, suitable for a 

hot primary structure, but it was clear that this future would not arrive during the 

NASP program.

Then there was beryllium. It had only two-thirds the density of aluminum and 

possessed good strength, but its temperature range was limited. The conventional 

metal had a limit of some 850°F, but an alloy from Lockheed called Lockalloy, which 

contained 38 percent aluminum, was rated only for 600°F. It had never become a 

mainstream engineering material like titanium, but for NASP 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 high-temperature formulations toward the development of beryllium as 

a material suitable for use at cryogenic temperatures. Standard forms of this metal 

became unacceptably brittle when only slightly colder than −100°F, but cryo-beryllium 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 but were winning only modest improvements in lowtemperature 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.62

In sum, the NASP materials effort scored a major advance with Beta-21S, 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 through the entire design, 

leading to a configuration that would have been smaller and lighter overall.

Overall, 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.63 This slow progress encouraged conservatism in 

drawing up the bill of materials, but this conservatism carried a penalty.
254

Facing the Heat Barrier: A History of Hypersonics

255

Why NASP Fell Short

33 Aviation Week, 3 April 1989, p. 58; Thompson, Space Log, Vol. 30 (1994), pp. 5-6.

34 Aviation Week, 11 May 1992, p. 54; AIAA Paper 92-5023; General Accounting Office Report 

NSIAD-92-5.

35 Author interview, Paul Czysz, 13 March 1986. Folder 18649, NASA Historical Reference Collection, NASA History Division, Washington, D.C. 20546.

36 DTIC AD-350952; also available as CASI 64X-17126.

37 AIAA Paper 92-3499; Aviation Week, 11 May 1992, p. 54.

38 DTIC ADB-191715 (quote, p. 3-22).

39 Escher, “Cryogenic,” pp. 5-6; DTIC ADB-191715, p. 3-20.

40 AIAA Paper 91-1944; DTIC ADB-191715, p. 3-20.

41 Crickmore, SR-71, p. 95.

42 Ejector ramjets: AIAA Paper 86-1680, Fig. 3. Turbojets: Aviation Week: 3 May 1954, p. 44; 2 

May 1960, p. 145.

43 AIAA Paper 89-0008 (quote, p. 1).

44 See also 1989 NASP Technology Maturation Plan, task 2.1.02.

45 DTIC ADB-197041, Attachment D.

46 Ibid. 

47 Schweikart, Quest, p. 300; quote: DTIC ADB-197041, p. 2.

48 Author interview, Paul Czysz, 13 March 1986. Notes available, Folder 18649, NASA Historical 

Reference Collection, NASA History Division, Washington, D.C. 20546.

49 Heppenheimer, High Technology, September 1986, pp. 54-55; author interview, Philip Parrish, 

21 March 1986. Folder 18649, NASA Historical Reference Collection, NASA History Division, Washington, D.C. 20546. 

50 AIAA Paper 90-5207; DTIC ADB-192559; Aviation Week, 3 May 1993, p. 36.

51 AIAA Papers 91-5008, 93-2564.

52 AIAA Paper 93-2564 (quote, p. 8); Aviation Week, 3 February 1992, p. 52.

53 DTIC ADB-196085.

54 Brown, Aerospace America, August 1992, pp. 66-67.

55 AIAA Paper 90-5206; Preston, American Steel, pp. 8-9.

56 General Accounting Office Report NSIAD-93-71; DTIC ADB-192559.

57 AIAA Paper 89-5010.

58 DTIC ADB-191699, ADB-192559; ADB-190937, pp. 3-5.

59 DTIC ADB-191699 (quotes, p. 9).

60 DTIC ADB-191402, pp. 117-49.

61 DTIC ADB-191627.

62 DTIC ADB-191898; AIAA Paper 95-6131.

63 General Accounting Office Report NSIAD-93-71, table, p. 20.

1 Author interviews, George Baum, 15 May 1986; Robert Williams, 1 May 1986. Folder 18649, 

NASA Historical Reference Collection, NASA History Division, Washington, D.C. 20546. 

2 Jameson, Science, 28 July 1989, pp. 361-71.

3 AIAA Paper 85-1509.

4 AIAA Paper 84-0387.

5 AIAA Paper 93-5101.

6 Aerospace America, September 1987, p. 37.

7 AIAA Paper 90-5232.

8 AIAA Papers 90-5231, 90-5233; Journal of Spacecraft and Rockets, January-February 1999, p. 13.

9 Dwoyer et al., Aerospace America, October 1987, pp. 32-35 (quote, p. 35).

10 AIAA Papers 90-5232 (quote, p. 12), 92-5049 (quote, p. 2).

11 DTIC ADA-274530 (quote, p. i).

12 Heppenheimer, “Tractable”; Lin, Statistical, pp. 15-17, 43-45.

13 Lee et al., Statistical, pp. 78-83; Kuethe and Chow, Foundations, pp. 393-94; AIAA Paper 83-

1693.

14 AIAA Paper 78-257. Quote: Author interview, Anthony Jameson, 4 June 1990. See Mosaic, 

Spring 1991, p. 30.

15 Quote: AIAA Paper 90-1480, p. 3.

16 Quote: AIAA Paper 83-1693, p. 1.

17 AIAA Paper 85-1652 (quotes, pp. 9, 12).

18 Shapiro, Compressible, p. 1085.

19 Quotes: AIAA Paper 90-5247, p. 1.

20 Author interviews: Anatol Roshko, 8 June 1990; Peter Bradshaw, 11 June 1990. See Mosaic, 

Spring 1991, p. 38. 

21 Quotes: AIAA Paper 93-0200, p. 2. Author interviews: John Lumley, 5 June 1990; Anthony 

Jameson, 6 June 1990. See Mosaic, Spring 1991, pp. 32, 38.

22 Author interview, Anatol Roshko, 8 June 1990, as noted. 

23 Author interview, William Dannevik, 11 June 1990. See Mosaic, Spring 1991, p. 38. 

24 Author interview, Anatol Roshko, 8 June 1990, as noted.

25 DTIC ADA-274530. Quote: Aerospace America, October 1987, 35.

26 AIAA Paper 95-6055, p. 3 (includes quote).

27 Ibid. 

28 1989 NASP Technical Maturation Plan, task 2.1.02; Air Power History, Fall 2000, diagram, p. 

26; DTIC ADB-191715, p. ES-4; ADB-968993, p. 24.

29 AIAA Paper 86-1680.

30 AIAA Paper 91-1944; author interview, Paul Czysz,13 March 1986. Folder 18649, NASA 

Historical Reference Collection, NASA History Division, Washington, D.C. 20546. 

31 DTIC ADB-191464, pp. 1, 10.

32 DTIC ADB-192311, task 1561.

On 7 December 1995 the entry probe of the Galileo spacecraft plunged into 

the atmosphere of Jupiter. It did not plummet directly downward but sliced into 

that planet’s hydrogen-rich envelope at a gentle angle as it followed a trajectory that 

took it close to Jupiter’s edge. The probe 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. Peak 

heating came to 11,800 BTU per square foot-second, corresponding to a radiative equilibrium temperature of 12,000°F. The heat load totaled 141,800 BTU per 

square foot, enough to boil 150 pounds of water for each square foot of heatshield 

surface.1

 The deceleration peaked at 228 g, which was tantamount to slamming 

from a speed of 5,000 miles per hour to a standstill in a single second. Yet the 

probe survived. It deployed a parachute and transmitted data from every one of its 

onboard instruments for nearly an hour, until it was overwhelmed within the depths 

of the atmosphere.2

It used an ablative heatshield, and as an exercise in re-entry technology, the design 

was straightforward. The nose cap was of chopped and molded carbon phenolic 

composite; the rest of the main heatshield was tape-wrapped carbon phenolic. The 

maximum thickness was 5.75 inches. The probe also mounted an aft heatshield, 

which was of phenolic nylon. The value of these simple materials, under the extreme 

conditions of Jupiter atmosphere entry, showed beyond doubt that the problem of 

re-entry was well in hand.3

Other activities have done less well. The X-33 and X-34 projects, which sought 

to build next-generation shuttles using NASP materials, failed utterly. Test scramjets 

have lately taken to flight but only infrequently. Still, work in CFD continues to 

flourish. Today’s best supercomputers offer a million times more power than the 

ancestral Illiac 4, the top computer of the mid-1970s. This ushers in the important 

new topic of Large Eddy Simulation (LES). It may enable us to learn, via computation, just how good scramjets may become.

257

Hypersonics

After NASP

9
258

Facing the Heat Barrier: A History of Hypersonics

259

Hypersonics After NASP

with a ground crew of only 15 people along with three more in its control center. It 

flew no higher than a few thousand feet, but it became the first rocket in history to 

abort a flight and execute a normal landing.5

The Clinton Administration came to Washington in January 1993. Dan Goldin, 

the NASA Administrator, soon chartered a major new study of launch options 

called Access to Space. Arnold Aldrich, Associate Administrator for Space Systems 

Development, served as its director. With NASP virtually on its deathbed, the work 

comprised three specific investigations. Each addressed a particular path toward a 

new generation of launch vehicles, which could include a new shuttle.

Managers at NASA Headquarters and at NASA-Johnson considered how 

upgrades to current expendables, and to the existing shuttle, might maintain them 

in service through the year 2030. At NASA-Marshall, a second group looked at 

prospects for new expendables that could replace existing rockets, including the 

shuttle, beginning in 2005. A collaboration between Headquarters and Marshall 

also considered a third approach: development of an entirely new reusable launch 

vehicle, to replace the shuttle and current expendables beginning in 2008.6

Engineers in industry were ready with ideas of their own. At Lockheed’s famous 

Skunk Works, manager David Urie already had a concept for a fully-reusable singlestage vehicle that was to fly to orbit. It used a lifting-body configuration that drew 

on an in-house study of a vehicle to rescue crews from the space station. Urie’s 

design was to be built as a hot structure with metal external panels for thermal protection and was to use high-performing rocket engines from Rocketdyne that would 

burn liquid hydrogen and liquid oxygen. This concept led to the X-33.7

Orbital Sciences was also stirring the pot. During the spring of 1993, this company conducted an internal study that examined prospects for a Pegasus follow-on. 

Pegasus used solid propellant in all three of its stages, but the new effort specifically 

considered the use of liquid propellants for higher performance. Its concept took 

shape as an air-launched two-stage vehicle, with the first stage being winged and 

fully reusable while the second stage, carried internally, was to fly to orbit without 

being recovered. Later that year executives of Orbital Sciences approached officials 

of NASA-Marshall to ask whether they might be interested, for this concept might 

complement that of Lockheed by lifting payloads of much lesser weight. This initiative led in time to the X-34.8

NASA’s Access to Space report was in print in January 1994. Managers of the 

three option investigations had sought to make as persuasive a case as possible for 

their respective alternatives, and the view prevailed that technology soon would be 

in hand to adopt Lockheed’s approach. In the words of the report summary,

The study concluded that the most beneficial option is to develop and 

deploy a fully reusable single-stage-to-orbit (SSTO) pure-rocket launch 

The X-33 and X-34

During the early 1990s, as NASP passed its peak of funding and began to falter, 

two new initiatives showed that there still was much continuing promise in rockets. 

The startup firm of Orbital Sciences Corporation had set out to become the first 

company to develop a launch vehicle as a commercial venture, and this rocket, 

called Pegasus, gained success on its first attempt. This occurred in April 1990, as 

NASA’s B-52 took off from Edwards AFB and dropped it into flight. Its first stage 

mounted wings and tail surfaces. Its third stage carried a small satellite and placed 

it in orbit.4

In a separate effort, the Strategic Defense Initiative Office funded the DC-X project of McDonnell Douglas. This single-stage vehicle weighed some 40,000 pounds 

when fueled and flew with four RL10 rocket engines from Pratt & Whitney. It took 

off and landed vertically, like Flash Gordon’s rocket ship, using rocket thrust during 

the descent and avoiding the need for a parachute. It went forward as an exercise in 

rapid prototyping, with the contract being awarded in August 1991 and the DC-X 

being rolled out in April 1993. It demonstrated both reusability and low cost, flying 

The DC-X, which flew, and the X-33, which did not. (NASA)
260

Facing the Heat Barrier: A History of Hypersonics

261

Hypersonics After NASP

The National Aeronautics and Space Administration will be the lead agency 

for technology development and demonstration for next generation reusable 

space transportation systems, such as the single-stage-to-orbit concept.13

The Pentagon’s assignment led to the Evolved Expendable Launch Vehicle Program, which brought development of the Delta 4 family and of new versions of the 

Atlas.14

The new policy broke with past procurement practices, whereby NASA had paid 

the full cost of the necessary research and development and had purchased flight 

vehicles under contract. Instead, the White House took the view that the private 

sector could cover these costs, developing the next space shuttle as if it were a new 

commercial airliner. NASA’s role still was critical, but this was to be the longstanding role of building experimental flight craft to demonstrate pertinent technologies. 

The policy document made this clear:

The objective of NASA’s technology development and demonstration effort 

is to support government and private sector decisions by the end of this 

decade on development of an operational next generation reusable launch 

system.

Research shall be focused on technologies to support a decision no 

later than December 1996 to proceed with a sub-scale flight demonstration 

which would prove the concept of single-stage-to-orbit....

It is envisioned that the private sector could have a significant role 

in managing the development and operation of a new reusable space 

transportation system. In anticipation of this role, NASA shall actively 

involve the private sector in planning and evaluating its launch technology 

activities.15

This flight demonstrator became the X-33, with the smaller X-34 being part of 

the program as well. In mid-October NASA issued Cooperative Agreement Notices, 

which resembled requests for proposals, for the two projects. At a briefing to industry representatives held at NASA-Marshall on 19 October 1994, agency officials 

presented year-by-year projections of their spending plans. The X-33 was to receive 

$660 million in federal funds—later raised to $941 million—while the X-34 was 

slated for $70 million. Contractors were to add substantial amounts of their own 

and to cover the cost of overruns. Orbital Sciences was a potential bidder and held 

no contract, but its president, David Thompson, was well aware that he needed 

deeper pockets. He turned to Rockwell International and set up a partnership.16

The X-34 was the first to go to contract, as NASA selected the Orbital Sciences 

proposal in March 1995. Matching NASA’s $70 million, this company and Rockvehicle fleet incorporating advanced technologies, and to phase out current 

systems beginning in the 2008 time period....

The study determined that while the goal of achieving SSTO fully 

reusable rocket launch vehicles had existed for a long time, recent advances 

in technology made such a vehicle feasible and practical in the near term 

provided that necessary technologies were matured and demonstrated prior 

to start of vehicle development.9

Within weeks NASA followed with a new effort, the Advanced Launch Technology Program. It sought to lay technical groundwork for a next-generation shuttle, 

as it solicited initiatives from industry that were to pursue advances in structures, 

thermal protection, and propulsion.10

The Air Force had its own needs for access to space and had generally been more 

conservative than NASA. During the late 1970s, while that agency had been building the shuttle, the Air Force had pursued the Titan 34D as a new version of its Titan 

3. More recently that service had gone forward with its upgraded Titan 4.11 In May 

1994 Lieutenant General Thomas Moorman, Vice Commander of the Air Force’s 

Space Command, released his own study that was known as the Space Launch Modernization Plan. It considered a range of options that paralleled NASA’s, including development of “a new reusable launch system.” However, whereas NASA had 

embraced SSTO as its preferred direction, the Air Force study did not even mention this as a serious prospect. Nor did it recommend a selected choice of launch 

system. In a cover letter to the Deputy Secretary of Defense, John Deutch, Moorman wrote that “this study does not recommend a specific program approach” but 

was intended to “provide the Department of Defense a range of choices.” Still, the 

report made a number of recommendations, one of which proved to carry particular 

weight: “Assign DOD the lead role in expendable launch vehicles and NASA the 

lead in reusables.”12

The NASA and Air Force studies both went to the White House, where in 

August the Office of Science and Technology Policy issued a new National Space 

Transportation Policy. It divided the responsibilities for new launch systems in the 

manner that the Air Force had recommended and gave NASA the opportunity to 

pursue its own wishes as well:

The Department of Defense (DoD) will be the lead agency for improvement 

and evolution of the current U.S. expendable launch vehicle (ELV) fleet, 

including appropriate technology development.

The National Aeronautics and Space Administration (NASA) will 

provide for the improvement of the Space Shuttle system, focusing on 

reliability, safety, and cost-effectiveness.
262

Facing the Heat Barrier: A History of Hypersonics

263

Hypersonics After NASP

Mach 15, well short of orbital velocity, but would subject its thermal protection to 

a demanding test.20

No rocket craft of any type had ever flown to orbit as a single stage. NASA hoped 

that vehicles such as VentureStar not only would do this but would achieve low 

cost, cutting the cost of a pound in orbit from the $10,000 of the space shuttle to 

as little as $1,000.21 The X-33 was to demonstrate the pertinent technology, which 

was being pursued under NASA’s Advanced Launch Technology Program of 1994. 

Developments based on this program were to support the X-34 as well.

Lightweight structures were essential, particularly for the X-33. Accordingly, 

there was strong interest in graphite-composite tanks and primary structure. This 

represented a continuation of NASP activity, which had anticipated a main hydrogen tank of graphite-epoxy. The DC-X supported the new work, as NASA took 

it over and renamed it the DC-XA. Its oxygen tank had been aluminum; a new 

one, built in Russia, used an aluminum-lithium alloy. Its hydrogen tank, also of 

aluminum, gave way to one of graphite-epoxy with lightweight foam for internal 

insulation. This material also served for an intertank structure and a feedline and 

valve assembly.22

Rapid turnaround offered a particularly promising road to low launch costs, and 

the revamped DC-XA gave support in this area as well. Two launches, conducted in 

June 1996, demonstrated turnaround and reflight in only 26 hours, again with its 

ground crew of only 15.23

Thermal protection raised additional issues. The X-34 was to fly only to Mach 8 

and drew on space shuttle technology. Its surface was to be protected with insulation 

blankets that resembled those in use on the shuttle orbiter. These included the High 

Heat Blanket for the X-34 undersurface, rated for 2,000°F, with a Nextel 440 fabric 

and Saffil batting. The nose cap as well as the wing and rudder leading edges were 

protected with Fibrous Refractory Composite Insulation, which formed the black 

silica tiles of the shuttle orbiter. For the X-34, these tiles were to be impregnated 

with silicone to make them water resistant, impermeable to flows of hot gas, and 

easier to repair.24

VentureStar faced the demands of entry from orbit, but its re-entry environment 

was to be more benign than that of the shuttle. The shuttle orbiter was compact in 

size and relatively heavy and lost little of its orbital energy until well into the atmosphere. By contrast, VentureStar would resemble a big lightweight balloon when 

it re-entered after expending its propellants. The VentureStar thermal protection 

system was to be tested in flight on the X-33. It had the form of a hot structure, with 

radiative surface panels of carbon-carbon, Inconel 617 nickel alloy, and titanium, 

depending on the temperature.25

In an effort separate from that of the X-33, elements of this thermal protection were given a workout by being mounted to the space shuttle Endeavour and 

tested during re-entry. Thoughts of such tests dated to 1981 and finally were realwell each agreed to put up $60 million, which meant that the two corporations 

together were to provide more than 60 percent of the funding. Their partnership, 

called American Space Lines, anticipated developing an operational vehicle, the X34B, that would carry 2,500 pounds to orbit. Weighing 108,500 pounds when 

fully fueled, it was to fly from NASA’s Boeing 747 that served as the shuttle’s carrier 

aircraft. Its length of 88 feet compared with 122 feet for the space shuttle orbiter.17

Very quickly an imbroglio developed over the choice of rocket engine for NASA’s 

test craft. The contract called for use of a Russian engine, the Energomash RD-120 

that was being marketed by Pratt & Whitney. Rockwell, which owned Rocketdyne, 

soon began demanding that its less powerful RS-27 engine be used instead. “The 

bottom line is Rockwell came in two weeks ago and said ‘Use our engine or we’ll 

walk,’” a knowledgeable industry observer told Aviation Week.

18

As the issue remained unresolved, Orbital Sciences missed program milestone 

dates for airframe design and for selecting between configurations. Early in November NASA responded by handing Orbital a 14-day suspension notice. This led to 

further discussions, but even the personal involvement of Dan Goldin failed to 

resolve the matter. In addition, the X-34B concept had grown to as much as 140,000 

pounds. Within the program, strong private-sector involvement meant that privatesector criteria of profitability were important, and Orbital determined that the new 

and heavy configuration carried substantial risk of financial loss. Early in 1996 company officials called for a complete redesign of NASA’s X-34 that would substantially reduce its size. The agency responded by issuing a stop-work order. Rockwell 

then made its move by bailing out as well. With this, the X-34 appeared dead.

But it soon returned to life, as NASA prepared to launch it anew. It now was 

necessary to go back to square one and again ask for bids and proposals, and again 

Orbital Sciences was in the running, this time without a partner. The old X-34 

had amounted to a prototype of the operational X-34B, approaching it in size and 

weight while also calling for use of NASA’s Boeing 747. The company’s new concept 

was only 58 feet long compared with 83; its gross weight was to be 45,000 pounds 

rather than 120,000. It was not to launch payloads into orbit but was to serve as a 

technology demonstrator for an eventual (and larger) first stage by flying to Mach 8. 

In June 1996 NASA selected Orbital again as the winner, choosing its proposal over 

competing concepts from such major players as McDonnell Douglas, Northrop 

Grumman, Rockwell, and the Lockheed Martin Skunk Works.19

Preparations for the X-33 had meanwhile been going forward as well. Design 

studies had been under way, with Lockheed Martin, Rockwell, and McDonnell 

Douglas as the competitors. In July 1996 Vice President Albert Gore announced 

that Lockheed had won the prize. This company envisioned a commercial SSTO 

craft named VentureStar as its eventual goal. It was to carry a payload of 59,000 

pounds to low Earth orbit, topping the 51,000 pounds of the shuttle. Lockheed’s 

X-33 amounted to a version of this vehicle built at 53 percent scale. It was to fly to 
264

Facing the Heat Barrier: A History of Hypersonics

265

Hypersonics After NASP

be higher to properly exercise the thermal protection. The X-33 then was to glide 

onward to a landing at Malmstrom AFB in northern Montana, 950 miles from 

Edwards.31

The original program plan called for rollout of a complete flight vehicle on 1 

November 1998. When that date arrived, though, the effort faced a five-month 

schedule slip. This resulted from difficulties with the rocket engines.32 Then in 

December, two days before Christmas, the program received a highly unwelcome 

present. A hydrogen fuel tank, under construction at a Lockheed Martin facility in 

Sunnyvale, California, sustained major damage within an autoclave. An inner wall 

of the tank showed delamination over 90 percent of its area, while another wall 

sprang loose from its frame. The tank had been inspected using ultrasound, but this 

failed to disclose the incipient problem, which raised questions as to the adequacy 

of inspection procedures as well as of the tank design itself. Another delay was at 

hand of up to seven months.

By May 1999 the weight at main engine cutoff was up to 83,000 pounds, including unburned residual propellant. Cleon Lacefield, the Lockheed Martin program 

manager, continued to insist bravely that the vehicle would reach at least Mach 13, 

but working engineers told Aviation Week that the top speed had been Mach 10 for 

quite some time and that “the only way it’s getting to Malmstrom is on the back of 

a truck.”33 The commercial VentureStar concept threatened to be far more demanding, and during that month Peter Teets, president and CEO of Lockheed Martin, 

told the U.S. Senate Commerce and Science Committee that he could not expect to 

attract the necessary private-sector financing. “Wall Street has spoken,” he declared. 

“They have picked the status quo; they will finance systems with existing technology. They will not finance VentureStar.”34

By then the VentureStar design had gone over to aluminum tanks. These were 

heavier than tanks of graphite-epoxy, but the latter brought unacceptable technical 

risks because no autoclave existed that was big enough to fabricate such tankage. 

Lockheed Martin designers reshaped VentureStar and accepted a weight increase 

from 2.6 million pounds to 3.3 million. (It had been 2.2 million in 1996.) The use 

of graphite-epoxy in the X-33 tank now no longer was relevant to VentureStar, but 

this was what the program held in hand, and a change to aluminum would have 

added still more weight to the X-33.

During 1999 a second graphite-epoxy hydrogen tank was successfully assembled at Lockheed Martin and then was shipped to NASA-Marshall for structural 

tests. Early in November it experienced its own failure, showing delamination and 

a ripped outer skin along with several fractures or breaks in the skin. Engineers had 

been concerned for months about structural weakness, with one knowledgeable 

specialist telling Aviation Week, “That tank belonged in a junkyard, not a test stand.” 

The program now was well on its way to becoming an orphan. It was not beloved 

ized during Mission STS-77 in May 1996. Panels of Inconel 617 and of Ti-1100 

titanium, measuring 7 by 10 inches, were mounted in recessed areas of the fuselage 

that lay near the vertical tail and which were heated only to approximately 1,000°F 

during re-entry. Both materials were rated for considerably higher temperatures, but 

this successful demonstration put one more arrow in NASA’s quiver.26

For both VentureStar and its supporting X-33, light weight was critical. The 

X-30 of NASP had been designed for SSTO operation, with a structural mass fraction—the ratio of unfueled weight to fully fueled weight—of 25 percent.27 This 

requirement was difficult to achieve because most of the fuel was slush hydrogen, 

which has a very low density. This ballooned the size of the X-30 and increased the 

surface area that needed structural support and thermal protection. VentureStar was 

to use rockets, which had less performance than scramjets. It therefore needed more 

fuel, and its structural mass fraction, including payload, engines, and thermal protection, was less than 12 percent. However, this fuel included a great deal of liquid 

oxygen, which was denser than water and drove up the weight of the propellant. 

This low structural mass fraction therefore appeared within reach, and for the X-33, 

the required value was considerably less stringent. Its design called for an empty 

weight of 63,000 pounds and a loaded weight of 273,000, for a structural mass 

fraction of 23 percent.28

Even this design goal imposed demands, for while liquid oxygen was dense and 

compact, liquid hydrogen still was bulky and again enlarged the surface area. Designers thus made extensive use of lightweight composites, specifying graphite-epoxy 

for the hydrogen tanks. A similar material, graphite-bismaleimide, was to serve for 

load-bearing trusses as well as for the outer shell that was to support the thermal 

protection. This represented the X-30’s road not taken, for the NASP thermal environment during ascent had been so severe that its design had demanded a primary 

structure of titanium-matrix composite, which was heavier. The lessened requirements of VentureStar’s thermal protection meant that Lockheed could propose to 

reach orbit using materials that were considerably less heavy—that indeed were 

lighter than aluminum. The X-33 design saved additional weight because it was to 

be unpiloted, needing no flight deck and no life-support system for a crew.29

But aircraft often gain weight during development, and the X-33 was no exception. Starting in mid-1996 with a dry weight of 63,000 pounds, it was at 80,000 a 

year later, although a weight-reduction exercise trimmed this to 73,000.30 Managers 

responded by cutting the planned top speed from Mach 15 or more to Mach 13.8. 

Jerry Rising, vice president at the Skunk Works that was the X-33’s home, explained 

that such a top speed still would permit validation of the thermal protection in 

flight test. The craft would lift off from Edwards AFB and follow a boost-glide trajectory, reaching a peak altitude of 300,000 feet. The vehicle then would be lower in 

the atmosphere than previously planned, and the heating rate would consequently 
266

Facing the Heat Barrier: A History of Hypersonics

267

Hypersonics After NASP

Scramjets Take Flight

On 28 November 1991 a Soviet engine flew atop an SA-5 surface-to-air missile in an attempt to demonstrate supersonic combustion. The flight was launched 

from the Baikonur center in Kazakhstan and proceeded ballistically, covering some 

112 miles. The engine did not produce propulsive thrust but rode the missile while 

mounted to its nose. The design had an axisymmetric configuration, resembling 

that of NASA’s Hypersonic Research Engine, and the hardware had been built at 

Moscow’s Central Institute of Aviation Motors (CIAM).

As described by Donat Ogorodnikov, the center director, the engine performed 

two preprogrammed burns during the flight. The first sought to demonstrate the 

important function of transition from subsonic to supersonic combustion. It was 

initiated at 59,000 feet and Mach 3.5, as the rocket continued to accelerate. Ogorodnikov asserted that after fifteen seconds, near Mach 5, the engine went over to 

supersonic combustion and operated in this mode for five seconds, while the rocket 

accelerated to Mach 6 at 92,000 feet. Within the combustor, internal flow reached 

a measured speed of Mach 3. Pressures within the combustor were one to two atmospheres.

The second engine burn lasted ten seconds. This one had the purpose of verifying the design of the engine’s ignition system. It took place on the downward leg of 

the trajectory, as the vehicle descended from 72,000 feet and Mach 4.5 to 59,000 

feet and Mach 3.5. This burn involved only subsonic combustion. Vyacheslav Vinogradov, chief of engine gasdynamics at CIAM, described the engine as mounting 

three rows of fuel injectors. Choice of an injector row, out of the three available, was 

to help in changing the combustion mode.

The engine diameter at the inlet was 9.1 inches; its length was 4.2 feet. The 

spike, inlet, and combustor were of stainless steel, with the spike tip and cowl leadRussian flight-test scramjet. (Aviation Week and Space Technology)

by NASA, which refused to increase its share of funding above $941 million, while 

the in-house cost at Lockheed Martin was mounting steadily.35

The X-33 effort nevertheless lingered through the year 2000. This was an election year, not a good time to cancel a billion-dollar federal program, and Al Gore 

was running for president. He had announced the contract award in 1996, and in 

the words of a congressional staffer, “I think NASA will have a hard time walking 

away from the X-33 until after the election. For better or worse, Al Gore now has 

ownership of it. They can’t admit it’s a failure.”36

The X-34 was still in the picture, as a substantial effort in its own right. Its loaded 

weight of 47,000 pounds approached the 56,000 of the X-15 with external tanks, 

built more than 30 years earlier.37 Yet despite this reduced weight, the X-34 was to 

reach Mach 8, substantially exceeding the Mach 6.7 of the X-15. This reflected the 

use of advanced materials, for whereas the X-15 had been built of heavy Inconel X, 

the X-34 design specified lightweight composites for the primary structure and fuel 

tank, along with aluminum for the liquid-oxygen tank.38

Its construction went forward without major mishaps because it was much 

smaller than the X-33. The first of them reached completion in February 1999, but 

during the next two years it never came close to powered flight. The reason was that 

the X-34 program called for use of an entirely new engine, the 60,000-pound-thrust 

Fastrak of NASA-Marshall that burned liquid oxygen and kerosene. This engine 

encountered development problems, and because it was not ready, the X-34 could 

not fly under power.39

Early in March 2001, with George W. Bush in the White House, NASA pulled 

the plug. Arthur Stephenson, director of NASA-Marshall, canceled the X-34. This 

reflected the influence of the Strategic Defense Initiative Office, which had maintained a continuing interest in low-cost access to orbit and had determined that the 

X-34’s costs outweighed the benefits. Stephenson also announced that the cooperative agreement between NASA and Lockheed Martin, which had supported the X33, would expire at the end of the month. He then pronounced an epitaph on both 

programs: “One of the things we have learned is that our technology has not yet 

advanced to the point that we can successfully develop a new reusable launch vehicle 

that substantially improves safety, reliability, and affordability.”40

One could say that the X-30 effort went farther than the X-33, for the former 

successfully exercised a complete hydrogen tank within its NIFTA project, whereas 

the latter did not. But the NIFTA tank was subscale, whereas those of the X-33 

were full-size units intended for flight. The reason that NIFTA appears to have done 

better is that NASP never got far enough to build and test a full-size tank for its 

hydrogen slush. Because that tank also was to have been of graphite-epoxy, as with 

the X-33, it is highly plausible that the X-30 would have run aground on the same 

shoal of composite-tank structural failure that sank Lockheed Martin’s rocket craft.41
268

Facing the Heat Barrier: A History of Hypersonics

269

Hypersonics After NASP

It nevertheless was clear that with better quality control in manufacturing and 

with better fault tolerance in the onboard control laws, full success might readily 

be achieved. However, the CIAM design was axisymmetric and hence was of a type 

that NASA had abandoned during the early 1970s. Such scramjets had played no 

role in NASP, which from the start had focused on airframe-integrated configurations. The CIAM project had represented an existing effort that was in a position 

to benefit from even the most modest of allocations; the 1992 flight, for instance, 

received as little as $200,000 from France.48 But NASA had its eye on a completely 

American scramjet project that could build on the work of NASP. It took the name 

Hyper-X and later X-43A.

Its background lay in a 1995 study conducted by McDonnell Douglas, with 

Pratt & Whitney providing concepts for propulsion. This effort, the Dual-Fuel 

Airbreathing Hypersonic Vehicle Study, gave conceptual designs for vehicles that 

could perform two significant missions: weapons delivery and reconnaissance, and 

operation as the airbreathing first stage of a two-stage-to-orbit launch system. This 

work drew interest at NASA Headquarters and led the Hypersonic Vehicles Office 

at NASA-Langley to commission the conceptual design of an experimental airplane 

that could demonstrate critical technologies required for the mission vehicles.

The Hyper-X design grew out of a concept for a Mach 10 cruise aircraft with 

length of 200 feet and range of 8,500 nautical miles. It broke with the NASP 

approach of seeking a highly integrated propulsion package that used an ejector 

ramLACE as a low-speed system. Instead it returned to the more conservative path 

of installing separate types of engine. Hydrocarbon-fueled turboramjets were to 

serve for takeoff, acceleration to Mach 4, and subsonic cruise and landing. Hydrogen-burning scramjets were to take the vehicle to Mach 10. The shape of this vehicle 

defined that of Hyper-X, which was designed as a detailed scale model that was 12 

feet long rather than 200.49

Like the Russian engines, Hyper-X was to fly to its test Mach using a rocket 

booster. But Hyper-X was to advance beyond the Russian accomplishments by separating from this booster to execute free flight. This separation maneuver proved to 

be trickier than it looked. Subsonic bombers had been dropping rocket planes into 

flight since the heyday of Chuck Yeager, and rocket stages had separated in nearvacuum at the high velocities of a lunar mission. However, Hyper-X was to separate 

at speeds as high as Mach 10 and at 100,000 feet, which imposed strong forces 

from the airflow. As the project manager David Reubush wrote in 1999, “To the 

program’s knowledge there has never been a successful separation of two vehicles (let 

alone a separation of two non-axisymmetric vehicles) at these conditions. Therefore, 

it soon became obvious that the greatest challenge for the Hyper-X program was, 

not the design of an efficient scramjet engine, but the development of a separation 

scenario and the mechanism to achieve it.”50

ing edge being fabricated using powder metallurgy. The fuel was liquid hydrogen, 

and the system used no turbopump. Pressure, within a fuel tank that also was stainless steel, forced the hydrogen to flow. The combustor was regeneratively cooled; 

this vaporized the hydrogen, which flowed through a regulator at rates that varied 

from 0.33 pounds per second in low-Mach flight to 0.11 at high Mach.42

The Russians made these extensive disclosures because they hoped for financial 

support from the West. They obtained initial assistance from France and conducted 

a second flight test a year later. The engine was slightly smaller and the trajectory 

was flatter, reaching 85,000 feet. It ignited near Mach 3.5 and sustained subsonic 

combustion for several seconds while the rocket accelerated to Mach 5. The engine 

then transitioned to supersonic combustion and remained in this mode for some 

fifteen seconds, while acceleration continued to Mach 5.5. Burning then terminated 

due to exhaustion of the fuel.43

On its face, this program had built a flightworthy scramjet, had achieved a supersonic internal airflow, and had burned hydrogen within this flow. Even so, this was 

not necessarily the same as accomplishing supersonic combustion. The alleged transition occurred near Mach 5, which definitely was at the low end for a scramjet.44 In 

addition, there are a number of ways whereby pockets of subsonic flow might have 

existed within an internal airstream that was supersonic overall. These could have 

served as flameholders, localized regions where conditions for combustion were particularly favorable.45

In 1994 CIAM received a contract from NASA, with NASA-Langley providing 

technical support. The goal now was Mach 6.5, at which supersonic combustion 

appeared to hold a particularly strong prospect. The original Russian designs had 

been rated for Mach 6 and were modified to accommodate the higher heat loads at 

this higher speed. The flight took place in February 1998 and reached Mach 6.4 at 

70,000 feet, with the engine operating for 77 seconds.46

It began operation near Mach 3.5. Almost immediately the inlet unstarted due 

to excessive fuel injection. An onboard control system detected the unstart and 

reduced the fuel flow, which enabled the inlet to start and to remain started. However, the onboard control failed to detect this restart and failed to permit fuel to flow 

through the first of the three rows of fuel injectors. Moreover, the inlet performance 

fell short of predictions due to problems in fabrication.

At Mach 5.5 and higher, airflow entered the fuel-air mixing zone within the 

combustor at speeds near Mach 2. However, only the two rear rows of injectors 

were active, and burning of their fuel forced the internal Mach number to subsonic 

values. The flow reaccelerated to sonic velocity at the combustor exit. The combination of degraded inlet performance and use of only the rear fuel injectors ensured 

that even at the highest flight speeds, the engine operated primarily in a subsoniccombustion mode and showed little if any supersonic combustion.47
270

Facing the Heat Barrier: A History of Hypersonics

271

Hypersonics After NASP

as a whole to provide sufficient control authority. The wings and tail surfaces were 

constructed of temperature-resistant Haynes 230 alloy. Leading edges of the nose, 

vertical fins, and wings used carbon-carbon. For thermal protection, the vehicle was 

covered with Alumina Enhanced Thermal Barrier tiles, which resembled the tiles of 

the space shuttle.54

Additional weight came from the scramjet. It was fabricated of a copper alloy 

called Glidcop, which was strengthened with very fine particles of aluminum oxide 

dispersed within. This increased its strength at high temperatures, while retaining 

the excellent thermal conductivity of copper. This alloy formed the external surface, 

sidewalls, cowl, and fuel injectors. Some internal surfaces were coated with zirconia 

to form a thermal barrier that protected the Glidcop in areas of high heating. The 

engine did not use its hydrogen fuel as a coolant but relied on water cooling for the 

sidewalls and cowl leading edge. Internal engine seals used braided ceramic rope.55

Because the X-43A was small, its engine tests were particularly realistic. This 

vehicle amounted to a scale model of a much larger operational craft of the future, 

but the engine testing involved ground-test models that were full size for the X-43A. 

Most of the testing took place at NASA-Langley, where the two initial series were 

conducted at the Arc-Heated Scramjet Test Facility. This wind tunnel was described 

in 1998 as “the primary Mach 7 scramjet test facility at Langley.”56

Development tests began at the very outset of the Hyper-X Program. The first 

test article was the Dual-Fuel Experiment (DFX), with a name that reflected links 

to the original McDonnell Douglas study. The DFX was built in 1996 by modifying 

existing NASP engine hardware. It provided a test scramjet that could be modified 

rapidly and inexpensively for evaluation of changes to the flowpath. It was fabricated 

primarily of copper and used no active cooling, relying on heat sink. This ruled out 

tests at the full air density of a flight at Mach 7, which would have overheated this 

engine too quickly for it to give useful data. Even so, tests at reduced air densities 

gave valuable guidance in designing the flight engine.

The DFX reproduced the full-scale height and length of the Hyper-X engine, 

correctly replicating details of the forebody, cowl, and sidewall leading edge. The 

forebody and afterbody were truncated, and the engine width was reduced to 44 

percent of the true value so that this test engine could fit with adequate clearances in 

the test facility. This effort conducted more than 250 tests of the DFX, in four different configurations. They verified predicted engine forces and moments as well as 

inlet and combustor component performances. Other results gave data on ignition 

requirements, flameholding, and combustor-inlet interactions.

Within that same facility, subsequent tests used the Hyper-X Engine Module 

(HXEM). It resembled the DFX, including the truncations fore and aft, and it 

too was of reduced width. But it replicated the design of the flight engine, thereby 

overcoming limitations of the DFX. The HXEM incorporated the active cooling of 

Engineers at Sandia National Laboratory addressed this issue. They initially 

envisioned that the rocket might boost Hyper-X to high altitude, with the separation taking place in near-vacuum. The vehicle then could re-enter and light its 

scramjet. This approach fell by the wayside when the heat load at Mach 10 proved 

to exceed the capabilities of the thermal protection system. The next concept called 

for Hyper-X to ride the underside of its rocket and to be ejected downward as if it 

were a bomb. But this vehicle then would pass through the bow shock of the rocket 

and would face destabilizing forces that its control system could not counter.

Sandia’s third suggestion called for holding the vehicle at the front of the rocket 

using a hinged adapter resembling a clamshell or a pair of alligator jaws. Pyrotechnics would blow the jaws open, releasing the craft into flight. The open jaws then 

were to serve as drag brakes, slowing the empty rocket casing while the flight vehicle 

sailed onward. The main problem was that if the vehicle rolled during separation, 

one of its wings might strike this adapter as it opened. Designers then turned to 

an adapter that would swing down as a single piece. This came to be known as the 

“drop-jaw,” and it served as the baseline approach for a time.51

NASA announced the Hyper-X Program in October 1996, citing a budget of 

$170 million. In February 1997 Orbital Sciences won a contract to provide the 

rocket, which again was to be a Pegasus. A month later the firm of Micro Craft Inc. 

won the contract for the Hyper-X vehicle, with GASL building the engine. Work 

at GASL went forward rapidly, with that company delivering a scramjet to NASALangley in August 1998. NASA officials marked the occasion by changing the name 

of the flight aircraft to X-43A.52

The issue of separation in flight proved not to be settled, however, and developments early in 1999 led to abandonment of the drop-jaw. This adapter extended 

forward of the end of the vehicle, and there was concern that while opening it would 

form shock waves that would produce increased pressures on the rear underside of 

the flight craft, which again could overtax its control system. Wind-tunnel tests 

showed that this indeed was the case, and a new separation mechanism again was 

necessary. This arrangement called for holding the X-43A in position with explosive bolts. When they were fired, separate pyrotechnics were to actuate pistons that 

would push this craft forward, giving it a relative speed of at least 13 feet per second. 

Further studies and experiments showed that this concept indeed was suitable.53

The minimal size of the X-43A meant that there was little need to keep its weight 

down, and it came in at 2,800 pounds. This included 900 pounds of tungsten at the 

nose to provide ballast for stability in flight while also serving as a heat sink. High 

stiffness of the vehicle was essential to prevent oscillations of the structure that could 

interfere with the Pegasus flight control system. The X-43A thus was built with steel 

longerons and with steel skins having thickness of one-fourth inch. The wings were 

stubby and resembled horizontal stabilizers; they did not mount ailerons but moved 
272

Facing the Heat Barrier: A History of Hypersonics

273

Hypersonics After NASP

models were only one foot in length, but they incorporated movable rudders and 

wings. Eighteen-inch models followed, which were as large as these tunnels could 

accommodate, and gave finer increments of the control-surface deflections. Thirtyinch models brought additional realism and underwent supersonic and transonic 

tests in the Unitary Plan Wind Tunnel and the 16-Foot Transonic Tunnel.60

Similar studies evaluated the methods proposed for separation of the X-43A 

from its Pegasus booster. Initial tests used Langley’s Mach 6 and Mach 10 tunnels. 

These were blowdown facilities that did not give long run times, while their test 

sections were too small to permit complete representations of vehicle maneuvers 

during separation. But after the drop-jaw concept had been selected, testing moved 

to tunnel B of the Von Karman Facility at the Arnold Engineering Development 

Center. This wind tunnel operated with continuous flow, in contrast to the blowdown installations of Langley, and provided a 50-inch-diameter test section for use 

at Mach 6. It was costly to test in that tunnel but highly productive, and it accommodated models that demonstrated a full range of relative orientations of Pegasus 

and the X-43A during separation.61

This wind-tunnel work also contributed to inlet development. To enhance 

overall engine performance, it was necessary for the boundary layer upstream of 

this inlet to be turbulent. Natural transition to turbulence could not be counted 

on, which meant that an aerodynamic device of some type was needed to trip the 

boundary layer into turbulence. The resulting investigations ran from 1997 into 

1999 and used both the Mach 6 and Mach 10 Langley wind tunnels, executing 

more than 300 runs. Hypulse, a shock tunnel at GASL, conducted more than two 

dozen additional tests.62

Computational fluid dynamics was used extensively. The wind-tunnel tests that 

supported studies of X-43A separation all were steady-flow experiments, which 

failed to address issues such as unsteady flow in the gap between the two vehicles 

as they moved apart. CFD dealt with this topic. Other CFD analyses examined 

relative orientations of the separating vehicles that were not studied at AEDC. To 

scale wind-tunnel results for use with flight vehicles, CFD solutions were generated 

both for the small models under wind-tunnel conditions and for full-size vehicles 

in flight.63

Flight testing was to be conducted at NASA-Dryden. The first X-43A flight vehicle arrived there in October 1999, with its Pegasus booster following in December. 

Tests of this Pegasus were completed in May 2000, with the flight being attempted 

a year later. The plan called for acceleration to Mach 7 at 95,000 feet, followed by 

10 seconds of powered scramjet operation. This brief time reflected the fact that the 

engine was uncooled and relied on copper heat sink, but it was long enough to take 

data and transmit them to the ground. In the words of NASA manager Lawrence 

Huebner, “we have ground data, we have ground CFD, we have flight CFD—all we 

need is the flight data.”64

the flight version, which opened the door to tests at Mach 7 and at full air density. 

These took place within the large Eight-Foot High Temperature Tunnel (HTT).

The HTT had a test section that was long enough to accommodate the full 12-

foot length of the X-43A underside, which provided major elements of the inlet 

and nozzle with its airframe-integrated forebody and afterbody. This replica of the 

underside initially was tested with the HXEM, thereby giving insight into the aerodynamic effects of the truncations. Subsequent work continued to use the HTT and 

replaced the HXEM with the full-width Hyper-X Flight Engine (HXFE). This was 

a flight-spare Mach 7 scramjet that had been assigned for use in ground testing.

Mounted on its undersurface, this configuration gave a geometrically accurate 

nose-to-tail X-43A propulsion flowpath at full scale. NASA-Langley had conducted 

previous tests of airframe-integrated scramjets, but this was the first to replicate 

the size and specific details of the propulsion system of a flight vehicle. The HTT 

heated its air by burning methane, which added large quantities of carbon dioxide 

and water vapor to the test gas. But it reproduced the Mach, air density, pressure, 

and temperature of flight at altitude, while gaseous oxygen, added to the airflow, 

enabled the engine to burn hydrogen fuel. Never before had so realistic a test series 

been accomplished.57

The thrust of the engine was classified, but as early as 1997 Vince Rausch, the 

Hyper-X manager at NASA-Langley, declared that it was the best-performing scramjet that had been tested at his center. Its design called for use of a cowl door that 

was to protect the engine by remaining closed during the rocket-powered ascent, 

with this door opening to start the inlet. The high fidelity of the HXFE, and of the 

test conditions, gave confidence that its mechanism would work in flight. The tests 

in the HTT included 14 unfueled runs and 40 with fuel. This cowl door was actuated 52 times under the Mach 7 test conditions, and it worked successfully every 

time.58

Aerodynamic wind-tunnel investigations complemented the propulsion tests 

and addressed a number of issues. The overall program covered all phases of the 

flight trajectory, using 15 models in nine wind tunnels. Configuration development 

alone demanded more than 5,800 wind-tunnel runs. The Pegasus rocket called for 

evaluation of its own aerodynamic characteristics when mated with the X-43A, and 

these had to be assessed from the moment of being dropped from the B-52 to separation of the flight vehicle. These used the Lockheed Martin Vought High Speed 

Wind Tunnel in Grand Prairie, Texas, along with facilities at NASA-Langley that 

operated at transonic as well as hypersonic speeds.59

Much work involved evaluating stability, control, and performance characteristics of the basic X-43A airframe. This effort used wind tunnels of McDonnell 

Douglas and Rockwell, with the latter being subsonic. At NASA-Langley, activity 

focused on that center’s 20-inch Mach 6 and 31-inch Mach 10 facilities. The test 
274

Facing the Heat Barrier: A History of Hypersonics

275

Hypersonics After NASP

The flight plan called for each Terrier-Orion to accelerate its scramjet onto a 

ballistic trajectory that was to reach an altitude exceeding 300 kilometers. Near the 

peak of this flight path, an attitude-control system was to point the rocket downward. Once it re-entered the atmosphere, below 40 kilometers, its speed would fall 

off and the scramjet would ignite. This engine was to operate while continuing to 

plunge downward, covering distance into an increasingly dense atmosphere, until it 

lost speed in the lower atmosphere and crashed into the outback.

The flights took place at Woomera Instrumented Range, north of Adelaide. The 

first launch attempt came at the end of October 2001. It flopped; the first stage 

performed well, but the second stage went off course. But nine months later, on 30 

July 2002, the second shot gained full success. The rocket was canted slightly away 

from the vertical as it leaped into the air, accelerating at 22 g as it reached Mach 3.6 

in only six seconds.

This left it still at low altitude while topping the speed of the SR-71, so after the 

second stage with payload separated, it coasted for 16 seconds while continuing to 

ascend. The second stage then ignited, and this time its course was true. It reached a 

peak speed of Mach 7.7. The scramjet went over the top; it pointed its nose downward, and at an altitude of 36 kilometers with its speed approaching Mach 7.8, 

gaseous hydrogen caused it to begin producing thrust. This continued until HyShot 

reached 25 kilometers, when it shut down.

It fired for only five seconds. But it returned data over 40 channels, most of 

which gave pressure readings. NASA itself provided support, with Lawrence Huebner, the X-43A manager, declaring, “We’re very hungry for flight data.” For the 

moment, at least, the Aussies were in the lead.68

But the firm of Micro Craft had built two more X-43As, and the second flight 

took place in March 2004. This time the Pegasus first stage had been modified by 

having part of its propellant removed, to reduce its performance, and the drop altitude was considerably higher.69 In the words of Aviation Week,

The B-52B released the 37,500-lb. stack at 40,000 ft. and the Pegasus 

booster ignited 5 sec. later.... After a few seconds it pulled up and reached 

a maximum dynamic pressure of 1,650 psf. at Mach 3.5 climbing through 

47,000 ft. Above 65,000 ft. it started to push over to a negative angle of 

attack to kill the climb rate and gain more speed. Burnout was 84 sec. 

after drop, and at 95 sec. a pair of pistons pushed the X-43A away from 

the booster at a target condition of Mach 7 and 95,000 ft. and a dynamic 

pressure of 1,060 psf. in a slight climb before the top of a ballistic arc.

After a brief period of stabilization, the X-43A inlet door was opened 

to let air in through the engine.... The X-43A stabilized again because the 

engine airflow changed the trim.... Then silane, a chemical that burns 

upon contact with air, was injected for 3 sec. to establish flame to ignite the 

Launch finally occurred in June 2001. Ordinarily, when flying to orbit, Pegasus was air-dropped at 38,000 feet, and its first stage flew to 207,000 feet prior to 

second-stage ignition. It used solid propellant and its performance could not readily 

be altered; therefore, to reduce its peak altitude to the 95,000 feet of the X-43A, 

it was to be air-dropped at 24,000 feet, even though this lower altitude imposed 

greater loads.

The B-52 took off from Edwards AFB and headed over the Pacific. The Pegasus fell away; its first stage ignited five seconds later and it flew normally for some 

eight seconds that followed. During those seconds, it initiated a pullout to begin its 

climb. Then one of its elevons came off, followed almost immediately by another. As 

additional parts fell away, this booster went out of control. It fell tumbling toward 

the ocean, its rocket motor still firing, and a safety officer sent a destruct signal. The 

X-43A never had a chance to fly, for it never came close to launch conditions.65

A year later, while NASA was trying to recoup, a small group in Australia beat 

the Yankees to the punch by becoming the first in the world to fly a scramjet and 

achieve supersonic combustion. Their project, called HyShot, cost under $2 million, compared with $185 million for the X-43A program. Yet it had plenty of 

technical sophistication, including tests in a shock tunnel and CFD simulations 

using a supercomputer.

Allan Paull, a University of Queensland researcher, was the man who put it 

together. He took a graduate degree in applied mathematics in 1985 and began 

working at that university with Ray Stalker, an engineer who had won a global reputation by building a succession of shock tunnels. A few years later Stalker suffered a 

stroke, and Paull found himself in charge of the program. Then opportunity came 

knocking, in the form of a Florida-based company called Astrotech Space Operations. That firm was building sounding rockets and wanted to expand its activities 

into the Asia and Pacific regions.

In 1998 the two parties signed an agreement. Astrotech would provide two Terrier-Orion sounding rockets; Paull and his colleagues would construct experimental 

scramjets that would ride those rockets. The eventual scramjet design was not airframe-integrated, like that of the X-43A. It was a podded axisymmetric configuration. But it was built in two halves, with one part being fueled with hydrogen while 

the other part ran unfueled for comparison.66

Paull put together a team of four people—and found that the worst of his problems was what he called an “amazing legal nightmare” that ate up half his time. In 

the words of the magazine Air & Space, “the team had to secure authorizations from 

various state government agencies, coordinate with aviation bodies and insurance 

companies in both Australia and the United States (because of the involvement of 

U.S. funding), perform environmental assessments, and ensure their launch debris 

would steer clear of land claimed by Aboriginal tribes.... All told, the preparations 

took three and a half years.”67
276

Facing the Heat Barrier: A History of Hypersonics

277

Hypersonics After NASP

The program still had one unexpended X-43A vehicle that was ready to fly, and 

it flew successfully as well, in November. The goal now was Mach 10. This called 

for beefing up the thermal structure by adding leading edges of solid carbon-carbon 

to the vertical tails along with a coating of hafnium carbide and by making the nose 

blunter to increase the detachment of the bow shock. These changes indeed were 

necessary. Nose temperatures reached 3,600°F, compared with 2,600°F on the Mach 

7 flight, and heating rates were twice as high.

The Pegasus rocket, with the X-43A at its front, fell away from its B-52 carrier 

aircraft at 40,000 feet. Its solid rocket took the combination to Mach 10 at 110,000 

feet. Several seconds after burnout, pistons pushed the X-43A away at Mach 9.8. 

Then, 2.5 seconds after separation, the engine inlet door opened and the engine 

began firing at Mach 9.65. It ran initially with silane to ensure ignition; then the 

engine continued to operate with silane off, for comparison. It fired for a total of 10 

to 12 seconds and then continued to operate with the fuel off. Twenty-one seconds 

after separation, the inlet door closed and the vehicle entered a hypersonic glide. 

This continued for 14 minutes, with the craft returning data by telemetry until it 

struck the Pacific Ocean and sank.

This flight gave a rare look at data taken under conditions that could not be 

duplicated on the ground using continuous-flow wind tunnels. The X-43A had 

indeed been studied in 0.005-second runs within shock tunnels, and Aviation Week

noted that Robert Bakos, vice president of GASL, described such tests as having 

done “a very good job of predicting the flight.” Dynamic pressure during the flight 

was 1,050 pounds per square foot, and the thrust approximately equaled the drag. 

In addition, the engine achieved true supersonic combustion, without internal 

pockets of subsonic flow. This meant that the observations could be scaled to still 

higher Mach values.71

Recent Advances in Fluid Mechanics

The methods of this field include ground test, flight test, and CFD. Ground-test 

facilities continue to show their limitations, with no improvements presently in view 

that would advance the realism of tests beyond Mach 10. A recently announced Air 

Force project, Mariah, merely underscores this point. This installation, to be built at 

AEDC, is to produce flows up to Mach 15 that are to run for as long as 10 seconds, 

in contrast to the milliseconds of shock tunnels. Mariah calls for a powerful electron 

beam to create an electrically charged airflow that can be accelerated with magnets. 

But this installation will require an e-beam of 200 megawatts. This is well beyond 

the state of the art, and even with support from a planned research program, Mariah 

is not expected to enter service until 2015.72

Similar slow progress is evident in CFD, for which the flow codes of recent 

projects have amounted merely to updates of those used in NASP. In designing 

hydrogen. Injection of the gaseous hydrogen fuel ramped up as the silane 

ramped down, lasting 8 sec. The hydrogen flow rate increased through 

and beyond a stoichiometric mixture ratio, and then ramped down to a 

very lean ratio that continued to burn until the fuel was shut off.... The 

hydrogen was stored in 8,000-psi bottles.

Accelerometers showed the X-43A gained speed while fuel was on.... 

Data was gathered all the way to the splashdown 450 naut. mi. offshore at 

about 11 min. after drop.

Aviation Week added that the vehicle accelerated “while in a slight climb at Mach 

7 and 100,000 ft. altitude. The scramjet field is sufficiently challenging that producing thrust greater than drag on an integrated airframe/engine is considered a major 

accomplishment.”70

In this fashion, NASA executed its first successful flight of a scramjet. The overall 

accomplishment was not nearly as ambitious as that planned for the Incremental 

Flight Test Vehicle of the 1960s, for which the velocity increase was to have been 

much greater. Nor did NASA have a follow-on program in view that could draw on 

the results of the X-43A. Still, the agency now could add the scramjet to its list of 

flight engines that had been successfully demonstrated.

X-43A mission to Mach 7. (NASA)
278

Facing the Heat Barrier: A History of Hypersonics

279

Hypersonics After NASP

by using vectors, strings of 256 numbers, but that much of its capability went 

unused when his vector held only five numbers, representing five chemical species. 

The computation also slowed when finding the value of a single constant or when 

taking square roots, which is essential when calculating the speed of sound. Still, he 

adds, “people are happy if they get 50 percent” of a computer’s rated performance. 

“I do get 50 percent, so I’m happy.”78

THE WORLD’S FASTEST SUPERCOMPUTERS (Nov. 2004; updated annually)

Name Manufacturer Location Year

Rated

speed

teraflops

Number 

of 

proces- sors

1 BlueGene IBM Rochester, NY 2004 70,720 32,768

2

Numerical

Aerodynamic

Simulator

Silicon

Graphics NASA-Ames 2004 51,870 10,160

3 Earth

Simulator

Nippon

Electric

Yokohama,

 Japan 2002 35,860 5,120

4 Mare Nostrum IBM Barcelona, Spain 2004 20,530 3,564

5 Thunder

California

Digital

Corporation

Lawrence

Livermore 

National 

Laboratory

2004 19,940 4,096

6 ASCI Q Hewlett-Packard

Los Alamos

National

Laboratory 

2002 13,880 8,192

7 System X Self-made Virginia Tech 2004 12,250 2,200

8 BlueGene

(prototype)

IBM,

Livermore Rochester, NY 2004 11,680 8,192

9 eServer p Series

655 IBM

Naval

Oceanographic

Office

2004 10,310 2,944

10 Tungsten Dell

National Center 

for Supercomputer

Applications

2003 9,819 2,500

500 Superdome 875 Hewlett-Packard SBC Service, Inc. 2004 850.6 416

Source: http://www.top500.org/list/2004/11

the X-43A, the most important such code was the General Aerodynamic Simulation Program (GASP). NASP had used version 2.0; the X-43A used 3.0. The latter 

continued to incorporate turbulence models. Results from the codes often showed 

good agreement with test, but this was because the codes had been benchmarked 

extensively with wind-tunnel data. It did not reflect reliance on first principles at 

higher Mach.

Engine studies for the X-43A used their own codes, which again amounted to those 

of NASP. GASP 3.0 had the relatively recent date of 1996, but other pertinent literature showed nothing more recent than 1993, with some papers dating to the 1970s.73

The 2002 design of ISTAR, a rocket-based combined-cycle engine, showed that 

specialists were using codes that were considerably more current. Studies of the 

forebody and inlet used OVERFLOW, from 1999, while analysis of the combustor 

used VULCAN version 4.3, with a users’ manual published in March 2002. OVERFLOW used equilibrium chemistry while VULCAN included finite-rate chemistry, 

but both solved the Navier-Stokes equations by using a two-equation turbulence 

model. This was no more than had been done during NASP, more than a decade 

earlier.74

The reason for this lack of progress can be understood with reference to Karl 

Marx, who wrote that people’s thoughts are constrained by their tools of production. The tools of CFD have been supercomputers, and during the NASP era the 

best of them had been rated in gigaflops, billions of floating-point operations per 

second.75 Such computations required the use of turbulence models. But recent 

years have seen the advent of teraflop machines. A list of the world’s 500 most powerful is available on the Internet, with the accompanying table giving specifics for 

the top 10 of November 2004, along with number 500.

One should not view this list as having any staying power. Rather, it gives a snapshot of a technology that is advancing with extraordinary rapidity. Thus, in 1980 

NASA was hoping to build the Numerical Aerodynamic Simulator, and to have it 

online in 1986. It was to be the world’s fastest supercomputer, with a speed of one 

gigaflop (0.001 teraflop), but it would have fallen below number 500 as early as 

1994. Number 500 of 2004, rated at 850 gigaflops, would have been number one 

as recently as 1996. In 2002 Japan’s Earth Simulator was five times faster than its 

nearest rivals. In 2004 it had fallen to third place.76

Today’s advances in speed are being accomplished both by increasing the number 

of processors and by multiplying the speed of each such unit. The ancestral Illiac4, for instance, had 64 processors and was rated at 35 megaflops.77 In 2004 IBM’s 

BlueGene was two million times more powerful. This happened both because it had 

512 times more processors—32,768 rather than 64—and because each individual 

processor had 4,000 times more power. Put another way, a single BlueGene processor could do the work of two Numerical Aerodynamic Simulator concepts of 1980.

Analysts are using this power. The NASA-Ames aerodynamicist Christian Stemmer, who has worked with a four-teraflop machine, notes that it achieved this speed 
280

Facing the Heat Barrier: A History of Hypersonics

281

Hypersonics After NASP

LES is farther along in development. It directly simulates the large energy-bearing eddies and goes onward into the upper levels of the cascade. Because its computations do not capture the complete physics of turbulence, LES continues to rely 

on turbulence models to treat the energy flow in the cascade along with the Kolmogorov-scale dissipation. But in contrast to the turbulence models of present-day 

codes, those of LES have a simple character that applies widely across a broad range 

of flows. In addition, their errors have limited consequence for a flow as a whole, in 

an inlet or combustor under study, because LES accurately captures the physics of 

the large eddies and therefore removes errors in their modeling at the outset.81

The first LES computations were published in 1970 by James Deardorff of the 

National Center for Atmospheric Research.82 Dean Chapman, Director of Astronautics at NASA-Ames, gave a detailed review of CFD in the 1979 AIAA Dryden 

Lectureship in Research, taking note of the accomplishments and prospects of 

LES.83 However, the limits of computers restricted the development of this field. 

More than a decade later Luigi Martinelli of Princeton University, a colleague of 

Antony Jameson who had established himself as a leading writer of flow codes, 

declared that “it would be very nice if we could run a large-eddy simulation on a full 

three-dimensional configuration, even a wing.” Large eddies were being simulated 

only for simple cases such as flow in channels and over flat plates, and even then the 

computations were taking as long as 100 hours on a Cray supercomputer.84

Since 1995, however, the Center for Turbulence Research has come to the forefront as a major institution where LES is being developed for use as an engineering 

tool. It is part of Stanford University and maintains close ties both with NASAAmes and with Lawrence Livermore National Laboratory. At this center, Kenneth 

Jansen published LES studies of flow over a wing in 1995 and 1996, treating a 

NACA 4412 airfoil at maximum lift.85 More recent work has used LES in studies 

of reacting flows within a combustor of an existing jet engine of Pratt & Whitney’s 

PW6000 series. The LES computation found a mean pressure drop across the injector of 4,588 pascals, which differs by only two percent from the observed value 

of 4,500 pascals. This compares with a value of 5,660 pascals calculated using a 

Reynolds-averaged Navier-Stokes code, which thus showed an error of 26 percent, 

an order of magnitude higher.86

Because LES computes turbulence from first principles, by solving the NavierStokes equations on a very fine computational grid, it holds high promise as a means 

for overcoming the limits of ground testing in shock tunnels at high Mach. The 

advent of LES suggests that it indeed may become possible to compute one’s way 

to orbit, obtaining accurate results even for such demanding problems as flow in a 

scramjet that is flying at Mach 17.

Parviz Moin, director of the Stanford center, cautions that such flows introduce 

shock waves, which do not appear in subsonic engines such as the PW6000 series, 

and are difficult to treat using currently available methods of LES. But his colleague 

Teraflop ratings, representing a thousand-fold advance over the gigaflops of 

NASP and subsequent projects, are required because the most demanding problems 

in CFD are four-dimensional, including three physical dimensions as well as time. 

William Cabot, who uses the big Livermore machines, notes that “to get an increase 

in resolution by a factor of two, you need 16” as the increase in computational 

speed because the time step must also be reduced. “When someone says, ‘I have a 

new computer that’s an order of magnitude better,’” Cabot continues, “that’s about 

a factor of 1.8. That doesn’t impress people who do turbulence.”79

But the new teraflop machines increase the resolution by a factor of 10. This 

opens the door to two new topics in CFD: Large-Eddy Simulation (LES) and Direct 

Numerical Simulation (DNS).

One approaches the pertinent issues by examining the structure of turbulence 

within a flow. The overall flowfield has a mean velocity at every point. Within it, 

there are turbulent eddies that span a very broad range of stress. The largest carry 

most of the turbulent energy and accomplish most of the turbulent mixing, as in a 

combustor. The smaller eddies form a cascade, in which those of different sizes are 

intermingled. Energy flows down this cascade, from the larger to the smaller ones, 

and while turbulence is often treated as a phenomenon that involves viscosity, the 

transfer of energy along the cascade takes place through inviscid processes. However, 

viscosity becomes important at the level of the smallest eddies, which were studied 

by Andrei Kolmogorov in the Soviet Union and hence define what is called the 

Kolmogorov scale of turbulence. At this scale, viscosity, which is an intermolecular 

effect, dissipates the energy from the cascade into heat. The British meteorologist 

Lewis Richardson, who introduced the concept of the cascade in 1922, summarized 

the matter in a memorable sendup of a poem by England’s Jonathan Swift:

Big whorls have little whorls

Which feed on their velocity;

And little whorls have lesser whorls,

And so on to viscosity.80

In studying a turbulent flow, DNS computes activity at the Kolmogorov scale 

and may proceed into the lower levels of the cascade. It cannot go far because the 

sizes of the turbulent eddies span several orders of magnitude, which cannot be 

captured using computational grids of realistic size. Still, DNS is the method of 

choice for studies of transition to turbulence, which may predict its onset. Such 

simulations directly reproduce the small disturbances within a laminar flow that 

grow to produce turbulence. They do this when they first appear, making it possible 

to observe their growth. DNS is very computationally intensive and remains far 

from ready for use with engineering problems. Even so, it stands today as an active 

topic for research.
282

Facing the Heat Barrier: A History of Hypersonics

283

Hypersonics After NASP

Rutan’s firm of Scaled Composites.89 Such lightweight materials also found use in 

the construction of SpaceShipOne, which was assembled within that plant.

SpaceShipOne brought the prospect of routine commercial flights having the 

performance of the X-15. Built entirely as a privately funded venture, it used a 

simple rocket engine that burned rubber, with nitrous oxide as the oxidizer, and 

reached altitudes as high as 70 miles. A movable set of wings and tail booms, rotating upward, provided stability in attitude during re-entry and kept the craft’s nose 

pointing upward as well. The craft then glided to a landing.

There was no commercial follow-on to Voyager, but today there is serious interest in building commercial versions of SpaceShipOne that will take tourists on brief 

hops into space—and enable them to win astronauts’ wings in the process. Richard Branson, founder of Virgin Airways, is currently sponsoring a new enterprise, 

Virgin Galactic, that aims to do just that. He has formed a partnership with Scaled, 

has sold more than 100 tickets at $200,000 each, and hopes for his first flight late 

in 2008.

And yet.... The top speed of SpaceShipOne was only 2,200 miles per hour, or 

Mach 3.3. Rutan’s vehicle thus stands today as a brilliant exercise in rocketry and 

the design of reusable piloted spacecraft. But it is too slow to qualify as a project in 

hypersonics.90

Is that it, then? Following more than half a century of effort, does the re-entry 

problem stand as the single unambiguous contribution of hypersonics? Air Force 

historian Richard Hallion has written of a “hypersonic revolution,” but from this 

perspective, one may regard hypersonics less as an extension of aeronautics than as 

a branch of materials science, akin to metallurgy. Specialists in that field introduced 

superalloys that extended the temperature limits of jet engines, thereby enhancing their range and fuel economy. Similarly, the hypersonics community developed 

lightweight thermal-protection systems that have found use even in exploring the 

planet Jupiter. Yet one does not speak of a “superalloy revolution,” and hypersonics 

has had similarly limited application.

There remains the issue of the continuing effort to develop the scramjet. This 

work has gone forward as part of an ongoing hope that better methods might be 

devised for ascent to orbit, corresponding perhaps to the jet airliners that drove 

their piston-driven counterparts to the boneyard. Access to space holds undeniable 

importance, and one may speak without challenge of a “satellite revolution” when 

we consider the vital role of such craft in a host of areas: weather forecasting, navigation, tactical warfare, reconnaissance, as well as telecommunications. Yet low-cost 

access remains out of reach and hence continues to justify work on advanced technologies, including scramjets.

Still, despite 40 years of effort, the scramjet continues to stand at two removes 

from importance. The first goal is simply to make it work, by demonstrating flight 

to orbit in a vehicle that uses such engines for propulsion. The X-30 was to fly in 

Heinz Pitsch anticipates rapid progress. He predicted in 2003 that LES will first be 

applied to scramjets in university research, perhaps as early as 2005. He adds that by 

2010 “LES will become the state of the art and will become the method of choice” 

for engineering problems, as it emerges from universities and begins to enter the 

mainstream of CFD.87

Hypersonics and the Aviation Frontier

Aviation has grown through reliance upon engines, and three types have been 

important: the piston motor, turbojet, and rocket. Hypersonic technologies have 

made their largest contributions, not by adding the scramjet to this list, but by 

enhancing the value and usefulness of rockets. This happened when these technologies solved the re-entry problem.

This problem addressed critical issues of the national interest, for it was essential 

to the success of Corona and of the return of film-carrying capsules from orbit. It 

also was a vital aspect of the development of strategic missiles. Still, if such weapons 

had proven to be technically infeasible, the superpowers would have fallen back 

on their long-range bombers. No such backup was available within the Corona 

program. During the mid-1960s the Lunar Orbiter Program used a high-resolution 

system for scanning photographic film, with the data being returned using telemetry.88 But this arrangement had a rather slow data rate and was unsuitable for the 

demands of strategic reconnaissance.

Success in re-entry also undergirded the piloted space program. In 40 years of 

effort, this program has failed to find a role in the mainstream of technical activity 

akin to the importance of automated satellites in telecommunications. Still, piloted 

flight brought the unforgettable achievements of Apollo, which grow warmer in 

memory as the decades pass. 

In a related area, the advent of thermal-protection methods led to the development of aircraft that burst all bounds on speed and altitude. These took form as the 

X-15 and the space shuttle. On the whole, though, this work has led to disappointment. The Air Force had anticipated that airbreathing counterparts of the X-15, 

powered perhaps by ramjets, would come along in the relatively near future. This 

did not happen; the X-15 remains sui generis, a thing unto itself. In turn, the shuttle 

failed to compete effectively with expendable launch vehicles.

This conclusion remains valid in the wake of the highly publicized flights of 

SpaceShipOne, built by the independent inventor Burt Rutan. Rutan showed an 

uncanny talent for innovation in 1986, when his Voyager aircraft, piloted by his 

brother Dick and by Dick’s former girlfriend Jeana Yeager, circled the world on a 

single load of fuel. This achievement had not even been imagined, for no sciencefiction writer had envisioned such a nonstop flight around the world. What made 

it possible was the use of composites in construction. Indeed, Voyager was built at 
284

Facing the Heat Barrier: A History of Hypersonics

285

Hypersonics After NASP

this fashion, although present-day thinking leans more toward using it merely in 

an airbreathing first stage. But at least within the next decade the most that anyone 

hopes for is to accelerate a small test vehicle of the X-43 class.91

Yet even if a large launch vehicle indeed should fly using scramjets, it then will 

face a subsequent test, for it will have to win success in the face of competition from 

existing launchers. The history of aerospace shows several types of craft that indeed 

flew well but that failed in the market. The classic example was the dirigible, which 

was abandoned because it could not be made safe.92

The world still remembers the Hindenburg, but the problems ran deeper than 

the use of hydrogen. Even with nonflammable helium, such airships proved to be 

structurally weak. The U.S. Navy built three large ones—the Shenandoah, Akron, 

and Macon—and quickly lost them all in storms and severe weather. Nor has this 

problem been solved. Dirigibles might be attractive today as aerial cruise ships, 

offering unparalleled views of Caribbean islands, but the safety problem persists.

More recently the Concorde supersonic airliner flew with great style and panache 

but faltered due to its high costs. The Saturn V Moon rocket proved to be too large 

to justify continued production; it lacked payloads that demanded its heft. Piloted 

space flight raises its own questions. It too is very costly, and in the light of experience with the shuttle, perhaps it too cannot be made completely safe.

Yet though scramjets face obstacles both in technology and in the market, they 

will continue to tantalize. Hallion writes that faith in a future for hypersonics “is 

akin to belief in the Second Coming: one knows and trusts that it will occur, but 

one can’t be certain when.” Scramjet advocates will continue to echo the defiant 

words of Eugen Sänger: “Nevertheless, my silver birds will fly!”93

1 Milos et al., Journal of Spacecraft and Rockets, May-June 1999, pp. 298-99.

2 Young et al., Science, 10 May 1996, pp. 837-38 (228 g: Table 3).

3 Milos et al., Journal of Spacecraft and Rockets, May-June 1999, pp. 298-99.

4 Aerospace America: December 1989, pp. 48-49; December 1990, p. 73.

5 Miller, X-Planes, pp. 341-42. Launch crew size: Acta Astronautica, February-April 1996, p. 

326.

6 NASA SP-4407, Vol. IV, pp. 585-88, 591.

7 Aviation Week, 8 July 1996, p. 21.

8 AIAA Paper 95-3777.

9 NASA SP-4407, Vol. IV, pp. 585-86.

10 Acta Astronautica, February-April 1996, p. 323.

11 Heppenheimer, Development, pp. 365-66; Hanley, Aerospace America, July 1991, pp. 34-38.

12 NASA SP-4407, Vol. IV, p. 604; quotes, pp. 605, 617, 624.

13 Ibid., pp. 626-27 (includes extended quote).

14 Acta Astronautica, February-April 1996, p. 324.

15 NASA SP-4407, Vol. IV, pp. 628-29.

16 Aviation Week, 31 October 1994, pp. 22-23; AIAA Paper 95-3777, p. 1. $941 million: NASA 

SP-4407, Vol. IV, p. 634.

17 Aviation Week, 3 April 1995, pp. 44, 48 (X-34B, p. 48). $60 million: Aviation Week, 5 February 

1996, p. 86.

18 Miller, X-Planes, pp. 352-353. Quote: Aviation Week, 6 November 1995, p. 30.

19 Miller, X-Planes, p. 353. Concept comparisons: AIAA Paper 95-3777, pp. 3-4; Aviation Week, 

17 June 1996, p. 31.

20 Miller, X-Planes, p. 344. Payloads: Aviation Week, 8 July 1996, table, pp. 20-21. 53 percent: 

Rising, X-33, p. 6.

21 General Accounting Office Report NSIAD-99-176, p. 4.

22 Acta Astronautica, February-April 1996, pp. 325-26, 329.

23 Aviation Week, 26 April 1999, p. 78.

24 Journal of Spacecraft and Rockets, March-April 1999, pp. 189, 218; Jenkins, Space Shuttle, p. 

395.

25 Miller, X-Planes, p. 346; AIAA Paper 96-4563.

26 Aviation Week, 3 June 1996, p. 92.

27 Air Force Studies Board, Hypersonic, p. 5.

28 Aviation Week, 8 July 1996, table, pp. 20-21.

29 AIAA Paper 96-4563.

30 Aviation Week: 8 July 1996, p. 21; 30 June 1997, p. 28; 10 November 1997, p. 50.

31 Rising, X-33, figure 10; author interview, Jerry Rising, 11 January 1999.

32 Miller, X-Planes, p. 344; Aviation Week, 2 November 1998, pp. 26-27.
286

Facing the Heat Barrier: A History of Hypersonics

287

Hypersonics After NASP

33 Aviation Week, 25 January 1999, pp. 68-69 (weight, p. 68; quote, p. 68).

34 Sietzen, “Wall Street” (includes quotes).

35 Aviation Week, 15 November 1999, pp. 28-30 (quote, p. 28). 2.2 million pounds: Aviation 

Week, 8 July 1996, p. 21.

36 Quote: Aviation Week, 15 November 1999, p. 30.

37 Miller, X-Planes, pp. 203, 357.

38 Ibid., p. 356; AIAA Paper 98-3516.

39 Miller, X-Planes, pp. 354-57.

40 Aviation Week, 5 March 2001, p. 24 (includes quote).

41 NIFTA: AIAA Paper 91-5008.

42 Aviation Week, 30 March 1992, pp. 18-20.

43 Aviation Week, 14/21 December 1992, pp. 70-73.

44 The Incremental Scramjet of the mid-1960s was to achieve transition between Mach 5.4 and 6. 

See Hallion, Hypersonic, pp. VI-xvii to VI-xx. Work at the Applied Physics Laboratory demonstrated transition between Mach 5.9 and 6.2: see AIAA Paper 91-2395.

45 Author interview, Lawrence Huebner, 24 July 2002.

46 NASA/TP 1998-206548, pp. 1, 8-10.

47 AIAA Paper 99-4848.

48 $200,000: Aviation Week, 14/21 December 1992, p. 71.

49 AIAA Papers 99-4818, p. 2; 2002-5251, pp. 1-2 (length and range, p. 2).

50 AIAA Paper 99-4818 (quote, p. 1).

51 Ibid.

52 Miller, X-Planes, p. 389. Contract awards: AIAA Paper 2001-1910, p. 791.

53 AIAA Paper 99-4818. 13 feet per second: AIAA Paper 2001-1802, p. 257.

54 AIAA Paper 2002-5251, p. 3; Miller, X-Planes, pp. 391-92 (2,800 pounds, p. 392).

55 Journal of Spacecraft and Rockets, November-December 2001, pp. 846-47. 

56 Quote: AIAA Paper 98-1532, p. 4.

57 AIAA Papers 98-1532 (numbers, p. 4), 2001-1809.

58 AIAA Paper 2001-1809, pp. 292-93; Vince Rausch: Aviation Week, 9 June 1997, p. 32.

59 Journal of Spacecraft and Rockets, November-December 2001, pp. 829-30.

60 Journal of Spacecraft and Rockets, November-December 2001, pp. 828-35.

61 Journal of Spacecraft and Rockets, November-December 2001, pp. 803-10.

62 Journal of Spacecraft and Rockets, November-December 2001, pp. 853-64 (300 runs, table 1).

63 Journal of Spacecraft and Rockets, November-December 2001, p. 806.

64 AIAA Paper 2001-1910, p. 793. Mission plan: Aviation Week, 11 June 2001, p. 50. Quote: 

author interview, Lawrence Huebner, 24 July 2002. 

65 Aviation Week, 11 June 2001, pp. 50-51.

66 Air & Space, November 2002, pp. 74-81; AIAA Paper 2003-7029.

67 Quote: Air & Space, November 2002, p. 81.

68 Internet postings: Google, “Hyshot Australia.” Quote: Air & Space, November 2002, p. 79.

69 Aviation Week, 28 July 2003, pp. 37-38.

70 Aviation Week, 5 April 2004, pp. 28-29.

71 Aviation Week, 22 November 2004, pp. 24-26.

72 Aviation Week, 5 March 2001, p. 66.

73 Journal of Spacecraft and Rockets, November-December 2001, p. 838.

74 AIAA Paper 2002-5127.

75 Dongarra, “Performance.”

76 Science: 16 January 1981, pp. 268-69; 29 November 2002, p. 1713. Also http://www.top500.

org/list/1994/11, http://www.top500.org/list/1996/11

77 AIAA Paper 83-0037, table 1.

78 Author interview, Christian Stemmer, 22 April 2003. 

79 Author interview, William Cabot, 23 April 2003. 

80 Pope, Turbulent, pp. 182-183.

81 Author interview, Christian Stemmer, 23 April 2003. 

82 Deardorff: Journal of Fluid Mechanics, 13 April 1970, pp. 453-80; Geophysical Fluid Dynamics, 

November 1970, pp. 377-410.

83 AIAA Journal, December 1979, pp. 1300-1302.

84 Heppenheimer, “Tractable,” pp. 36-37 (includes quote).

85 Jansen: “Preliminary”; “Large-Eddy.”

86 Mahesh et al., “Large-Eddy.”

87 Author interviews, Parviz Moin and Heinz Pitsch, 22 April 2003. 

88 Scientific American, May 1968, pp. 59-78.

89 Heppenheimer, “Voyager”; 1989 Yearbook of Science and the Future (Encyclopedia Britannica), 

pp. 142-59.

90 Time, 29 November 2004, pp. 63-67; Air & Space, March 2006, p. 25.

91 Malakoff, Science, 9 May 2003, pp. 888-889.

92 Brooks, “Airship”; Botting, The Giant Airships.

93 Quotes: Hallion, Hypersonic, p. 98-iii; Sänger: Spaceflight, May 1973, p. 166.

“Advanced Technology Program: Technical Development Plan for Aerothermodynamic/elastic Structural Systems Environmental Tests (ASSET).” Air Force Systems Command, 9 September 1963. 

Air Force Studies Board. “Hypersonic Technology for Military Application.” Washington, DC: 

National Academies Press, 1989.

Allen, Thomas B., and Norman Polmar. Code-Name Downfall. New York: Simon & Schuster, 1995.

Anderson, John. A History of Aerodynamics. New York: Cambridge University Press, 1997.

Anspacher, William B., Betty H. Gay, Donald E. Marlowe, Paul B. Morgan, and Samuel J. Raff. The 

Legacy of the White Oak Laboratory. Dahlgren, VA: Naval Surface Warfare Center, 2000.

Augenstein, Bruno. “Rand and North American Aviation’s Aerophysics Laboratory: An Early 

Interaction in Missiles and Space.” Paper IAA-98-IAA.2.2.06 presented at the 49th Annual 

International Astronautical Federation, Melbourne, Australia, 1998.

Bagwell, Margaret. “History of the Bomarc Weapon System 1953-1957.” Wright-Patterson AFB, 

OH: Air Materiel Command, Office of Information Services, February 1959.

Baucom, Donald R. The Origins of SDI, 1944-1983. Lawrence, KS: University Press of Kansas, 1992.

Becker, John, Norris F. Dow, Maxime A. Faget, Thomas A. Toll, and J.B. Whitten. “Research 

Airplane Study.” NACA-Langley, April 1954.

Becker, Paul R. “Leading-Edge Structural Material System of the Space Shuttle.” American Ceramic 

Society Bulletin 60, no. 11 (1981): 1210-1214.

Bond, Aleck C. “Big Joe.” Project Mercury Technical History, 27 June 1963.

Botting, Douglas. The Giant Airships. Alexandria, VA: Time-Life Books, 1980.

Boyne, Walter J. The Messerschmitt Me 262: Arrow to the Future. Washington, DC: Smithsonian 

Institution Press, 1980.

Broad, William. Teller’s War. New York: Simon & Schuster, 1992.

Brooks, Peter W. “Why the Airship Failed.” Aeronautical Journal 79 (1975): 439-449.

Brown, C. E., W. J. O’Sullivan, and C. H. Zimmerman. “A Study of the Problems Related to High 

Speed, High Altitude Flight.” NACA-Langley, 1953.

Chapman, John L. Atlas: The Story of a Missile. New York: Harper, 1960.

Clarke, Arthur C. Profiles of the Future. New York: Bantam, 1964.

289

Bibliography
290

Facing the Heat Barrier: A History of Hypersonics

291

Bibliography

“Evaluation Report on X-15 Research Aircraft Design Competition.” In RG 255, National Archives, 

Philadelphia, 5 August 1955.

Faget, Maxime, and Milton Silveira. “Fundamental Design Considerations for an Earth-Surface to 

Orbit Shuttle.” Paper presented at the 21st Annual International Astronautical Congress, Konstanz, 

West Germany, 4-10 October 1970.

Fahrney, Rear Admiral D. S. “The History of Pilotless Aircraft and Guided Missiles.” Washington, 

DC: Archives, Federation of American Scientists, undated.

Ferri, Antonio. “Possible Directions of Future Research in Air-Breathing Engines.” In Combustion 

and Propulsion, Proceedings of the Fourth AGARD Colloquium, 1960. New York: Pergamon Press, 

1961.

———. “Statement Regarding the Military Service of Antonio Ferri.” Typescript in NASA-Langley 

Archives.

Ferri, Antonio, and Herbert Fox. “Analysis of Fluid Dynamics of Supersonic Combustion Process 

Controlled by Mixing.” In Twelfth Symposium on Combustion. Pittsburgh, PA: Combustion Institute, 

1968, 1105-1113.

Fitzgerald, Frances. Way Out There in the Blue. New York: Simon & Schuster, 2000.

Fletcher, Edward A. “Early Supersonic Combustion Studies at NACA and NASA.” In Eleventh 

Symposium on Combustion. Pittsburgh, PA: Combustion Institute, 1967, 729-737.

Geiger, Clarence J. “Termination of the X-20A Dyna-Soar.” In History of Aeronautical Systems 

Division, July-December 1963, Volume III. Air Force Systems Command: Historical Division, 

Report 64 ASE-39.

Gibson, James N. The Navaho Missile Project. Atglen, PA: Schiffer Publishing, 1996.

Greer, Kenneth E. “Corona.” In Corona, by Kevin C. Ruffner, ed., 3-39. Langley, VA: CIA, Center 

for the Study of Intelligence, 1995. 

Grimwood, James M., and Frances Strowd. “History of the Jupiter Missile System.” Redstone 

Arsenal, AL: U.S. Army Ordnance Missile Command, 27 July 1962.

Gunston, Bill. Fighters of the Fifties. Osceola, WI: Specialty Press, 1981.

Hallion, Richard P. “American Rocket Aircraft: Precursors to Manned Flight Beyond the 

Atmosphere.” In History of Rocketry and Astronautics. Edited by Kristan R. Lattu. AAS History Series, 

Volume 8. San Diego: Univelt, 1989, 283-312.

———, ed. The Hypersonic Revolution. Bolling AFB, DC: Air Force History and Museums Program, 

1998.

Hansen, Harry, ed. The World Almanac and Book of Facts for 1956. New York: New York WorldTelegram Corporation, 1956.

Harshman, D. L. “Design and Test of a Mach 7-8 Supersonic Combustion Ramjet Engine.” Paper 

“Columbia Accident Investigation Report.” NASA, August 2003.

Cornett, Lloyd H. “History of the Air Force Missile Development Center 1 January–30 June 1964.” 

Holloman AFB, NM. Report AFMDC 64-5791, 1964.

———. “An Overview of ADC Weapons 1946-1972.” USAF Aerospace Defense Command, April 

1973.

Crickmore, Paul F. Lockheed SR-71 Blackbird. London: Osprey Publishing, 1986.

Day, Dwayne A., John M. Logsdon, and Brian Latell. Eye in the Sky. Washington, DC: Smithsonian 

Institution Press, 1998.

Day, Leroy. “NASA Space Shuttle Task Group Report, Volume II: Desired System Characteristics.” 

Washington, DC: Space Shuttle Task Group, NASA, 12 June 1969.

Deardorff, J. W. “A Numerical Study of Three-Dimensional Turbulent Channel Flows at Large 

Reynolds Numbers.” Journal of Fluid Mechanics 41 (1970): 453-480.

———. “A Three-Dimensional Numerical Investigation of the Idealized Planetary Boundary Layer.” 

Geophysical Fluid Dynamics 1 (1970): 377-410.

“Development of the Bomarc Guided Missile 1950-53.” U.S. Air Force. Archives, Maxwell AFB, 

Montgomery, Alabama.

“Development of the Navaho Guided Missile 1945-1953.” U.S. Air Force. Archives, Maxwell AFB, 

Montgomery, Alabama.

“Development of the SM-64 Navaho Missile 1954-1958.” Wright-Patterson AFB, OH: Historical 

Division, ARDC, January 1961.

Dongarra, Jack J. “Performance of Various Computers Using Standard Linear Equations Software 

in a Fortran Environment.” Argonne, IL: Technical Memorandum 23, Mathematics and Computer 

Science Division, Argonne National Laboratory, 30 September 1988.

Dornberger, Walter. V-2. New York: Bantam Books, 1979.

Drake, Hubert M., and L. Robert Carman. “A Suggestion of Means for Flight Research at 

Hypersonic Velocities and High Altitudes.” In RG 255, National Archives, Philadelphia: NACALangley, 26 May 1952.

Emme, Eugene. The History of Rocket Technology. Detroit, MI: Wayne State University Press, 1964.

Erb, R. Bryan, and Emily W. Stephens. “Project Mercury: An Analysis of Mercury Heat-Shield 

Performance During Entry.” NASA Project Mercury Working Paper No. 193. Langley Field, VA: 

NASA Space Task Group, 21 June 1961.

Escher, William J. D. “Cryogenic Hydrogen-Induced Air Liquefaction Technologies for CombinedCycle Propulsion Applications.” European RBCC Workshop, TNO Prins Maurits Laboratory, Delft, 

Netherlands, 6-9 November 1995.
292

Facing the Heat Barrier: A History of Hypersonics

293

Bibliography

Korb, L. J., C.A. Morant, R.M. Calland and C.S. Thatcher. “The Shuttle Orbiter Thermal Protection 

System.” American Ceramic Society Bulletin 60 (1981): 1188-1193.

Krieth, Frank. Principles of Heat Transfer. Scranton, PA: International Textbook Company, 1964.

Kuethe, Arnold M., and Chuen-Yen Chow. Foundations of Aerodynamics. New York: John Wiley, 

1986.

Lay, Joachim E. Thermodynamics. Columbus, OH: C. E. Merrill, 1963.

Lee, John F., Francis W. Sears, and Donald L. Turcotte. Statistical Thermodynamics. Reading, MA: 

Addison-Wesley, 1963.

Ley, Willy. Rockets, Missiles, and Man in Space. New York: Signet, 1968.

Lin, C. C.. Statistical Theories of Turbulence. Princeton, NJ: Princeton University Press, 1961.

Lukasiewicz, J. Experimental Methods of Hypersonics. New York: Marcel Dekker, 1973.

MacKenzie, Donald. Inventing Accuracy. Cambridge, MA: MIT Press, 1990.

Mackley, Ernest A. “Historical Aspects of NASA Scramjet Technology Development.” NASALangley, undated.

Mahesh, Krishnan, et al. “Large-Eddy Simulation of Gas Turbine Combustors.” Annual Research 

Briefs. Stanford, CA: Center for Turbulence Research, Stanford University, 2001, 3-17.

Malia, Martin E. The Soviet Tragedy. New York: Free Press, 1994.

Manchester, William. American Caesar. New York: Dell, 1978.

Martin, James A. “History of NACA-Proposed High-Mach-Number, High-Altitude Research 

Airplane.” 16 December 1954.

McDougall, Walter A. The Heavens and the Earth. New York: Basic Books, 1985.

McPhee, John. The Curve of Binding Energy. New York: Ballantine Books, 1975.

Miller, Jay. The X-Planes, X-1 to X-45. North Branch, MN: Specialty Press, 2001.

Miller, Ron, and Frederick C. Durant. Worlds Beyond: The Art of Chesley Bonestell. Norfolk, VA: 

Donning, 1983.

Murray, Russ. Lee Atwood...Dean of Aerospace, Rockwell International Corporation, 1980.

Naval Research Laboratory. “Upper Atmosphere Research Report No. 1.” Report R-2955. 1 October 

1946.

Neufeld, Jacob. The Development of Ballistic Missiles in the United States Air Force, 1945-1960. 

Washington, DC: Office of Air Force History, USAF, 1990.

Neufeld, Michael J. The Rocket and the Reich. New York: Free Press, 1995.

presented at the American Institute of Aeronautics and Astronautics Propulsion Specialist Meeting, 

Washington, DC, 17-21 July 1967.

Heppenheimer, T. A. Countdown: A History of Space Flight. New York: John Wiley, 1997.

———. Development of the Space Shuttle, 1972-1981. Washington, DC: Smithsonian Institution 

Press, 2002.

———. First Flight. New York: John Wiley, 2003.

———. Hypersonic Technologies and the National Aerospace Plane. Arlington, VA: Pasha Publications, 

1990.

———. The Man-Made Sun: The Quest for Fusion Power. Boston: Little, Brown, 1984.

———. “Some Tractable Mathematics for Some Intractable Physics.” Mosaic. Vol. 22, No. 1, Spring 

1991, pp. 28-39.

———. Turbulent Skies: The History of Commercial Aviation. New York: John Wiley, 1995.

Hermann, R. “The Supersonic Wind Tunnel Installations at Peenemunde and Kochel and Their 

Contributions to the Aerodynamics of Rocket-Powered Vehicles.” In Space: Mankind’s Fourth 

Environment. Edited by L. G. Napolitano. New York: Pergamon Press, 1982, 435-446.

“History of the Arnold Engineering Development Center.” AFSC Historical Publications, U.S. Air 

Force.

Hunley, J. D. “The Significance of the X-15.” NASA-Dryden History Office, 1999.

Hunter, Maxwell W. “The Origins of the Shuttle (According to Hunter).” Sunnyvale, CA: Lockheed, 

September 1972. Reprinted in part in Earth/Space News, November 1976, 5-7.

Huxley, Aldous, Brave New World. New York: HarperCollins, 1998.

Jane’s All the World’s Aircraft. New York: McGraw-Hill. Annual editions; began in 1930.

Jansen, Kenneth. “Preliminary Large-Eddy Simulations of Flow Around a NACA 4412 Airfoil 

Using Unstructured Grids.” Annual Research Briefs. Stanford, CA: Center for Turbulence Research, 

Stanford University, 1995, 61-72.

———. “Large-Eddy Simulation of Flow Around a NACA 4412 Airfoil Using Unstructured Grids.” 

Annual Research Briefs. Stanford, CA: Center for Turbulence Research, Stanford University, 1996, 

225-232.

Jenkins, Dennis. Space Shuttle. Stillwater, MN: Voyageur Press, 2001.

Kantrowitz, Arthur. “Shock Tubes for High Temperature Gas Kinetics.” Avco-Everett Research 

Laboratory, Research Report 141, October 1962.

Kaufman, Louis, Barbara Fitzgerald, and Tom Sewell. Moe Berg: Athlete, Scholar, Spy. Boston: Little, 

Brown, 1974.
294

Facing the Heat Barrier: A History of Hypersonics

295

Bibliography

Ruffner, Kevin C. “Corona: America’s First Satellite Program.” Washington: Central Intelligence 

Agency, 1995.

Schlaifer, Robert, and S. D. Heron. Development of Aircraft Engines and Fuels. Boston: Harvard 

University, 1950.

Schnyer, A. Dan and R. G. Voss. “Review of Orbital Transportation Concepts—Low Cost 

Operations.” Washington, DC: NASA-OMSF, 18 December 1968.

Schramm, Wilson. “HRSI and LRSI—The Early Years.” American Ceramic Society Bulletin 60 

(1981): 1194-1195.

Schweikart, Larry. The Quest for the Orbital Jet. Volume III of The Hypersonic Revolution. Edited by 

Richard P. Hallion. Bolling AFB, DC: Air Force History and Museums Program, 1998.

Shapiro, Ascher. The Dynamics and Thermodynamics of Compressible Fluid Flow. New York: Ronald 

Press. Two volumes, 1953 and 1954.

Sietzen, Frank. “Wall Street Rejects VentureStar.” Washington, DC: News release, 21 May 1999.

Speer, Albert. Inside the Third Reich. New York: Macmillan, 1970.

“Standard Missile Characteristics: IM-99A Bomarc.” U.S. Air Force, 8 May 1958.

“Standard Missile Characteristics: XSM-64 Navaho.” U.S. Air Force, 1 November 1956.

Sunday, Terry L., and John R. London. “The X-20 Space Plane: Past Innovation, Future Vision.” In 

History of Rocketry and Astronautics. Edited by John Becklake. AAS History Series, Volume 17. San 

Diego: Univelt, 1995, 253-284.

Sutton, George P. Rocket Propulsion Elements. New York: John Wiley, 1986.

Thompson, Tina D. ed. TRW Space Log. Redondo Beach, CA: TRW Space and Electronics Group. 

Annual editions.

Tsien, Hsue-shen. “Similarity Laws of Hypersonic Flows.” Journal of Mathematics and Physics 25 

(1946): 247-251.

Von Karman, Theodore with Lee Edson. The Wind and Beyond. Boston: Little, Brown, 1967.

Walker, Martin. The Cold War. New York: Henry Holt, 1994.

Waltrup, Paul J., Griffin Y. Anderson, and Frank S. Stull. “Supersonic Combustion Ramjet 

(Scramjet) Engine Development in the United States.” Paper presented at the Third International 

Symposium on Air Breathing Engines, Munich, 7-12 March 1976, pp. 42-1 to 42-27.

Wattendorf, Frank L. “German Wind Tunnels.” AAF Technical Report 5268, Air Technical Service 

Command, 20 August 1945.

Wegener, Peter P. The Peenemunde Wind Tunnels. New Haven, CT: Yale University Press, 1996.

Ordway, Frederick and Mitchell R. Sharpe. The Rocket Team. New York: Thomas Y. Crowell, 1979.

Pace, Scott. “Engineering Design and Political Choice: The Space Shuttle 1969-1972.” Masters 

thesis, MIT, May 1982.

Pedigree of Champions: Boeing Since 1916. Seattle: Boeing, 1985.

Peterson, Lee L. “Evaluation Report on X-7A.” Report AFMDC ADJ 57-8184. Holloman AFB, 

NM, 3 October 1957.

Pfeifer, John L. “The BOMARC Weapon System.” Bomarc Service News Article. Seattle, WA: 

Boeing archives, 4 March 1959.

Pope, Stephen B. . Turbulent Flows. New York: Cambridge University Press, 2000.

Preston, Richard. American Steel. New York: Prentice Hall, 1991.

“Re-Entry Studies,” Volume 1. Redstone Arsenal, AL: Army Ballistic Missile Agency, 25 November 

1958.

“Re-Entry Test Vehicle X-17.” In “History, Air Force Missile Test Center, 1 July – 31 December 

1957.” Volume IV, “Supporting Documents,” Appendix F. Air Force archives, Maxwell AFB, 

Montgomery, AL.

“Research-Airplane-Committee Report on Conference on the Progress of the X-15 Project.” Langley 

Field, VA: Langley Aeronautical Laboratory, October 25-26, 1956.

“Research-Airplane-Committee Report on Conference on the Progress of the X-15 Project.” Los 

Angeles: IAS Building, July 28-30, 1958.

Rhodes, Richard. Dark Sun: The Making of the Hydrogen Bomb. New York: Simon & Schuster, 1995.

———. The Making of the Atomic Bomb. New York: Simon & Schuster, 1986.

Richelson, Jeffrey. American Espionage and the Soviet Target. New York: William Morrow, 1987.

———. America’s Secret Eyes in Space. New York: Harper & Row, 1990.

Riddell, F. R. and J. D. Teare. “The Differences Between Satellite and Ballistic Missile Re-Entry 

Problems.” In Vistas in Astronautics, Vol. 2. Edited by Morton Alperin and Hollingsworth F. Gregory. 

New York: Pergamon Press, 1959, 174-190.

Rising, Jerry J. “The Lockheed Martin/NASA X-33 Program: A Stepping Stone to VentureStar.” 

Washington, DC: Lockheed Martin Technology Symposium, 3 November 1998.

Ritchie, William A. “Evaluation Report on X-7A (System 601B).” Report AFMDC DAS 58-8129. 

Holloman AFB, NM, January 1959. Montgomery, AL: Archives, Maxwell AFB.

Rose, P. H. “Physical Gas Dynamics Research at the Avco Research Laboratory.” AGARD Report 

145, July 1957.
296

Facing the Heat Barrier: A History of Hypersonics

297

Bibliography

70-1249. P. P. Antonatos, A. C. Draper, and R. D. Neumann. “Aero Thermodynamic and 

Configuration Development.” October 1970.

71-443. F. M. Anthony, R. R. Fisher, and R. G. Helenbrook. “Selection of Space Shuttle Thermal 

Protection Systems.” April 1971.

71-804. J. F. Yardley. “Fully Reusable Shuttle.” July 1971.

71-805. Bastian Hello. “Fully Reusable Shuttle.” July 1971.

75-1212. Carl A. Trexler. “Inlet Performance of the Integrated Langley Scramjet Module (Mach 2.3 

to 7.6).” September 1975.

78-257. B. S. Baldwin and H. Lomax. “Thin Layer Approximation and Algebraic Model for 

Separated Turbulent Flows.” January 1978.

78-485. M. J. Suppanz and C. J. Schroeder. “Space Shuttle Orbiter Thermal Protection 

Development and Verification Test Program.” April 1978.

79-0219. L. L. Cronvich. “Aerodynamic Development of Fleet Guided Missiles in Navy’s Bumblebee 

Program.” January 1979.

79-7045. Robert W. Guy and Ernest A. Macklay. “Initial Wind Tunnel Tests at Mach 4 and 7 of a 

Hydrogen-Burning, Airframe-Integrated Scramjet.” April 1979.

83-0037. P. Kutler. “A Perspective of Theoretical and Applied Computational Fluid Dynamics 

(Invited).” January 1983.

83-1693. T. J. Coakley. “Turbulence Modeling Methods for the Compressible Navier-Stokes 

Equations.” July 1983.

84-0387. F. S. Billig, M. E. White, and D. M. Van Wie. “Application of CAE and CFD Techniques 

to a Complete Tactical Missile Design.” January 1984.

84-2414. S. A. Tremaine and Jerry B. Arnett. “Transatmospheric Vehicles—A Challenge for the Next 

Century.” November 1984.

85-1509. J. S. Shang and S. J. Scherr. “Navier-Stokes Solution of the Flow Field Around a Complete 

Aircraft.” July 1985.

85-1652. B. Lakshminarayana. “Turbulence Modelling for Complex Flows (Invited Paper).” July 

1985.

86-0159. G. Burton Northam and G. Y. Anderson. “Supersonic Combustion Ramjet Research at 

Langley.” January 1986.

86-1680. W. J. D. Escher, R. R. Teeter, and E. E. Rice. “Airbreathing and Rocket Propulsion 

Synergism: Enabling Measures for Tomorrow’s Orbital Transports.” June 1986.

88-0230. R. J. Johnston and W. C. Sawyer. “An Historical Perspective on Hypersonic Aerodynamic 

Research at the Langley Research Center.” January 1988.

Wolfe, Tom. The Right Stuff. New York: Bantam Books, 1980.

Yeager, Chuck, and Leo Janos. Yeager, an Autobiography. New York: Bantam Books, 1985.

Books, Papers, and Reports by Source

Acta Astronautica

Dale D. Myers. “The Navaho Cruise Missile—A Burst of Technology.” 26 (1992): 741-748.

Delma C. Freeman and Theodore A. Talay. “Single-Stage-to-Orbit—Meeting the Challenge.” 38 

(1996): 323-331.

Aerospace America

Adrian M. Messner. “Solid Rockets.” December 1979, pp. 48-49.

___, “Solid Rockets.” December 1980, p. 73.

“Hypersonic Flight Revisited.” June 1984, p. 1.

Patrick L. Johnston, Allen B. Whitehead, and Gary T. Chapman. “Fitting Aerodynamics and 

Propulsion Into the Puzzle.” September 1987, pp. 32-37.

Douglas L. Dwoyer, Paul Kutler, and Louis A. Povinelli. “Retooling CFD for Hypersonic Aircraft.” 

October 1987, pp. 32-35.

Art Hanley. “Titan IV: Latest in a Family of Giants.” July 1991, pp. 34-38.

Alan S. Brown. “NASA Funds Titanium Composite Plant.” August 1992, pp. 66-67.

AIAA (American Institute of Aeronautics and Astronautics)

63075. Perry V. Row and Jack Fischel. “Operational Flight-Test Experience with the X-15 Airplane.” 

March 1963.

64-17. Euclid C. Holleman and Elmor J. Adkins. “Contributions of the X-15 Program to Lifting 

Entry Technology.” January 1964.

67-166. B. J. Griffith. “Comparison of Data from the Gemini Flights and AEDC-VKF Wind 

Tunnels.” January 1967.

67-838. Robert L. Perry. “The Ballistic Missile Decisions.” October 1967.

68-1142. R. Bryan Erb, D.H. Greenshields, L.T. Chauvin, J.E. Pavlosky and C.L. Statham. “Apollo 

Thermal-Protection System Development.” December 1968.

69-98. Eugene P. Bartlett, Larry W. Anderson, and Donald M. Curry. “An Evaluation of Ablation 

Mechanisms for the Apollo Heat Shield Material.” January 1969.

70-715. A. J. Metzler and T. W. Mertz. “Large Scale Supersonic Combustor Testing at Conditions 

Simulating Mach 8 Flight.” June 1970.
298

Facing the Heat Barrier: A History of Hypersonics

299

Bibliography

93-2328. William O. T. Peschke. “Approach to In Situ Analysis of Scramjet Combustor Behavior.” 

June 1993.

93-2329. Frederick S. Billig. “SCRAM—A Supersonic Combustion Ramjet Missile.” June 1993.

93-2564. K. E. Dayton. “National AeroSpace Plane Integrated Fuselage/Cryotank Risk Reduction 

Program.” June 1993.

93-5101. E. Kraft and G. Chapman. “A Critical Review of the Integration of Computations, 

Ground Tests, and Flight Test for the Development of Hypersonic Vehicles.” December 1993.

95-3777. A. Elias, D. Hays, and J. Kennedy. “Pioneering Industry/Government Partnerships: X-34.” 

September 1995.

95-6031. R. L. Chase and M. 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.” April 1995.

95-6055. R. Voland and K. Rock. “NASP Concept Demonstration Engine and Subscale Parametric 

Engine Tests.” April 1995.

95-6131. T. Ronald. “Status and Applications of Materials Developed for NASP.” April 1995.

96-4563. S. Cook. “X-33 Reusable Launch Vehicle Structural Technologies.” November 1996.

98-1532. R. T. Voland, K.E. Rock, L.D. Huebner, D.W. Witte, K.E. Fisher and C.R. McClinton. 

“Hyper-X Engine Design and Ground Test Program.” April 1998.

98-2506. R. C. Rogers, D. P. Capriotti, and R. W. Guy. “Experimental Supersonic Combustion 

Research at NASA Langley.” June 1998.

98-3516. R. B. Sullivan and B. Winters. “X-34 Program Overview.” July 1998.

99-4818. David E. Reubush. “Hyper-X Stage Separation—Background and Status.” November 

1999.

99-4848. R. T. Voland, A.H. Auslander, M.K. Smart, A.S. Rousakov, V.L. Semenov and V. 

Kapchenov. “CIAM/NASA Mach 6.5 Scramjet Flight and Ground Test.” November 1999.

99-4896. D. E. Pryor, E. H. Hyde, and W. J. D. Escher. “Development of a 12-Thrust Chamber 

Kerosene/Oxygen Primary Rocket Subsystem for an Early (1964) Air-Augmented Rocket Ground 

Test System.” November 1999.

2001-1802. David E. Reubush, John G. Martin, Jeffrey S. Robinson, David M. Bose and Brian K. 

Strovers. “Hyper-X Stage Separation—Simulation Development and Results.” April 2001.

2001-1809. Lawrence D. Huebner, Kenneth E. Rock, Edward G. Ruf, David W. Witte and Earl H. 

Andrews. “Hyper-X Flight Engine Ground Testing for X-43 Flight Risk Reduction.” April 2001.

2001-1910. Charles R. McClinton, David R. Rausch, Joel Sitz and Paul Reukauf. “Hyper-X 

Program Status.” April 2001.

89-0008. J. Der. “Improved Methods of Characterizing Ejector Pumping Performance.” January 

1989.

89-5010. T. Ronald. “Materials Challenges for NASP.” July 1989.

89-5014. N. Hannum and F. Berkopec. “Fueling the National Aerospace Plane with Slush 

Hydrogen.” July 1989.

90-1480. P. Bradshaw. “Progress in Turbulence Research.” June 1990.

90-5206. N. Newman. “NASP Materials and Structures Augmentation Program (NMASAP) 

Overview.” October 1990.

90-5207. J. Sorensen. “Titanium Matrix Composites—NASP Materials and Structures 

Augmentation Program.” October 1990.

90-5231. F. Owen. “Transition and Turbulence Measurements in Hypersonic Flows.” October 1990.

90-5232. M. Malik, T. Zang, and D. Bushnell. “Boundary Layer Transition in Supersonic Flows.” 

October 1990.

90-5233. T. Elias and E. Eiswirth. “Stability Studies of Planar Transition in Supersonic Flows.” 

October 1990.

90-5247. T. Gatski, S. Sarkar, C.Speziale, L. Balakrishnan, R. Abid and E. Anderson. “Assessment 

and Application of Reynolds-Stress Closure Models to High-Speed Compressible Flows.” October 

1990.

91-1944. Jerry Rosevear. “Liquid Air Cycle Engines.” June 1991.

91-2395. G. A. Sullins, D. A. Carpenter, M.W. Thompson, F.T. Kwok and L.A. Mattes. “A 

Demonstration of Mode Transition in a Scramjet Combustor.” June 1991.

91-5008. B. Waldman and F. Harsha. “The First Year of Teaming: A Progress Report.” December 

1991.

92-3499. L. Q. Maurice, J. L. Leingang, and L. R. Carreiro. “Airbreathing Space Boosters Using InFlight Oxidizer Collection.” July 1992.

92-5023. M. Togawa, T. Aoki, and Y. Kaneko. “On LACE Research.” December 1992.

92-5049. M. Malik and P. Balakumar. “Instability and Transition in Three-Dimensional Supersonic 

Boundary Layers.” December 1992.

93-0200. P. G. Huang and T. J. Coakley. “Turbulence Modeling for Complex Hypersonic Flows.” 

January 1993.

93-0309. William H. Dana. “The X-15 Airplane—Lessons Learned.” January 1993.

93-2323. Earl H. Andrews and Ernest A. Mackley. “Review of NASA’s Hypersonic Research Engine 

Project.” June 1993.
300

Facing the Heat Barrier: A History of Hypersonics

301

Bibliography

John V. Becker. “The X-15 Project.” February 1964, pp. 52-61.

Thomas A. Toll and Jack Fischel. “The X-15 Project: Results and New Research.” March 1964, pp. 

20-28.

Antonio Ferri. “Supersonic Combustion Progress.” August 1964, pp. 32-37.

Max Faget, “Space Shuttle: A New Configuration.” January 1970, pp. 52-61.

Alfred C. Draper, Melvin L. Buck, and William H. Goesch. “A Delta Shuttle Orbiter.” January 1971, 

pp. 26-35.

John Sloop. “Looking for the Sweet Combination.” October 1972, pp. 52-57.

George Strouhal and Douglas J. Tillian. “Testing the Shuttle Heat-Protection Armor.” January 1976, 

pp. 57-65.

Robert A. Jones and Paul W. Huber. “Toward Scramjet Aircraft.” February 1978, pp. 38-48.

Astronautics and Aerospace Engineering

Perry V. Row and Jack Fischel. “X-15 Flight-Test Experience.” 1 (1963): 25-32.

Aviation Week (includes Aviation Week & Space Technology)

“Ford Beats Schedule on J57 Delivery.” 3 May 1954: 44. 

“How Mercury Capsule Design Evolved.” 21 September 1959, pp. 52-59.

“Rolls-Royce RCo. 12, Mark 509.” 2 May 1960, p. 145.

Larry Booda. “USAF Plans Radical New Space Plane.” 31 October 1960, p. 26.

“Fourth ASSET Glider Gathers Flutter Data.” 2 November 1964, p. 25.

William J. Normyle. “Manned Flight Tests to Seek Lifting-Body Technology.” 16 May 1966, pp. 64-

75.

B. K. Thomas. “USAF Nears Manned Lifting Body Tests.” 10 July 1967, pp. 99-101.

“Heading Shift Cited in X-15 Loss.” 12 August 1968, pp. 104-117.

“X-15 Crash Spurs Adaptive Control Study.” 26 August 1968, pp. 85-105.

Michael L. Yaffee. “Program Changes Boost Grumman Shuttle.” 12 July 1971, pp. 36-39.

“Sortie Module May Cut Experiment Cost.” 17 January 1972, p. 17.

William S. Hieronymus. “Two Reusable Materials Studied for Orbiter Thermal Protection.” 27 

March 1972, p. 48.

Benjamin L. Elson. “New Unit to Test Shuttle Thermal Guard.” 31 March 1975, pp. 52-53.

2002-5127. M. E. M. Stewart, A. Suresh, M.S. Liou, A.K. Owen and D.G. Messitt. 

“Multidisciplinary Analysis of a Hypersonic Engine.” October 2002.

2002-5251. George F. Orton. “Air-Breathing Hypersonics Research at Boeing Phantom Works.” 

October 2002.

2003-7029. Russell R. Boyce, Sullivan Gerard, and Allan Paull. “The HyShot Scramjet Flight 

Experiment—Flight Data and CFD Calculations Compared.” December 2003.

AIAA Journal

Henry Hidalgo and Leo P. Kadanoff. “Comparison Between Theory and Flight Ablation Data.” 1 

(1963): 41-45.

Dean R. Chapman. “Computational Aerodynamics Development and Outlook.” 17 (1979): 1293-

1313.

Air & Space (Smithsonian)

Luba Vangelova. “Outback Scramjet.” November 2002, pp. 74-81.

Craig Mellow. “Go Ballistic.” March 2006, pp. 20-25.

Air Enthusiast

Richard A. DeMeis. “The Trisonic Titanium Republic.” No. 7, July-September 1978, pp. 198-213.

Air Power History

Larry Schweikart. “Hypersonic Hopes: Planning for NASP, 1986-1991.” Spring 1994, pp. 36-48.

Roy Houchin. “Hypersonic Technology and Aerospace Doctrine.” Fall 1999, pp. 4-17.

T. A. Heppenheimer. “Origins of Hypersonic Propulsion: A Personal History.” Fall 2000, pp. 14-27.

American Heritage of Invention and Technology

James I. Killgore. “The Planes That Never Leave the Ground.” Winter 1989, pp. 56-63.

James E. Tomayko. “The Airplane as Computer Peripheral.” Winter 1992, pp. 19-24.

ARS (American Rocket Society)

Thomas R. Brogan. “Electric Arc Gas Heaters for Re-Entry Simulation and Space Propulsion.” ARS 

paper 724-58, November 1958.

Leo Steg. “Materials for Re-Entry Heat Protection of Satellites.” ARS Journal, September 1960, pp. 

815-822.

Astronautics & Aeronautics (continues as Aerospace America)

Melvin W. Root and George M. Fuller. “The Astro Concept.” January 1964, pp. 42-51.
302

Facing the Heat Barrier: A History of Hypersonics

303

Bibliography

Bruce A. Smith. “Elevon Failure Precedes Loss of First X-43A.” 11 June 2001, pp. 50-51.

Michael Dornheim. “X-43 to Fly in Fall.” 28 July 2003, pp. 36-37.

———. “A Breath of Fast Air.” 5 April 2004, pp. 28-29.

———. “But Now What?” 22 November 2004, pp. 24-26.

Bell Aircraft

D143-981-009. “MX-2276: Advanced Strategic Weapon System Progress Report No. 9.” 15 

February 1955.

D143-945-055. “Brass Bell Reconnaissance Aircraft Weapon System Summary Report.” 31 August 

1957.

Boeing Airplane Company

D-11532. “Final Summary Report: Development of USAF XIM-99A Weapon System January 

1951-March 1959.”

CASI (Center for Aerospace Information)

64X-17126. R. J. Bassett and R. M. Knox. “Preliminary Design Analysis of a 10,000 lb Thrust 

LACE Engine.” Marquardt: 5 November 1963.

72A-10764. 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, October 1971, pp. 453-464.

75N-75668. J. F. Foote. “Operating Experience and Performance Documentation of a LACE 

System.” Marquardt: 15 November 1963.

77A-35304. David H. Greenshields. “Orbiter Thermal Protection System Development.” Canaveral 

Council of Technical Societies: 14th Space Congress, Cocoa Beach, FL, April 1977, pp. 1-28 to 1-

42.

81A-44344. L. J. Korb and H. M. Clancy. “The Shuttle Thermal Protection System—A Material and 

Structural Overview.” SAMPE 26th National Symposium, Los Angeles, April 1981, pp. 232-249.

88N-13231. P. J. Waltrup. “Hypersonic Airbreathing Propulsion: Evolution and Opportunities.”

Comptes Rendus

Paul Vieille. “Sur les discontinuites produites par la detente brusque de gaz compreimes.” 129 

(1988): 1228-1230.

Convair (see also General Dynamics)

ZP-M-095. “Project Spaceplane Status Presentation.” April 1960.

“Reusable Space Launch Vehicle Systems Study.” Montgomery, AL: Archives, Maxwell AFB.

“Japan Explores Liquid Air Cycle Engine for Future Rocket Propulsion Needs.” 3 April 1989, p. 58.

Stanley W. Kandebo. “Lifting Body Design Is Key to Single-Stage-to-Orbit.” 29 October 1990, pp. 

36-37.

———. “NASP Team Narrows Its Options As First Design Cycle Nears Completion.” 1 April 1991, 

p. 80.

Michael A. Dornheim. “NASP Fuselage Model Undergoes Tests for Temperatures, Stresses of Flight.” 

3 February 1992, p. 52.

Stanley W. Kandebo. “Russians Want U.S. to Join Scramjet Tests.” 30 March 1992, pp. 18-20.

Michael A. Dornheim. “Mitsubishi Planning to Test Liquid Air Cycle Rocket Engine.” 11 May 

1992: 54.

Stanley W. Kandebo. “Franco-Russian Tests May Spur New Projects.” 14/21 December 1992, pp. 

70-73.

———. “Boeing 777 to Incorporate New Alloy Developed for NASP.” 3 May 1993, p. 36.

James R. Asker. “New Space Launchers a Tough Test for NASA.” 31 October 1994, pp. 22-23.

———. “Racing to Remake Space Economics.” 3 April 1995, pp. 44-45.

“X-34 to Be Acid Test for Space Commerce.” 3 April 1995, pp. 44-53.

Joseph C. Anselmo and James R. Asker. “NASA Suspends X-34 Booster.” 6 November 1995, p. 30.

“Endeavour Tests Metallic TPS.” 3 June 1996, p. 92.

Joseph C. Anselmo. “NASA Gives Orbital Second Shot at X-34.” 17 June 1996, p. 31.

Michael A. Dornheim. “Follow-on Plan Key to X-33 Win.” 8 July 1996, pp. 20-22.

Bruce A. Smith. “Hyper-X Undergoes Design Review.” 9 June 1997, p. 32.

James R. Asker. “NASA, Industry Hit Snags on X-33.” 30 June 1977, pp. 27-28.

Michael Dornheim. “X-33 Design Gets Go-Ahead.” 10 November 1997.

———. “X-33 Flight Slips Half-Year to December 1999.” 2 November 1998, pp. 26-27.

———. “Bonding Bugs Delay X-33’s First Flight.” 25 January 1999, pp. 46-47.

Joseph C. Anselmo. “A Mach 8 Vehicle on the RLV Frontier.” 26 April 1999, pp. 78-80.

Michael A. Dornheim. “Engineers Anticipated X-33 Tank Failure.” 15 November 1999, pp. 28-30.

Frank Morring. “NASA Kills X-33, X-34, Trims Space Station.” 5 March 2001, pp. 24-25.

Edward H. Phillips. “‘Mariah’ Team Eyes Hypersonic Tunnel.” 5 March 2001, p. 66.
304

Facing the Heat Barrier: A History of Hypersonics

305

Bibliography

AD-351239. “The Dynatech High-Temperature, Chemical Air Separator for ACES.” Dynatech 

Corporation, 15 February 1962.

AD-351581. “Republic Aerospaceplane. Volume 1.” Republic Aviation, 28 January 1964.

AD-351582. “Evaluation of P + W Turboramjet Engines in Republic Aerospaceplane. Volume II.” 

Republic Aviation, 28 January 1964.

AD-351583. “Evaluation of G.E. Turboramjet Engines in Republic Aerospaceplane. Volume III.” 

Republic Aviation, 28 January 1964.

AD-351584. “Evaluation of Marquardt SuperLACE Engines in Republic Aerospaceplane. Volume 

IV.” Republic Aviation, 28 January 1964.

AD-353898. L. C. Hoagland and J. R. Turner. “High-Temperature Air Enrichment Program.” 

Dynatech Corporation, 15 September 1964.

AD-357523. “ASSET ASV-3 Flight Test Report.” McDonnell Aircraft Corporation, 4 January 1965.

AD-359545. “Phase I Report on Recoverable Orbital Launch System (ROLS).” Holloman AFB, 

NM, 23 April 1965.

AD-359763. “Experimental and Theoretical Investigation of the Rocket Engine-Nozzle Ejector 

(RENE) Propulsion System.” Martin, April 1965.

AD-362539. T. E. Shaw, W.S. Mertz, N.M. Keough and I.M Musser. “Thermal Flight Test Summary 

Report for Mark 3 Mod 1 Re-Entry Vehicles.” General Electric, 2 August 1960.

AD-366546. “ASSET ASV-4 Flight Test Report.” McDonnell Aircraft Corporation, 25 June 1965.

AD-372320. R. L. Doebler and M. G. Hinton. “Current Status and Challenge of Airbreathing 

Propulsion for Launch Vehicle Systems.” Aerospace Corporation, 15 March 1966.

AD-381504. “Air Separator Test Program.” Union Carbide, October 1966.

AD-384302. R. J. Bassett and R. M. Knox. “Component Specifications for the MA116 Advanced 

LACE System.” Marquardt, 5 November 1963.

AD-386653. C. Edward Kepler and Arthur J. Karanian. “Hydrogen-Fueled Variable Geometry 

Scramjet.” United Aircraft Research Laboratories, January 1968.

AD-388239. M. L. Brown and R. L. Maxwell. “Scramjet Incremental Flight Test Program.” 

Marquardt, February 1968.

AD-393374. David J. McFarlin and C. Edward Kepler. “Mach 5 Test Results of Hydrogen-Fueled 

Variable-Geometry Scramjet.” United Aircraft Research Laboratories, 24 October 1968.

AD-441740. D. Cona. “Coatings for Refractory Metals.” Boeing, 27 December 1963.

AD-449685. Howard J. Middendorf. “Materials and Processes for X-20A (Dyna-Soar).” Air Force 

Systems Command, June 1964.

DTIC (Defense Technical Information Center)

AD-098980. Daniel E. Bloxsom. “Production of High-Temperature, Moderate-Pressure Gases by 

Means of Electrical Spark Discharge.” Arnold Engineering Development Center, November 1956.

AD-299774. James C. Sivells. “Aerodynamic Design and Calibration of the VKF 50-inch 

Hypersonic Wind Tunnels.” Arnold Engineering Development Center, March 1963.

AD-318628. F. Herr and R. L. Maxwell. “Hypersonic, Air-Breathing Engine Research Program.” 

Marquardt, 31 August 1960.

AD-320753. “Preliminary Flight Test Report. RVX-1, Flight No. 1. AVCO Vehicle No. 5. Missile 

No. 133 (Thor-Able Phase II).” General Electric, 29 April 1959.

AD-322199. “Hypersonic, Air-Breathing Engine Research Program.” Marquardt, 28 February 1961.

AD-328815. F. A. Costello and J. A. Segletes. “Development and Thermal Performance of the 

Discoverer Heat Shield.” General Electric, 1 March 1962.

AD-335212. J. F. Foote and R. J. Bassett. “Summary of Initial Tests of the Marquardt Model 

MA117 LACE Engine.” Marquardt, 15 February 1963.

AD-336169. J. R. Turner and L. C. Hoagland. “Study of High Temperature Chemical Separation of 

Air into O2

-N2 Components.” Dynatech Corporation, February 1963.

AD-342828. “Thor-Able Phase II, RVX-1, Flight-3, Re-entry Vehicle No. 172, Missile No. 132, 

AMR Test No. 15.” General Electric, 9 May 1959.

AD-342992. “Spaceplane Technical Status Report.” Convair, February 1961.

AD-345178. “Airframe Support of a Boilerplate Air Separator Air Enrichment Program.” General 

Dynamics, October 1963.

AD-346912. “Proceedings of 1962 X-20A (Dyna-Soar) Symposium. Volume III. Structures and 

Materials.” Aeronautical Systems Division, U.S. Air Force, March 1963.

AD-350952. R. J. Bassett and R. M. Knox. “Preliminary Design Analysis of a 10,000-lb Thrust 

LACE Engine.” Marquardt, 5 November 1963.

AD-351039. “Aerospaceplane Propulsion Integration Study. Volume 1.” Republic Aviation, 15 

February 1963.

AD-351040. “Aerospaceplane Propulsion Integration Study. Volume 3.” Republic Aviation, 15 

February 1963.

AD-351041. “Aerospaceplane Propulsion Integration Study. Volume 2.” Republic Aviation, 15 

February 1963.

AD-351207. “Feasibility Study of a High Capacity Distillation Separator for an Air Enrichment 

System.” Union Carbide, February 1964.
306

Facing the Heat Barrier: A History of Hypersonics

307

Bibliography

ADB-968993. “NASP Security Classification Guide.” Department of Defense and NASA, 30 

November 1990.

ADC-053100. “Proposed 730 Orbital Airplane.” Republic Aviation, November 22, 1960.

General Accounting Office

NSIAD-92-5. Nancy R. Kingsbury. “Aerospace Plane Technology: Research and Development 

Efforts in Japan and Australia.” 4 October 1992.

NSIAD-93-71. Frank C. Conahan. “National Aero-Space Plane: Restructuring Further Research and 

Development Efforts.” 3 December 1992.

NSIAD-99-176. “Space Transportation: Status of the X-33 Reusable Launch Vehicle Program.” 11 

August 1999.

General Dynamics

GD/C-DCJ 65-004. “Reusable Space Launch Vehicle Study.” 18 May 1965.

GDC-DCB-67-031. “Multipurpose Reusable Spacecraft Preliminary Design Effort.” 25 September 

1967.

GDC-DCB-68-017. “Reusable Launch Vehicle/Spacecraft Concept.” November 1968.

GDC-DCB-69-046. “Space Shuttle Final Technical Report.” 31 October 1969.

Grumman

B35-43 RP-33. “Space Shuttle System Program Definition. Phase B Extension Final Report.” 15 

March 1972.

High Technology

T. A. Heppenheimer. “Making Planes from Powder.” September 1986, pp. 54-55.

ISABE (International Society for Air Breathing Engines)

97-7004. E. T. Curran. “Scramjet Engines: The First Forty Years.”

Journal of the British Interplanetary Society

Joel W. Powell. “Thor-Able and Atlas-Able.” (1984): 219-225.

Jet Propulsion

Lester Lees. “Laminar Heat Transfer Over Blunt-Nosed Bodies at Hypersonic Flight Speeds.” April 

1956, pp. 259-269.

Johns Hopkins APL Technical Digest

William Garten, and Frank A. Dean. “Evolution of the Talos Missile.” 3, No. 2 (1982): 117-122.

AD-461220. A. R. Reti and J. R. Turner. “Study of Oxides for Chemical Separation of Air Into 

O2

–N2 Components.” Dynatech Corporation, March 1965.

AD-500489. “Air Enrichment Study Program.” Union Carbide, February 1969.

AD-609169. William Cowie. “Utilization of Refractory Metals on the X-20A (Dyna-Soar).” Air 

Force Systems Command, June 1964.

AD-832686. “Flight Test Evaluation Report, Missile 7D.” Convair, 3 June 1959.

ADA-274530. “Report of the Defense Science Board Task Force on National Aero-Space Plane 

(NASP) Program.” November 1992.

ADA-303832. Roy Franklin Houchin. “The Rise and Fall of Dyna-Soar: A History of Air Force 

Hypersonic R&D, 1944-1963.” Air Force Institute of Technology, August 1995.

ADB-190937. “NASP Materials and Structures Augmentation Program Final Report. Refractory 

Composites.” General Dynamics, March 1993.

ADB-191402. “1993 National Aero-Space Plane Technology Review. Volume V—Materials.” NASP 

Joint Program Office, July 1993.

ADB-191464. G. James Van Fossen. “Performance Testing and Modelling of a Water Alleviation 

System.” NASA-Lewis, November 1991.

ADB-191627. John Bradley. “Test and Evaluation Report for Carbon/Carbon Wing Box 

Component.” General Dynamics, 21 February 1992.

ADB-191699. “Summary of Results of Carbon-Carbon Composites Technology Maturation 

Program for NASP.” NASA-Langley, May 1992.

ADB-191715. P. C. Davis. “Low-Speed Uncertainty Analysis Report.” General Dynamics, 30 

September 1992.

ADB-191898. “High Conductivity Composites. Volume 1—Executive Summary, Copper Materials, 

Beryllium Materials, Coatings, Ceramic Materials and Joining.” Rockwell International, March 

1993.

ADB-192311. “National Program Office Review.” Engine review at West Palm Beach, Florida—6 

August 1991. NASP National Program Office, August 1991.

ADB-192559. “NASP Materials and Structures Augmentation Program. Titanium Matrix 

Composites.” McDonnell Douglas, 31 December 1991.

ADB-196085. Lecia L. Mayhew and James L. McAfee. “Full Scale Assembly Summary Status 

Report.” McDonnell Douglas, 8 November 1994.

ADB-197041. R. L. Campbell, T. J. Treinen, and G. H. Ratekin. “Low Speed Propulsion Summary 

Report.” Rocketdyne, 19 May 1993.

ADB-197189. “Composite Configuration Steering Committee Report.” NASP Joint Program 

Office, 14 February 1990.
308

Facing the Heat Barrier: A History of Hypersonics

309

Bibliography

Memorial Lecture. (1964): 575-597.

Journal of Spacecraft and Rockets

B. J. Griffith and D. E. Boylan. “Postflight Apollo Command Module Aerodynamic Simulation 

Tests.” 5 (1968): 843-848.

F. S. Billig, R. C. Orth, and M. Lasky. “Effects of Thermal Compression on the Performance 

Estimates of Hypersonic Ramjets.” 5 (1968): 1076-1081.

George W. Sutton. “The Initial Development of Ablation Heat Protection, An Historical 

Perspective.” 19 (1982): 3-11.

Steven P. Schneider. “Flight Data for Boundary-Layer Transition at Hypersonic and Supersonic 

Speeds.” 36 (1999): 8-20.

Frank S. Milos and Thomas H. Squire. “Thermostructural Analysis of X-34 Wing Leading-Edge Tile 

Thermal Protection System.” 36 (1999): 189-198.

Kathryn E. Wurster, Christopher J. Riley, and E. Vincent Zoby. “Engineering Aerothermal Analysis 

for X-34 Thermal Protection System Design.” 36 (1999): 216-227.

Walter C. Engelund, Scott D. Holland, Charles E. Cockrell and Robert D. Bittner. “Aerodynamic 

Database Development for the Hyper-X Airframe-Integrated Scramjet Propulsion Experiments.” 38 

(2001): 803-810.

Scott D. Holland, William C. Woods, and Walter C. Engelund. “Hyper-X Research Vehicle 

Experimental Aerodynamics Test Program Overview.” 38 (2001): 828-835.

Charles E Cockrell, Walter C. Engelund, Robert D. Bittner, Tom N. Jentink, Arthur D. Dilley and 

Abdelkader Frendi. “Integrated Aeropropulsive Computational Fluid Dynamics Methodology for the 

Hyper-X Flight Experiment.” 38 (2001): 836-843.

Lawrence D. Huebner, Kenneth E. Rock, Edward G. Ruf, David W. Witte and Earl H. Andrews. 

“Hyper-X Flight Engine Ground Testing for Flight Risk Reduction.” 38 (2001): 844-852.

Scott A. Berry, Aaron H. Auslender, Arthur D. Dilley and John F. Calleja. “Hypersonic BoundaryLayer Trip Development for Hyper-X.” 38 (2001): 853-864.

Langley Working Papers

LWP-54. James L. Raper. “Preliminary Results of a Flight Test of the Apollo Heat Shield Material at 

28,000 Feet per Second.” NASA-Langley, 27 October 1964.

Lockheed

LAC 571375. “Aerospace Plane.” 20 April 1962.

LMSC-A959837. “Final Report: Integrated Launch and Re-Entry Vehicle.” 22 December 1969.

Lockheed Horizons

Harold E. Gilreath. “The Beginning of Hypersonic Ramjet Research at APL.” 11, Nos. 3 and 4 

(1990): 319-335.

James L Keirsey. “Airbreathing Propulsion for Defense of the Surface Fleet.” Vol. 13, No. 1 (1992): 

57-67.

Johns Hopkins Magazine

Alan Newman. “Speed Ahead of Its Time.” December 1988, pp. 26-31.

Journal of the Aeronautical Sciences.

Theodore von Karman. “Supersonic Aerodynamics—Principles and Applications.” Tenth Wright 

Brothers Lecture. (1947): 373-409.

Charles H. McLellan. “Exploratory Wind-Tunnel Investigation of Wings and Bodies at M = 6.9.” 

(1951): 641-648.

J. A. Fay and F. R. Riddell. “Theory of Stagnation Point Heat Transfer in Dissociated Air.” (1958): 

73-85.

F. H. Rose and W. I. Stark. “Stagnation Point Heat-Transfer Measurements in Dissociated Air.” 

(1958): 86-97.

Journal of the Aero/Space Sciences

George W. Sutton. “Ablation of Reinforced Plastics in Supersonic Flow.” (1960): 377-385.

Mac C. Adams, William E. Powers, and Steven Georgiev. “An Experimental and Theoretical Study of 

Quartz Ablation at the Stagnation Point.” (1960): 535-543.

Journal of Aircraft

A. Ferri, “Review of SCRAMJET Propulsion Technology.” 5 (1968): 3-10.

Journal of Applied Physics

John V. Becker. “Results of Recent Hypersonic and Unsteady Flow Research at the Langley 

Aeronautical Laboratory.” 21 (1950): 619-628.

E. L. Resler, Shao-Chi Lin, and Arthur Kantrowitz. “The Production of High Temperature Gases in 

Shock Tubes.” 23 (1952): 1390-1399.

Journal of Guidance and Control

Charles Stark Draper. “Origins of Inertial Navigation.” 4 (1981), pp. 449-63.

Journal of the Royal Aeronautical Society

R. L. Schleicher. “Structural Design of the X-15.” (1963): 618-636.

Antonio Ferri. “Review of Problems in Application of Supersonic Combustion.” Seventh Lanchester 
310

Facing the Heat Barrier: A History of Hypersonics

311

Bibliography

RM L58E07a. Maxime A. Faget, Benjamine J. Garland, and James J. Buglia. “Preliminary Studies of 

Manned Satellites. Wingless Configuration: Nonlifting.” 11 August 1958.

TN 2171. Charles H. McLellan, Thomas W. Williams, and Mitchel H. Bertram. “Investigation of a 

Two-Step Nozzle in the Langley 11-inch Hypersonic Tunnel.” September 1950.

TN 2206. Irving Pinkel and John S. Serafini. “Graphical Method for Obtaining Flow Field in TwoDimensional Supersonic Stream to Which Heat Is Added.” 1 November 1950.

TN 2223. Charles H. McLellan, Thomas W. Williams, and Ivan E. Beckwith,. “Investigation of the 

Flow Through a Single-Stage Two-Dimensional Nozzle in the Langley 11-inch Hypersonic Tunnel.” 

December 1950.

TN 2773. Mitchel H. Bertram. “An Approximate Method for Determining the Displacement Effects 

and Viscous Drag of Laminar Boundary Layers in Two-Dimensional Hypersonic Flow.” September 

1952.

TN 3302. Charles H. McLellan and Thomas W. Williams. “Liquefaction of Air in the Langley 11-

inch Hypersonic Tunnel.” October 1954.

TN 4046. Alfred J. Eggers, H. Julian Allen, and Stanford E. Neice. “A Comparative Analysis of the 

Performance of Long-Range Hypervelocity Vehicles.” October 1957.

TN 4386. Richard J. Weber and John S. MacKay. “An Analysis of Ramjet Engines Using Supersonic 

Combustion.” September 1958.

NASA (National Aeronautics and Space Administration)

CP-3105. “Proceedings of the X-15 First Flight 30th Anniversary Celebration.” 1991.

RP-1028. Joseph Adams Shortal. “A New Dimension. Wallops Island Flight Test Range: The First 

Fifteen Years.” December 1978.

RP-1132. “Aeronautical Facilities Catalogue Volume 1. Wind Tunnels.” January 1985.

SP-60. Wendell H. Stillwell. “X-15 Research Results.” 1965.

SP-292. “Vehicle Technology for Civil Aviation. The Seventies and Beyond.” 1971.

SP-440. Donald D. Baals and William R. Corliss. “Wind Tunnels of NASA.” 1981.

SP-2000-4029. Richard W. Orloff. “Apollo By the Numbers: A Statistical Reference.” 2000.

SP-4201. Loyd S. Swenson, James M. Grimwood, and Charles C. Alexander. This New Ocean: A 

History of Project Mercury. 1966.

SP-4206. Roger E. Bilstein. Stages to Saturn: A Technological History of the Apollo/Saturn Launch 

Vehicles. 1980.

SP-4220. R. Dale Reed and Darlene Lister. Wingless Flight: The Lifting Body Story. 1997.

Clarence “Kelly” Johnson. “Development of the Lockheed SR-71 Blackbird.” Issue 9, Winter 

1981/82, pp. 2-7, 13-18.

Wilson B. Schramm, Ronald P. Banas, and Y. Douglas Izu. “Space Shuttle Tile—The Early Lockheed 

Years.” Issue 13, 1983, pp. 2-15.

Martin Marietta

M-63-1. C. W. Spieth and W. T. Teegarden, “Astrorocket Progress Report.” December 1962.

ER 14465. “SV-5D PRIME Final Flight Test Summary.” September 1967.

Marquardt Corporation

“Scramjet Flight Test Program.” Brochure, September 1965.

McDonnell Douglas

MDC E0056. “A Two-Stage Fixed Wing Space Transportation System.” 15 December 1969.

MDC E0308. “Space Shuttle System Phase B Study Final Report.” 30 June 1971.

MDC E076-1. “External LH2

 Tank Study Final Report.” 30 June 1971.

“Interim Report to OMSF: Phase B System Study Extension.” 1 September 1971.

NACA (National Advisory Committee for Aeronautics)

1381. H. Julian Allen and A. J. Eggers. “A Study of the Motion and Aerodynamic Heating of 

Ballistic Missiles Entering the Earth’s Atmosphere at High Supersonic Speeds.” 28 April 1953.

RM L51D17. Charles H. McLellan, Mitchel H. Bertram, and John A. Moore. “An Investigation 

of Four Wings of Square Plan Form at a Mach Number of 6.86 in the Langley 11-inch Hypersonic 

Tunnel.” 29 June 1951.

RM L51J09. Ralph D. Cooper and Raymond A. Robinson. “An Investigation of the Aerodynamic 

Characteristics of a Series of Cone-Cylinder Configurations at a Mach Number of 6.86.” 17 

December 1951.

RM E51K26. Irving Pinkel, John S. Serafini, and John L. Gregg, “Pressure Distribution and 

Aerodynamic Coefficients Associated with Heat Addition to Supersonic Air Stream Adjacent to TwoDimensional Supersonic Wing.” February 14, 1952.

RM L54F21. Charles H. McLellan. “A Method for Increasing the Effectiveness of Stabilizing 

Surfaces at High Supersonic Mach Numbers.” 3 August 1954.

RM A55E26. Alvin Seiff and H. Julian Allen, “Some Aspects of the Design of Hypersonic BoostGlide Aircraft.” August 15, 1955.

RM A55L05. A. J. Eggers and Clarence Syvertson, “Aircraft Configurations Developing High LiftDrag Ratios at High Supersonic Speeds.” March 5, 1956.
312

Facing the Heat Barrier: A History of Hypersonics

313

Bibliography

TM X-2273. “NASA Space Shuttle Technology Conference.” Volume 2—Structures and Materials. 

1 April 1971.

TM X-2572. “Hypersonic Research Engine Project Technological Status 1971.” September 1972.

TM X-2895. John R. Henry and Griffin Y. Anderson. “Design Considerations for the AirframeIntegrated Scramjet.” December 1973.

TM X-56008. Lawrence W. Taylor and Elmor J. Adkins. “Adaptive Flight Control Systems—Pro and 

Con.” 15 May 1964.

TN D-1157. Lawrence W. Taylor and George B. Merrick. “X-15 Airplane Stability Augmentation 

System.” March 1962.

TN D-1159. Robert G. Hoey and Richard E. Day. “Mission Planning and Operational Procedures 

for the X-15 Airplane.” March 1962.

TN D-1278. Joseph Weil. “Review of the X-15 Program.” 6 April 1962.

TN D-1402. Robert A. Tremant. “Operational Experiences and Characteristics of X-15 Flight 

Control System.” December 1962.

TN D-2996. William I. Scallion and John L. Lewis. “Flight Parameters and Vehicle Performance for 

Project Fire Flight 1, Launched April 14, 1964.” 3 June, 1965.

TN D-4713. Randolph A. Graves and William G. Witte. “Flight-Test Analysis of Apollo HeatShield Material Using the Pacemaker Vehicle System.” 21 May 1968.

TN D-6208. “Experience with the X-15 Adaptive Flight Control System.” March 1971.

TN D-7564. James E. Pavlosky and Leslie G. St. Leger. “Apollo Experience Report—Thermal 

Protection Subsystem.” January 1974.

TP-1998-206548. Alexander S. Roudakov, Vyacheslav L. Semenov, and John W. Hicks. “Recent 

Flight Test Results of the Joint CIAM-NASA Mach 6.5 Scramjet Flight Test Program.” April 1998.

Nature

R. N. Hollyer, A.C. Hunting, Otto Laporte and E.B. Turner. “Luminosity Generated by Shock 

Waves.” 171 (1953): 395-396.

North American Aviation

AL-1347. “Development of a Strategic Missile and Associated Projects.” October 1951.

North American Rockwell

SD 69-573-1. “Study of Integral Launch and Reentry Vehicle System.” December 1969.

SD 71-114-1. “Space Shuttle Phase B Final Report.” 25 June 1971.

SV 71-28. “Fully Reusable Shuttle.” 19 July 1971.

SP-4221. T. A. Heppenheimer. The Space Shuttle Decision. 1999.

SP-4222. J. D. Hunley, ed. Toward Mach 2: The Douglas D-558 Program. 1999.

SP-2004-4236. Dennis R. Jenkins, The X-15. 2004.

SP-4302. Edwin P. Hartman. Adventures in Research: A History of Ames Research Center, 1940-1976. 

1970.

SP-4303. Richard P. Hallion. On the Frontier: Flight Research at Dryden, 1946-1981. 1981.

SP-4305. James R. Hansen. Engineer in Charge: A History of the Langley Aeronautical Laboratory, 

1917-1958. 1987.

SP-4308. James R. Hansen. Spaceflight Revolution: NASA Langley Research Center from Sputnik to 

Apollo. 1995.

SP-4404. John L. Sloop. Liquid Hydrogen as a Propulsion Fuel, 1945-1959. 1978.

SP-4407. John M. Logsdon, ed. Exploring the Unknown: Selected Documents in the History of the U.S. 

Civil Space Program. Volume 1, Organizing for Exploration, 1995; Volume IV, Accessing Space, 1999.

SP-2000-4408. Asif A. Siddiqi, Challenge to Apollo: The Soviet Union and the Space Race, 1945-

1974. 2000.

SP-2000-4518. Dennis R. Jenkins, Hypersonics Before the Shuttle: A Concise History of the X-15 

Research Airplane. June 2000.

TM X-490. Robert L. O’Neal and Leonard Rabb. “Heat-Shield Performance During Atmosphere 

Entry of Project Mercury Research and Development Vehicle.” 12 October 1960.

TM X-1053. Richard C. Dingeldein. “Flight Measurements of Re-entry Heating at Hyperbolic 

Velocity (Project Fire).” 14 October 1964.

TM X-1120. Elden S. Cornette. “Forebody Temperatures and Total Heating Rates Measured During 

Project Fire 1 Reentry at 38,000 Feet per Second.” 20 April 1965.

TM X-1130. William T. Schaefer. “Characteristics of Major Active Wind Tunnels at the Langley 

Research Center.” July 1965.

TM X-1182. James Raper. “Results of a Flight Test of the Apollo Heat-Shield Material at 28,000 

Feet per Second.” 2 September 1965.

TM X-1222. Dona L. Cauchon. “Project Fire 1 Radiative Heating Experiment.” 8 December 1965.

TM X-1305. Elden S. Cornette. “Forebody Temperatures and Calorimeter Heating Rates Measured 

During Project Fire II Reentry at 11.35 Kilometers per Second.” 7 July 1966.

TM X-1407. Richard M. Raper. “Heat-Transfer and Pressure Measurements Obtained During 

Launch and Reentry of the First Four Gemini-Titan Missions and Some Comparisons with WindTunnel Data.” 1 March 1967.
314

Facing the Heat Barrier: A History of Hypersonics

315

Bibliography

Antony Jameson. “Computational Aerodynamics for Aircraft Design.” 245 (1989): 361-371.

Richard E. Young, Martha A. Smith, and Charles K. Sobeck. “Galileo Probe: In Situ Observations of 

Jupiter’s Atmosphere.” 272 (1996): 837-838.

“World’s Fastest Supercomputers.” 298 (2002): 1713.

David Malakoff. “Mach 12 by 2012?” 300 (2003): 888-889.

Scientific American

Ellis Levin, Donald D. Viele, and Lowell B. Eldrenkamp. “The Lunar Orbiter Missions to the 

Moon.” May 1968, pp. 58-78.

Spaceflight

Irene Sanger-Bredt. “The Silver Bird Story.” (1973): 166-181.

Curtis Peebles. “The Origin of the U.S. Space Shuttle—1.” (1979): 435-442.

———. “Project Bomi.” (1980): 270-272.

Space World

“Sophisticated Insulation Barrier to Protect Crew of Space Shuttle.” June-July 1979, pp. 18, 23-24.

Time

“Journey Into Space.” 8 December 1952, pp. 62-73.

“The Rough & the Smooth.” 18 November 1957, pp. 19-20.

“‘Like a Bullet.’” 8 December 1958, p. 15.

“The Great Capsule Hunt.” 27 April 1959, pp. 16-17.

“Educated Satellites.” 27 April 1959, p. 65.

“Back from Space.” 13 June 1960, pp. 70-76.

“A Place in Space.” 27 October 1961, pp. 89-94.

“A Decade of Deadly Birds.” 22 May 1964, p. 25.

“Moonward Bound.” 17 November 1967, pp. 84-85.

“The Sky’s the Limit.” 29 November 2004, pp. 63-67.

U.S. Patents

2,735,263. J. O. Charshafian. “Combination Rocket and Ram-Jet Power Plant.” Filed 6 December 

1947.

SV 71-50. “Space Shuttle Phase B Extension, 4th Month Review.” 3 November 1971.

Pratt & Whitney

“Dependable Engines...Since 1925.” Product summary, July 1990.

Proceedings of the Royal Society

William Payman and Wilfred Charles Furness Shepherd, “Explosion Waves and Shock Waves VI. 

The Disturbances Produced by Bursting Diaphragms with Compressed Air.” Series A, 186 (1946): 

293-321.

Quarterly Review (Air Research and Development Command)

Lt. Col. D. D. Carlson. “Research at AEDC.” Summer 1959, pp. 80-88.

Leon S. Jablecki. “Development of High-Priority Aerial Weapons Advanced in AEDC’s New Mach 8 

Tunnel.” Fall 1959, pp. 56-61.

Rand Corporation

R-217. J. E. Lipp, R. M. Salter, and R. S. Wehner. “The Utility of a Satellite Vehicle for 

Reconnaissance.” April 1951.

SM-11827. “Preliminary Design of an Experimental World-Circling Spaceship.” 2 May 1946. 

Reprinted in part in NASA SP-4407, Vol. I, pp. 236-244.

Raumfahrtforschung

John V. Becker. “The X-15 Program in Retrospect.” (1969): 45-53.

Republic Aviation News

“Kartveli Envisions a Mach 25 Vehicle As ‘Ultimate Airplane.’” 9 September 1960, pp. 1, 5.

Review of Scientific Instruments

Walter Bleakney, D. K. Weimer, and C. H. Fletcher. “The Shock Tube: A Facility for Investigations 

in Fluid Dynamics.” 20 (1949): 807-815.

Rocketdyne

RI/RD87-142. “Space Shuttle Main Engine.” 1 May 1987.

“Thirty Years of Rocketdyne.” Company publication, 1985.

Science

Gina Bari Kolata. “Who Will Build the Next Supercomputer?” 211 (1981): 268-269.

R. Jeffrey Smith. “Estrangement on the Launch Pad.” 224 (1984): 1407-1409.
316

Facing the Heat Barrier: A History of Hypersonics

317

Bibliography

2,883,829. Alfred Africano. “Rocket Engine Convertible to a Ramjet Engine.” Filed 24 January 

1955.

2,922,286. Randolph Rae. “Production of Liquid Oxygen.” Filed 13 August 1954.

3,040,519. Randolph Rae. “Jet Propulsion Unit with Cooling Means for Incoming Air.” Filed 13 

August 1954.

3,040,520. Randolph Rae. “Jet Power Unit for an Aircraft.” Filed 22 November 1954.

3,172,253. Helmut Schelp and Frederick Hughes. “Combination Turbo and Ramjet Propulsion 

Apparatus.” Filed 2 January 1959.

3,525,474. Hans von Ohain, Roscoe M. Mills, and Charles A. Scolatti. “Jet Pump or Thrust 

Augmentor.” Filed 9 December 1968.

3,690,102. Anthony A. duPont. “Ejector Ram Jet Engine.” Filed 29 October 1970.

Vought Missiles and Space Company

“Technical Overview: Oxidation Resistant Carbon-Carbon for the Space Shuttle.” Undated; circa 

1971.

NASA History Series

Reference Works, NASA SP-4000:

Grimwood, James M. Project Mercury: A Chronology. NASA SP-4001, 1963.

Grimwood, James M., and C. Barton Hacker, with Peter J. Vorzimmer. Project 

Gemini Technology and Operations: A Chronology. NASA SP-4002, 1969.

Link, Mae Mills. Space Medicine in Project Mercury. NASA SP-4003, 1965.

Astronautics and Aeronautics, 1963: Chronology of Science, Technology, and Policy. 

NASA SP-4004, 1964.

Astronautics and Aeronautics, 1964: Chronology of Science, Technology, and Policy. 

NASA SP-4005, 1965.

Astronautics and Aeronautics, 1965: Chronology of Science, Technology, and Policy. 

NASA SP-4006, 1966.

Astronautics and Aeronautics, 1966: Chronology of Science, Technology, and Policy. 

NASA SP-4007, 1967.

Astronautics and Aeronautics, 1967: Chronology of Science, Technology, and Policy. 

NASA SP-4008, 1968.

Ertel, Ivan D., and Mary Louise Morse. The Apollo Spacecraft: A Chronology, Volume I, Through November 7, 1962. NASA SP-4009, 1969.

Morse, Mary Louise, and Jean Kernahan Bays. The Apollo Spacecraft: A Chronology, 

Volume II, November 8, 1962–September 30, 1964. NASA SP-4009, 1973.

Brooks, Courtney G., and Ivan D. Ertel. The Apollo Spacecraft: A Chronology,

Volume III, October 1, 1964–January 20, 1966. NASA SP-4009, 1973.

Ertel, Ivan D., and Roland W. Newkirk, with Courtney G. Brooks. The Apollo 

Spacecraft: A Chronology, Volume IV, January 21, 1966–July 13, 1974. 

NASA SP-4009, 1978.

Astronautics and Aeronautics, 1968: Chronology of Science, Technology, and Policy. 

NASA SP-4010, 1969.

Newkirk, Roland W., and Ivan D. Ertel, with Courtney G. Brooks. Skylab: A 

Chronology. NASA SP-4011, 1977.

319
Facing the Heat Barrier: A History of Hypersonics

320

The NASA History Series

321

NASA SP-4022, 1986.

Astronautics and Aeronautics, 1978: Chronology of Science, Technology, and Policy. 

NASA SP-4023, 1986.

Astronautics and Aeronautics, 1979–1984: Chronology of Science, Technology, and 

Policy. NASA SP-4024, 1988.

Astronautics and Aeronautics, 1985: Chronology of Science, Technology, and Policy. 

NASA SP-4025, 1990.

Noordung, Hermann. The Problem of Space Travel: The Rocket Motor. Edited by 

Ernst Stuhlinger and J. D. Hunley, with Jennifer Garland. NASA SP-4026, 1995.

Astronautics and Aeronautics, 1986–1990: A Chronology. NASA SP-4027, 1997.

Astronautics and Aeronautics, 1990–1995: A Chronology. NASA SP-2000-4028, 

2000.

Management Histories, NASA SP-4100:

Rosholt, Robert L. An Administrative History of NASA, 1958–1963. NASA SP4101, 1966.

Levine, Arnold S. Managing NASA in the Apollo Era. NASA SP-4102, 1982.

Roland, Alex. Model Research: The National Advisory Committee for Aeronautics, 

1915–1958. NASA SP-4103, 1985.

Fries, Sylvia D. NASA Engineers and the Age of Apollo. NASA SP-4104, 1992.

Glennan, T. Keith. The Birth of NASA: The Diary of T. Keith Glennan. J. D. Hunley, 

editor. NASA SP-4105, 1993.

Seamans, Robert C., Jr. Aiming at Targets: The Autobiography of Robert C. Seamans, 

Jr. NASA SP-4106, 1996.

Garber, Stephen J., editor. Looking Backward, Looking Forward: Forty Years of U.S. 

Human Spaceflight Symposium. NASA SP-2002-4107, 2002.

Mallick, Donald L. with Peter W. Merlin. The Smell of Kerosene: A Test Pilot’s Odyssey. 

NASA SP-4108, 2003. 

Iliff, Kenneth W. and Curtis L. Peebles. From Runway to Orbit: Reflections of a 

Van Nimmen, Jane, and Leonard C. Bruno, with Robert L. Rosholt. NASA Historical Data Book, Volume I: NASA Resources, 1958–1968. NASA SP-4012, 1976, 

rep. ed. 1988.

Ezell, Linda Neuman. NASA Historical Data Book, Volume II: Programs and Projects, 1958–1968. NASA SP-4012, 1988.

Ezell, Linda Neuman. NASA Historical Data Book, Volume III: Programs and Projects, 1969–1978. NASA SP-4012, 1988.

Gawdiak, Ihor Y., with Helen Fedor, compilers. NASA Historical Data Book, Volume IV: NASA Resources, 1969–1978. NASA SP-4012, 1994.

Rumerman, Judy A., compiler. NASA Historical Data Book, 1979–1988: Volume V, 

NASA Launch Systems, Space Transportation, Human Spaceflight, and Space Science. 

NASA SP-4012, 1999.

Rumerman, Judy A., compiler. NASA Historical Data Book, Volume VI: NASA 

Space Applications, Aeronautics and Space Research and Technology, Tracking and 

Data Acquisition/Space Operations, Commercial Programs, and Resources, 1979–

1988. NASA SP-2000-4012, 2000.

Astronautics and Aeronautics, 1969: Chronology of Science, Technology, and Policy. 

NASA SP-4014, 1970.

Astronautics and Aeronautics, 1970: Chronology of Science, Technology, and Policy. 

NASA SP-4015, 1972.

Astronautics and Aeronautics, 1971: Chronology of Science, Technology, and Policy. 

NASA SP-4016, 1972.

Astronautics and Aeronautics, 1972: Chronology of Science, Technology, and Policy. 

NASA SP-4017, 1974.

Astronautics and Aeronautics, 1973: Chronology of Science, Technology, and Policy. 

NASA SP-4018, 1975.

Astronautics and Aeronautics, 1974: Chronology of Science, Technology, and Policy. 

NASA SP-4019, 1977.

Astronautics and Aeronautics, 1975: Chronology of Science, Technology, and Policy. 

NASA SP-4020, 1979.

Astronautics and Aeronautics, 1976: Chronology of Science, Technology, and Policy. 

NASA SP-4021, 1984.

Astronautics and Aeronautics, 1977: Chronology of Science, Technology, and Policy. 
Facing the Heat Barrier: A History of Hypersonics

322

The NASA History Series

323

NASA Engineer. NASA SP-2004-4109, 2004. 

Chertok, Boris. Rockets and People, Volume 1. NASA SP-2005-4110, 2005. 

Laufer, Alexander, Todd Post, and Edward Hoffman. Shared Voyage: Learning and 

Unlearning from Remarkable Projects. NASA SP-2005-4111, 2005. 

Dawson, Virginia P. and Mark D. Bowles. Realizing the Dream of Flight: Biographical Essays in Honor of the Centennial of Flight, 1903-2003. NASA SP-2005-4112, 

2005.

Project Histories, NASA SP-4200:

Swenson, Loyd S., Jr., James M. Grimwood, and Charles C. Alexander. This New 

Ocean: A History of Project Mercury. NASA SP-4201, 1966; rep. ed. 1998.

Green, Constance McLaughlin, and Milton Lomask. Vanguard: A History. NASA 

SP-4202, 1970; rep. ed. Smithsonian Institution Press, 1971.

Hacker, Barton C., and James M. Grimwood. On the Shoulders of Titans: A History 

of Project Gemini. NASA SP-4203, 1977.

Benson, Charles D., and William Barnaby Faherty. Moonport: A History of Apollo 

Launch Facilities and Operations. NASA SP-4204, 1978.

Brooks, Courtney G., James M. Grimwood, and Loyd S. Swenson, Jr. Chariots for 

Apollo: A History of Manned Lunar Spacecraft. NASA SP-4205, 1979. 

Bilstein, Roger E. Stages to Saturn: A Technological History of the Apollo/Saturn 

Launch Vehicles. NASA SP-4206, 1980, rep. ed. 1997.

SP-4207 not published.

Compton, W. David, and Charles D. Benson. Living and Working in Space: A History of Skylab. NASA SP-4208, 1983.

Ezell, Edward Clinton, and Linda Neuman Ezell. The Partnership: A History of the 

Apollo-Soyuz Test Project. NASA SP-4209, 1978.

Hall, R. Cargill. Lunar Impact: A History of Project Ranger. NASA SP-4210, 1977.

Newell, Homer E. Beyond the Atmosphere: Early Years of Space Science. NASA SP4211, 1980.

Ezell, Edward Clinton, and Linda Neuman Ezell. On Mars: Exploration of the Red 

Planet, 1958–1978. NASA SP-4212, 1984.

Pitts, John A. The Human Factor: Biomedicine in the Manned Space Program to 

1980. NASA SP-4213, 1985.

Compton, W. David. Where No Man Has Gone Before: A History of Apollo Lunar 

Exploration Missions. NASA SP-4214, 1989.

Naugle, John E. First Among Equals: The Selection of NASA Space Science Experiments. NASA SP-4215, 1991.

Wallace, Lane E. Airborne Trailblazer: Two Decades with NASA Langley’s Boeing 

737 Flying Laboratory. NASA SP-4216, 1994.

Butrica, Andrew J., editor. Beyond the Ionosphere: Fifty Years of Satellite Communication. NASA SP-4217, 1997.

Butrica, Andrew J. To See the Unseen: A History of Planetary Radar Astronomy. 

NASA SP-4218, 1996.

Mack, Pamela E., editor. From Engineering Science to Big Science: The NACA and 

NASA Collier Trophy Research Project Winners. NASA SP-4219, 1998.

Reed, R. Dale, with Darlene Lister. Wingless Flight: The Lifting Body Story. NASA 

SP-4220, 1997.

Heppenheimer, T. A. The Space Shuttle Decision: NASA’s Search for a Reusable Space 

Vehicle. NASA SP-4221, 1999.

Hunley, J. D., editor. Toward Mach 2: The Douglas D-558 Program. 

NASA SP-4222, 1999.

Swanson, Glen E., editor. “Before this Decade Is Out . . .”: Personal Reflections on the 

Apollo Program. NASA SP-4223, 1999.

Tomayko, James E. Computers Take Flight: A History of NASA’s Pioneering Digital 

Fly-by-Wire Project. NASA SP-2000-4224, 2000.

Morgan, Clay. Shuttle-Mir: The U.S. and Russia Share History’s Highest Stage. NASA 

SP-2001-4225, 2001.

Leary, William M. “We Freeze to Please”: A History of NASA’s Icing Research Tunnel 

and the Quest for Flight Safety. NASA SP-2002-4226, 2002.

Mudgway, Douglas J. Uplink-Downlink: A History of the Deep Space Network 

1957–1997. NASA SP-2001-4227, 2001. 
Facing the Heat Barrier: A History of Hypersonics

324

The NASA History Series

325

Dawson, Virginia P. and Mark D. Bowles. Taming Liquid Hydrogen: The Centaur 

Upper Stage Rocket, 1958-2002. NASA SP-2004-4230, 2004.

Meltzer, Michael. Mission to Jupiter: A History of the Galileo Project. NASA SP2007-4231.

Center Histories, NASA SP-4300:

Rosenthal, Alfred. Venture into Space: Early Years of Goddard Space Flight Center. 

NASA SP-4301, 1985.

Hartman, Edwin P. Adventures in Research: A History of Ames Research Center, 

1940–1965. NASA SP-4302, 1970.

Hallion, Richard P. On the Frontier: Flight Research at Dryden, 1946–1981. NASA 

SP-4303, 1984.

Muenger, Elizabeth A. Searching the Horizon: A History of Ames Research Center, 

1940–1976. NASA SP-4304, 1985.

Hansen, James R. Engineer in Charge: A History of the Langley Aeronautical Laboratory, 1917–1958. NASA SP-4305, 1987.

Dawson, Virginia P. Engines and Innovation: Lewis Laboratory and American Propulsion Technology. NASA SP-4306, 1991.

Dethloff, Henry C. “Suddenly Tomorrow Came . . .”: A History of the Johnson Space 

Center. NASA SP-4307, 1993.

Hansen, James R. Spaceflight Revolution: NASA Langley Research Center from Sputnik to Apollo. NASA SP-4308, 1995.

Wallace, Lane E. Flights of Discovery: 50 Years at the NASA Dryden Flight Research 

Center. NASA SP-4309, 1996.

Herring, Mack R. Way Station to Space: A History of the John C. Stennis Space Center. NASA SP-4310, 1997.

Wallace, Harold D., Jr. Wallops Station and the Creation of the American Space 

Program. NASA SP-4311, 1997.

Wallace, Lane E. Dreams, Hopes, Realities: NASA’s Goddard Space Flight Center, The 

First Forty Years. NASA SP-4312, 1999.

Dunar, Andrew J., and Stephen P. Waring. Power to Explore: A History of the Marshall Space Flight Center. NASA SP-4313, 1999.

Bugos, Glenn E. Atmosphere of Freedom: Sixty Years at the NASA Ames Research 

Center. NASA SP-2000-4314, 2000.

Schultz, James. Crafting Flight: Aircraft Pioneers and the Contributions of the Men 

and Women of NASA Langley Research Center. NASA SP-2003-4316, 2003.

General Histories, NASA SP-4400:

Corliss, William R. NASA Sounding Rockets, 1958–1968: A Historical Summary. 

NASA SP-4401, 1971.

Wells, Helen T., Susan H. Whiteley, and Carrie Karegeannes. Origins of NASA 

Names. NASA SP-4402, 1976.

Anderson, Frank W., Jr. Orders of Magnitude: A History of NACA and NASA, 

1915–1980. NASA SP-4403, 1981.

Sloop, John L. Liquid Hydrogen as a Propulsion Fuel, 1945–1959. NASA SP-4404, 

1978.

Roland, Alex. A Spacefaring People: Perspectives on Early Spaceflight. NASA SP4405, 1985.

Bilstein, Roger E. Orders of Magnitude: A History of the NACA and NASA, 1915–

1990. NASA SP-4406, 1989.

Logsdon, John M., editor, with Linda J. Lear, Jannelle Warren-Findley, Ray A. 

Williamson, and Dwayne A. Day. Exploring the Unknown: Selected Documents in 

the History of the U.S. Civil Space Program, Volume I, Organizing for Exploration. 

NASA SP-4407, 1995.

Logsdon, John M., editor, with Dwayne A. Day and Roger D. Launius. Exploring the Unknown: Selected Documents in the History of the U.S. Civil Space Program, 

Volume II, Relations with Other Organizations. NASA SP-4407, 1996.

Logsdon, John M., editor, with Roger D. Launius, David H. Onkst, and Stephen 

J. Garber. Exploring the Unknown: Selected Documents in the History of the U.S. 

Civil Space Program, Volume III, Using Space. NASA SP-4407, 1998.

Logsdon, John M., general editor, with Ray A. Williamson, Roger D. Launius, 

Russell J. Acker, Stephen J. Garber, and Jonathan L. Friedman. Exploring the Unknown: Selected Documents in the History of the U.S. Civil Space Program, Volume IV, 
Facing the Heat Barrier: A History of Hypersonics

326

The NASA History Series

327

Accessing Space. NASA SP-4407, 1999.

Logsdon, John M., general editor, with Amy Paige Snyder, Roger D. Launius, 

Stephen J. Garber, and Regan Anne Newport. Exploring the Unknown: Selected 

Documents in the History of the U.S. Civil Space Program, Volume V, Exploring the 

Cosmos. NASA SP-2001-4407, 2001.

Siddiqi, Asif A. Challenge to Apollo: The Soviet Union and the Space Race, 1945–

1974. NASA SP-2000-4408, 2000.

Hansen, James R., editor. The Wind and Beyond: Journey into the History of 

Aerodynamics in America, Volume 1, The Ascent of the Airplane. NASA SP-2003-

4409, 2003.

Monographs in Aerospace History, NASA SP-4500:

Launius, Roger D. and Aaron K. Gillette, compilers, Toward a History of the Space 

Shuttle: An Annotated Bibliography. Monograph in Aerospace History, No. 1, 1992.

Launius, Roger D., and J. D. Hunley, compilers, An Annotated Bibliography of the 

Apollo Program. Monograph in Aerospace History, No. 2, 1994.

Launius, Roger D. Apollo: A Retrospective Analysis. Monograph in Aerospace History, No. 3, 1994.

Hansen, James R. Enchanted Rendezvous: John C. Houbolt and the Genesis of the 

Lunar-Orbit Rendezvous Concept. Monograph in Aerospace History, No. 4, 1995.

Gorn, Michael H. Hugh L. Dryden’s Career in Aviation and Space. Monograph in 

Aerospace History, No. 5, 1996.

Powers, Sheryll Goecke. Women in Flight Research at NASA Dryden Flight Research 

Center, from 1946 to 1995. Monograph in Aerospace History, No. 6, 1997. 

Portree, David S. F. and Robert C. Trevino.Walking to Olympus: An EVA Chronology. Monograph in Aerospace History, No. 7, 1997.

Logsdon, John M., moderator. Legislative Origins of the National Aeronautics and 

Space Act of 1958: Proceedings of an Oral History Workshop. Monograph in Aerospace History, No. 8, 1998.

Rumerman, Judy A., compiler, U.S. Human Spaceflight, A Record of Achievement 

1961–1998. Monograph in Aerospace History, No. 9, 1998.

Portree, David S. F. NASA’s Origins and the Dawn of the Space Age. Monograph in 

Aerospace History, No. 10, 1998.

Logsdon, John M. Together in Orbit: The Origins of International Cooperation in the 

Space Station. Monograph in Aerospace History, No. 11, 1998. 

Phillips, W. Hewitt. Journey in Aeronautical Research: A Career at NASA Langley 

Research Center. Monograph in Aerospace History, No. 12, 1998. 

Braslow, Albert L. A History of Suction-Type Laminar-Flow Control with Emphasis 

on Flight Research. Monograph in Aerospace History, No. 13, 1999.

Logsdon, John M., moderator. Managing the Moon Program: Lessons Learned From 

Apollo. Monograph in Aerospace History, No. 14, 1999. 

Perminov, V. G. The Difficult Road to Mars: A Brief History of Mars Exploration in 

the Soviet Union. Monograph in Aerospace History, No. 15, 1999. 

Tucker, Tom. Touchdown: The Development of Propulsion Controlled Aircraft at 

NASA Dryden. Monograph in Aerospace History, No. 16, 1999.

Maisel, Martin D., Demo J. Giulianetti, and Daniel C. Dugan. The History of the XV15 Tilt Rotor Research Aircraft: From Concept to Flight. NASA SP-2000-4517, 2000.

Jenkins, Dennis R. Hypersonics Before the Shuttle: A Concise History of the X-15 

Research Airplane. NASA SP-2000-4518, 2000. 

Chambers, Joseph R. Partners in Freedom: Contributions of the Langley Research 

Center to U.S. Military Aircraft in the 1990s. NASA SP-2000-4519, 2000.

Waltman, Gene L. Black Magic and Gremlins: Analog Flight Simulations at NASA’s 

Flight Research Center. NASA SP-2000-4520, 2000.

Portree, David S. F. Humans to Mars: Fifty Years of Mission Planning, 1950–2000.

NASA SP-2001-4521, 2001.

Thompson, Milton O., with J. D. Hunley. Flight Research: Problems Encountered 

and What They Should Teach Us. NASA SP-2000-4522, 2000.

Tucker, Tom. The Eclipse Project. NASA SP-2000-4523, 2000.

Siddiqi, Asif A. Deep Space Chronicle: A Chronology of Deep Space and Planetary 

Probes, 1958–2000. NASA SP-2002-4524, 2002. 

Merlin, Peter W. Mach 3+: NASA/USAF YF-12 Flight Research, 1969–1979. 

NASA SP-2001-4525, 2001.
Facing the Heat Barrier: A History of Hypersonics

328

Anderson, Seth B. Memoirs of an Aeronautical Engineer—Flight Tests at Ames Research Center: 1940–1970. NASA SP-2002-4526, 2002. 

Renstrom, Arthur G. Wilbur and Orville Wright: A Bibliography Commemorating 

the One-Hundredth Anniversary of the First Powered Flight on December 17, 1903. 

NASA SP-2002-4527, 2002.

No monograph 28.

Chambers, Joseph R. Concept to Reality: Contributions of the NASA Langley Research Center to U.S. Civil Aircraft of the 1990s. SP-2003-4529, 2003.

Peebles, Curtis, editor. The Spoken Word: Recollections of Dryden History, The Early 

Years. SP-2003-4530, 2003.

Jenkins, Dennis R., Tony Landis, and Jay Miller. American X-Vehicles: An InventoryX-1 to X-50. SP-2003-4531, 2003.

Renstrom, Arthur G. Wilbur and Orville Wright: A Chronology Commemorating 

the One-Hundredth Anniversary of the First Powered Flight on December 17, 1903.

NASA SP-2003-4532, 2002.

Bowles, Mark D. and Robert S. Arrighi. NASA’s Nuclear Frontier: The Plum Brook 

Research Reactor. SP-2004-4533, 2003.

Matranga, Gene J. and C. Wayne Ottinger, Calvin R. Jarvis with D. Christian 

Gelzer. Unconventional, Contrary, and Ugly: The Lunar Landing Research Vehicle. 

NASA SP-2006-4535.

McCurdy, Howard E. Low Cost Innovation in Spaceflight: The History of the Near 

Earth Asteroid Rendezvous (NEAR) Mission. NASA SP-2005-4536, 2005.

Seamans, Robert C. Jr. Project Apollo: The Tough Decisions. NASA SP-2005-4537, 

2005.

Lambright, W. Henry. NASA and the Environment: The Case of Ozone Depletion.

NASA SP-2005-4538, 2005.

Chambers, Joseph R. Innovation in Flight: Research of the NASA Langley Research 

Center on Revolutionary Advanced Concepts for Aeronautics. NASA SP-2005-4539, 

2005.

Phillips, W. Hewitt. Journey Into Space Research: Continuation of a Career at NASA 

Langley Research Center. NASA SP-2005-4540, 2005.
Index

A

ablation, 39-40, 50, 152, 153, 165, 173, 174, 257. 

 See also carbon-carbon; hot structure.

ACES (Air Collection and Enrichment System), 111, 116, 122-25, 127-28

Adams, Major Mike, 85-86, 204.

Aerospaceplane, xi-xii, 91, 112-28, 178-79, 198, 200

 Aerospaceplane (Lockheed), 117-19

 Aerospaceplane (Republic), 119

 Astrorocket (Martin), 126-27, 179

 POBATO, 120

 Point Design, 1964 (Convair), 122

 press coverage, 115

 PROFAC (Northrop), 114

 ROLS (Holloman AFB), 122-23

 Space Plane (Convair), 112-14, 115, 117, 125

aircraft, speed of, ix, 65, 197

Allen, H. Julian, 29, 137-38

Allen-Eggers blunt-body principle, 29-30, 34, 55, 59-60, 6l, 133, 144,

151, 168, 173, 175

Apollo program, 133, 154-55, 156-57, 159, 173, 282

 Project Fire, 157-58

 Saturn V, 159, 284

Applied Physics Laboratory (Johns Hopkins), 95, 100-02, 112, 199, 202

arc tunnel, 40-42

Armstrong, Neil 76, 150

Arnold Engineering Development Center, 65-67, 70, 120,

 125, 156, 273, 277-78

ASSET, 165-68, 182, 191

Atlas ICBM, 25-29, 36, 44, 48-49, 72, 96, 135, 152, 153-54, 157, 261

Atwood, Lee, quoted, 96

Avery, William, 101, 102

A-4b, 5-6

A-9, 4-5, 24

B

Becker, John, xi, 2, 14-15, 20, 71, 82, 203, 206

 and boost-glide studies, 136, 138-40, 173

 and 11-inch wind tunnel, 2, 14-20, 30, 61, 70, 152, 155, 156

 and X-15 feasibility study, 58-61, 70, 72

329
Facing the Heat Barrier: A History of Hypersonics

330

Index

331

F

Faget, Maxime, 59, 155, 173

 Project Mercury, 151-52, 153 

 space shuttle, 179-85, 186, 190

Fay-Riddell theory of heating, 35-36, 151

Ferri, Antonio, xi, 91, 103-07, 116, 198, 199, 201-02, 207

Flax, Alexander, 116, 143

flow visualization, 17-18

flutter, aerodynamic, 145-46, 165, 168, 245

F-86 (also XP-86), 56, 65

F-100A, 65

F-104, ix, 64, 65, 197, 198

F-105 (also XF-105), 56, 65, 70, 114, 198

G

Galileo (mission to Jupiter), 257, 283

Gardner, Trevor, 27, 28, 69

Garrett AiResearch, 120, 203-04

Gemini program, 152, 155-56, 173

General Applied Science Laboratories (GASL), 103, 198, 201-02, 208, 210,

217, 242-43, 270

H

Hall, Colonel Edward, 115

Hallion, Richard, quoted, x-xi, 167-68, 204, 283, 284

Holloman AFB, 93, 122

hot structures, 142-48, 165-66, 174, 181, 185-87, 189, 222, 264

 columbium, 186

Huebner, Lawrence, 274, 275

Hunter, Maxwell, 57, 175, 187

hydrogen bomb, xii, 23, 27-29

hypersonics, xii, 29

 and Air Force, 31, 58, 66-67, 68-69, 200-01

 defined, ix

 difficulty of, 128, 237

 and Naxi Germany, 1-2, 3

 origin of term, x

 significance, 282-84

ICBM, 20, 23-24. See also Atlas; Minuteman.

Inconel X, 59, 71, 72, 74, 150, 266

Billig, Frederick, 102, 107, 199, 202, 216, 231

Bogdonoff, Seymour, 116, 220

Bomarc, 94-95, 197

Bomi, 57, 70, 133-35

boundary layers, 223, 224, 238. See also turbulence.

Branson, Richard, 283

Bush, Vannevar, quoted, 24

B-50 bomber, 92, 93, 200 

B-52 bomber, 59, 83, 197, 258, 272, 274, 275, 277

B-58 bomber, 116, 144, 183, 198

C

carbon-carbon, 165, 191-93, 222, 249-52

combined-cycle engines, 108, 215. See also LACE; ejectors.

computational fluid mechanics, 229, 230-32, 257, 273, 278-82.

See also boundary layers; Direct Numerical Simulation,

Large Eddy Simulation, Navier-Stokes equations; turbulence.

Concorde, ix, 284

Cooper, Robert, 214, 215, 216-17, 218, 222

Corona program, 152-53, 282

Czysz, Paul, quoted, 239, 244

D

DARPA, 214-16, 218

DC-X, 258-59, 263

Direct Numerical Simulation, 280, 281 

dirigibles, 284

Dornberger, Walter, 5, 57, 70, 133

Douglas Skyrocket, 56, 71, 79, 80

Draper, Alfred, 182-85

Dryden, Hugh, 68, 69, 203

Dugger, Gordon, 102, 199

DuPont, Anthony, 203-04, 215, 216, 218, 219, 222, 229

Dyna-Soar, xii, 12, 126, 127, 133, 137-43, 150, 159-60, 173, 174, 182

 early concepts, 133-37, 138-41, 148-49, 151 

 technology, 143-50, 191. See also Rene 41, Q-felt.

Dynatech, 124-25, 127-28

E

Eastham, James, quoted, 106

Eggers, Alfred, 29, 116, 137-38

Eisenhower, President Dwight, 46

ejectors, xii, 108-10, 120, 124, 215, 217, 229, 239, 242-44
Facing the Heat Barrier: A History of Hypersonics

332

Index

333

J

Jenkins, Dennis, quoted, 71, 73

Johnson, Clarence “Kelly”, 68-69, 197

Jupiter missile, 37, 46, 48

Jupiter-C missile, 44 K

Kantrowitz, Arthur, 23, 31-33, 40, 41, 50

Kartveli, Alexander, 56, 70, 91, 104, 114

Kennedy, President John F., quoted, 86

Keyworth, George, 216-17

Kincheloe, Iven, 67, 75

Knight, William “Pete”, 79, 81, 204-05 L

Large Eddy Simulation (LES), xii, 257, 280, 281-82

Lees theory of heating, 35, 36, 38, 151

lifting body, 165, 168-73, 179-82.

 See also PRIME.

Linde (division of Union carbide), 125, 127-28

Liquid Air Cycle Engine (LACE), 110-11, 114, 116, 119, 120-21, 123-24,

127-28, 240-41

 deicing, 239-40

 DuPont version, 215-16, 217, 269

 See also ACES, Dynatech, Linde.

Lombard, Albert, quoted, 57-58 M

Manned Orbiting Laboratory, 141, 143

Marquardt Aviation, 93, 94, 98, 101, 120-21, 123-24, 199, 201, 202,

203-04, 216, 242, 243

McLellan, Charles, 19-20, 55, 61, 173

McNamara, Robert, 70, 126, 133, 141-43

Mercury program, 80, 151-52, 153-54, 158, 166, 173

Messerschmitt Me-262, 12

Minuteman ICBM, 141, 200

Missiles, and aviation, 197

National Aerospace Plane (NASP), xii, 198, 218, 219

 Copper Canyon, 216-17, 219, 229-30

 engine design, 238

 funding, 218, 222, 224

 program reviews, 222-24, 252

 See also ejectors; LACE; X-30.

Navaho missile, 24-25, 56, 57, 65, 66, 72, 95-96, 101, 151, 197

Navier-Stokes equations, 230-31, 233, 234-35, 278, 282

Nicholls, James, 102, 104

nose cone, 29-30, 36-38, 41

 heat sink, 36, 37, 38-39, 44, 153, 189-90, 210

 Mark 3, 48-49

 RVX-1, 47, 49

 RVX-2, 48, 49

 See also ablation.

Nucci, Louis, 104, 201-02

nuclear freeze, 212-13 O

Orbital Sciences Corp., 258, 261-62, 270

Ordnance Aerophysics Laboratory, 13, 98, 201, 204 P

Paull, Allan, 274-75

Pegasus rocket, 258, 259, 270, 272, 273-74, 275, 277

Post-shuttle launch vehicles, 258-61. See also

 Transatmospheric vehicles; X-30; X-33; X-34.

Prandtl, Ludwig, 2

PRIME, 165, 170, 172-73, 174 Q

Q-felt , 146, 147, 177, 247 R

ramjets, 56, 66, 91-99, 114-15, 147, 197-98, 200, 202, 214-15

Rand Corp., 26, 27,57, 122, 214

Rausch, Vince, 214, 272

Reagan, President Ronald, xi, 198, 212-13, 218, 224

re-entry, x, 23, 30-31, 34-36. 

 of a satellite, 152

 See also hot structures; nose cones.

Rene 41, 143-44, 146, 150, 174

RL10 rocket engine, 123-24, 258

Roshko, Anatol, 236-37

Rutan, Burt, 283
Facing the Heat Barrier: A History of Hypersonics

334

Index

335

S

Sanger, Eugen, xi, 2, 8-12, 137, 284

scramjets, xi, xii, 98-99, 103-07, 111, 114, 117, 199-200, 215-16,

220, 221, 222, 231

 Australian, 274-75

 engine-airframe integration, 119, 206-11

 at General Electric, 199, 200, 203-04, 205

 Hypersonic Research Engine, 202-06, 210, 215

 inlets, 106-07, 200, 207, 208, 210, 231, 273

 origins, 100-02

 at Pratt & Whitney, 199, 201, 205, 243-44

 produces net thrust, 211

 and ramjets

 and rockets, 99, 253, 284

 Russian, 267-69

 significance, 284

 struts for fuel injection, 207-08, 210

 thermal compression, 105-06, 107, 207

 thrust and drag, 232, 272, 276, 277

 transition to supersonic combustion, 107-08, 268, 277, 286

 See also Ferri, Antonio; X-43A.

Shock tubes and tunnels, 23, 32-34, 231-32, 277

Single-stage-to-orbit, 121-22, 128, 259-61, 263

Skantze, General Lawrence, 198, 213-14, 217, 218

Soviet activities, 121, 151

SpaceShipOne, 283

Space Shuttle, xii, 121, 123, 127, 160, 165, 175-76, 193, 260, 284

 Columbia, 165, 191, 192-93

 configuration, 179-87, 190

 first stage, 187, 189-90

 early concepts, 175-76, 178-79

 Space Shuttle Main engine, 213-14, 219

 tests thermal protection

 thermal protection “tiles”, 165, 174-78, 181, 186-88, 191

space station, 193, 218, 284

specific impulse, 99, 123, 219, 220, 221

SR-71, ix, 43, 94, 106-07, 109, 110, 183, 186, 197, 198, 242

Stack, John, 13, 15

Storms, Harrison, 116, 151

Strategic Defense Initiative (SDI), xii, 198, 212-13, 216, 218, 224, 258, 266

 See also X-ray laser.

supercomputers, 230, 278-80

Sutton, George, 23, 36, 38-40, 47-48

T

Talos, 95, 101

Tether, Anthony, 215, 216

Thermal protection, 174

 See also ablation, hot structure nose cone re-entry space shuttle “tiles.”

Thomas, Arthur, 107, 112, 201, 202, 216, 218

Thompson, Floyd (NACA-Langley), 58

Thor missile, 38, 45, 166

Thor-Able missile, 47-48, 166-67

Thor-Delta launch vehicle, 166-67

Titan rockets, 44, 140, 147, 182, 200, 260

titanium, xii, 65, 72, 74, 134, 186, 187, 222

 Beta-21S, 246-47, 252

 powder metallurgy, 245-46

Transatmospheric Vehicles (TAVs), 213-14, 215, 217, 218

turbojets, 62-64, 66, 91-92, 116, 197-98, 199

turboramjet, 66, 108

turbulence, 223, 229, 232-37, 280-81

Twining, General Nathan, 28, 153

U

U. S. Air Force, 56, 61-62, 70, 112, 114-15

 Scientific Advisory Board, 31, 33, 58, 116-17, 121, 125-26, 145

 and space shuttle, 182-85, 198, 214

Von Braun, Wernher, 36-37

Von Karman, Theodore, 13, 27, 50

V-2, ix-x, 2, 3-6, 14, 24, 122, 133

W

Watkins, Admiral James, 212-13, 217

Weber, Richard and John MacKay, 101, 104

Weinberg, Steven, 50

Weinberger, Caspar, 217-18

Williams, Robert, 215-18, 229-30

wind tunnels, 2-3, 6-8, 13, 20, 66-67, 156, 201, 208, 231-32, 238,

271-73, 277-78

wings, 182-85, 190

Woods, Robert, 57, 70

Worth, Weldon, 114, 199

Wright Aeronautical Corp., 93, 95, 97, 98
Facing the Heat Barrier: A History of Hypersonics

336

X

X-ray laser, 212

X-l, 56, 57, 70, 80

X-lA, 57, 6l, 93, 173

X-lB, 75-76

X-2, 6, 56, 57, 58, 65, 67, 70, 75, 80-81, 93

X-3, 29, 55, 71

X-7, ix, 65, 92-94

X-15, x, xi, xii, 6, 20, 43, 55-56, 61, 65, 67-68, 116, 137, 138, 150, 151,

173, 189-90, 200-01, 266, 283

 aerodynamics and heating, 83-84

 auxiliary power units, 85

 engines, 80-8l

 flight controls, 67, 74-78

 flight simulators, 78-79

 and Hypersonic Research Engine, 202-03, 204-05

 Langley design concept, 58-60

 lubricanta, 74, 85

 piloted, 80, 85 

 pressure suits, 79

 and Q-ball, 67-68, 74

 selection of contractor, 70-73

 skin thickness, 56, 60

 stability of, 55, 61

 X-15A-2, 81-82

 X-15B, 151

 See also Becker, John; Inconel X.

X-17, 23, 42-44, 48

X-30, 198, 212, 218, 24l, 244, 284

 materials, 248-53

 NIFTA, 247-48, 266-67

 test program, 249

 weight and performance, 219-22, 264

 See also National Aerospace Plane.

X-33, xii, 257, 259, 261, 262-66

X-34, xii, 257, 259, 261-62, 263, 266

X-43A, x, xii, 269-74, 275-77, 278, 284

XB-70 bomber, 116, 122, 183, 197

XF-91 fighter, 70

XF-103 fighter, 56, 65, 66, 96-97, 197

Y

Yeager, Chuck, 61, 85, 93, 173 
The NASA History Series

National Aeronautics and Space Administration

NASA History Division

Office of External Relations

Washington, D.C.

2007

Thomas A. Heppenheimer has been a freelance 

writer since 1978. He has written extensively on 

aerospace, business and government, and the 

history of technology. He has been a frequent 

contributor to American Heritage and its affiliated 

publications, and to Air & Space Smithsonian. 

He has also written for the National Academy of 

Sciences, and contributed regularly to Mosaic of the 

National Science Foundation. He has written some 

300 published articles for more than two dozen 

publications.

He has also written twelve hardcover books. 

Three of them–Colonies in Space (1977), Toward 

Distant Suns (1979) and The Man-Made Sun

(1984)-have been alternate selections of the 

Book-of-the-Month Club. His Turbulent Skies

(1995), a history of commercial aviation, is part 

of the Technology Book Series of the Alfred P. 

Sloan Foundation. It also has been produced as 

a four-part, four-hour Public Broadcasting System 

television series Chasing the Sun.

Under contract to NASA, Heppenheimer has written 

that agency’s authorized history of the Space 

Shuttle, in two volumes. Volume 1, The Space 

Shuttle Decision (1999), has been reissued in 

paperback by the Smithsonian Institution Press and 

has been selected as an Outstanding Academic 

Title. The present book reflects his longstanding 

activity in hypersonics, for which he has written 

three technical reviews for Pasha Publications.

He holds a Ph.D. in aerospace engineering from 

the University of Michigan, and is an associate 

fellow of the American Institute of Aeronautics and 

Astronautics. He has also held research fellowships 

in planetary science at the California Institute of 

Technology and at the Max Planck Institute in 

Heidelberg, Germany.

About the Cover: Hypersonic Plane by Leslie 

Bossinas. Artist’s concept of an aerospace plane 

showing aero-thermal heating effects caused by 

friction as the vehicle flies hypersonically through 

the atmosphere. The National Aero-Space Plane 

program provided technology for space launch 

vehicles and hypersonic cruise vehicles. This 

vehicle with advanced airbreathing engines would 

have the capability to take off horizontally from 

and land on conventional runways, accelerate to 

orbit, and cruise hypersonically in the atmosphere 

between Earth destinations. (NASA Art Program, 

Image 86-HC-217).