NASA SP-2000-4522

Flight Research:

Problems Encountered and

What They Should Teach Us

by

Milton O. Thompson

with

A Background Section by J.D. Hunley

Monographs in Aerospace History # 22
NASA SP-2000-4522

Flight Research:

Problems Encountered and

What They Should Teach Us

by

Milton O. Thompson

with

A Background Section by J.D. Hunley

NASA History Division

Office of Policy and Plans

NASA Headquarters

Washington, DC 20546

Monographs in

Aerospace History

Number 22

2000
Library of Congress Cataloging-in-Publication Data

Thompson, Milton O.

 Flight Research : problems encountered and what they should teach us / Milton O.

Thompson ; with a background section by J.D. Hunley.

 p. cm. – (Monographs in aerospace history ;) (NASA history series) (NASA SP-2000

 ; 4522)

 Includes bibliographical references and index.

1. Aeronautics—Research—United States. 2. Airplanes—Flight testing. 3. High-speed

Aeronautics. I. Title. II. Series. III. Series: NASA history series IV. NASA SP ; 4522.

TL565.T46 2000

629.13’07’2073—dc2l 00-048072

For sale by the U.S. Government Printing Office

Superintendent of Documents, Mail Stop: SSOP, Washington, DC 20402-9328

ii
Foreword ......................................................................................................................................vi

Background: Flight Research at Dryden, 1946-1979 ....................................................................1

Problems Encountered in Flight Research ..................................................................................19

Aerodynamic Problems ........................................................................................................19

Environmental Problems ......................................................................................................24

Control-System Problems ....................................................................................................27

Structural Problems ..............................................................................................................28

Landing Gear Problems........................................................................................................28

Aerodynamic Heating Problems ..........................................................................................29

Auxiliary-Power-Unit Problems ..........................................................................................30

MH-96 Problems ..................................................................................................................30

Fatigue Problems..................................................................................................................32

X-15 Program Results ..........................................................................................................32

Control-System Problems in General...................................................................................32

First Flight Preparation ........................................................................................................35

First Flight and Envelope Expansion ...................................................................................37

Remotely Piloted Research Vehicles (RPRVs) ....................................................................38

Flight Test Errors..................................................................................................................40

Conclusions ..........................................................................................................................44

The Future ............................................................................................................................46

Index ............................................................................................................................................48

About the Author .........................................................................................................................50

Monographs in Aerospace History ..............................................................................................50

Table of Contents

iii
Foreword

The document by Milt Thompson that is reproduced here was an untitled rough draft found in Thompson’s

papers in the Dryden Historical Reference Collection. Internal evidence suggests that it was written about

1974. Readers need to keep this date in mind, since Milt writes in the present tense. Apparently, he never

edited the document. Had he prepared it for publication, he would have done lots of editing and refined much

of what he said.

I have not attempted to second guess what Milt might have done in revising the paper, but I have made some

minor stylistic changes to make it more readable without changing the sense of what Milt initially wrote.

Where I have qualified what Milt said or added information for the reader’s benefit, I have done so either in

footnotes or inside square brackets [like these]. The draft itself indicated that it should contain numerous

figures to illustrate what he wrote, but no such figures were associated with the manuscript. I have searched

out figures that appear to illustrate what Milt intended to show, but in some cases I have found none. When

that has been the case, I have deleted his references to figures and simply kept his text, which does stand on

its own.

For the most part, I have not attempted to bring his comments up to date, although in a few instances I have

inserted footnotes that indicate some obvious changes since he wrote the paper. Despite—or perhaps because

of—the paper’s age, it offers some perspectives on flight research that engineers and managers not familiar

with the examples Milt provides can still profit from in today’s flight-research environment. For that reason, I

have gone to the trouble to edit Milt’s remarks and make them available to those who would care to learn

from the past.

For readers who may not be familiar with the history of what is today the NASA Dryden Flight Research

Center and of its predecessor organizations, I have added a background section. Those who do know the

history of the Center may wish to skip reading it, but for others, it should provide context for the events Milt

describes. Milt’s biography appears at the end of the monograph for those who would like to know more

about the author of the document.

Many people have helped me in editing the original manuscript and in selecting the figures. The process has

gone on for so long that I am afraid to provide a list of their names for fear of leaving some important contributors out. A couple of them, in any event, requested anonymity. Let me just say a generic ‘thank you’ to

everyone who has assisted in putting this document into its present form, with a special thanks to the Dryden

Graphics staff members, especially Jim Seitz, for their work on the figures; to Jay Levine and Steve Lighthill

for laying the monograph out; to Darlene Lister for her assistance with copy editing; and to Camilla McArthur

for seeing the monograph through the printing process.

J. D. Hunley, Historian

iv
speed of sound (Mach 1) in level

flight.1

Even so, during the early 1940s,

airplanes like Lockheed’s P-38 Lightning began to face the problem of

compressibility in dives—characterized (among other things) by increased

density, a sharp rise in drag, and

disturbed airflow at speeds approaching Mach 1. The effects of compressibility included loss of elevator effectiveness and even the break-up of

structural members such as the tail,

killing pilots in the process. This

problem was compounded by the

absence of accurate wind-tunnel data

Milt Thompson’s account of lessons to be

learned assumes some familiarity with

the history of flight research at what was

then called the Flight Research Center

and its predecessor organizations—

redesignated in 1976 the Hugh L. Dryden

Flight Research Center. The following

account provides a brief history of the

subject that perhaps will provide a useful

backdrop to what Milt had to say.

From the time of the Wright brothers’

first flight in 1903 until the end of World

War II, airplane technology evolved

considerably. The early decades’ monoand biplanes of wooden framework,

typically braced with wire and covered

with cloth, gradually gave way to an allmetal construction and improved aerodynamic shapes, but most aircraft in World

War II still featured propellers and even

the fastest of them flew at maximum

speeds of about 450 miles per hour. For

example, the North American P-51

Mustang, one of the finest prop fighters

used in the war, had a top speed of 437

miles per hour when flying a level

course at low altitude. This compared

with low-level maximum speeds of 514

and 585 mph respectively for the

Messerschmitt Me 262A and Gloster

Meteor F.Mk. jet fighters, both of

which thus still flew well below the

1

 See, e.g., Laurence K. Loftin, Jr., Quest for Performance: The Evolution of Modern Aircraft (Washington, DC, NASA

SP-468, 1985), Chs. 1-5 and 9-10, esp. pp. ix, 7-45, 77-88, 128-136, 281-286, 484-490; Roger E. Bilstein, Flight in

America: From the Wrights to the Astronauts (rev. ed.; Baltimore, MD.: Johns Hopkins Univ. Press, 1994), pp. 3-40,

129-145.

Background:

Flight

Research at

Dryden,

1946-1979

J. D. Hunley

A P-51 Mustang

on the lakebed

next to the

NACA HighSpeed Flight

Station in 1955.

(NASA photo

E55-2078)

1
for portions of the transonic speed

range in a narrow band on either side of

the speed of sound.2

This situation led to the myth of a sound

barrier that some people believed could

not be breached. Since it appeared that jet

aircraft would soon have the capability of

flying in level flight into the transonic

region—where the dreaded compressibility effects abound—a solution was needed

for the lack of knowledge of transonic

aerodynamics. A number of people

(including Ezra Kotcher with the Army

Air Forces [AAF] at Wright Field in Ohio,

John Stack at the Langley Memorial

Aeronautical Laboratory of the National

Advisory Committee for Aeronautics

[NACA] in Virginia, Robert Woods with

Bell Aircraft, L. Eugene Root with

Douglas Aircraft, and Abraham Hyatt at

the Navy Bureau of Aeronautics) concluded that the solution could best result

from a research airplane capable of flying

at least transonically and even supersonically.

The emphases of these different organizations resulted in two initial aircraft—the

XS-1 (XS standing for eXperimental

Sonic, later shortened to X), for which

Bell did the detailed design and construction for the AAF, and the D-558-1

Skystreak, designed and constructed by

Douglas for the Navy. The XS-1 was the

faster of the two, powered by an XLR-11

rocket engine built by Reaction Motors

and launched from a B-29 or later a B-50

“mothership” to take full advantage of the

limited duration provided by its rocket

2

 James O. Young, Meeting the Challenge of Supersonic Flight (Edwards AFB, CA: Air Force Flight Test Center

History Office, 1997), pp. 1-2; John V. Becker, The High-Speed Frontier: Case Histories of Four NACA Programs

(Washington, DC: NASA SP-445, 1980), esp. p. 95. It should be noted here that the first studies of compressibility

involved tip speeds of propellers and date from 1918 to 1923. On these, see especially John D. Anderson, Jr., “Research

in Supersonic Flight and the Breaking of the Sound Barrier” in From Engineering Science to Big Science: The NACA

and NASA Collier Trophy Research Project Winners, ed. Pamela Mack (Washington, DC: NASA SP-4219, 1998), pp.

66-68. This article also provides excellent coverage of the early research of John Stack and his associates at the NACA’s

Langley Memorial Aeronautical Laboratory on the compressibility issue for aircraft (as opposed to propellers).

A P-38 Lightning in flight in

1943. (NASA

photo E95-

43116-2)

2
propulsion. Stack and the other NACA

engineers were skeptical about the rocket

engine and less concerned about breaking

the sound barrier than gathering flight

data at transonic speeds, so they supported the AAF-Bell project with critical

design data and recommendations but

were more enthusiastic about the NavyDouglas Skystreak. This was designed

with an early axial-flow turbojet

powerplant and was capable of flying

only up to Mach 1. However, with

comparably designed wings and a

movable horizontal stabilizer recommended by the NACA, plus the ability

to fly in the transonic region for a

XS-1 Number 2

on the ramp at

Edwards Air

Force Base

with its B-29

mothership.

(NASA photo

E-9)

A D-558-1 in

flight. (NASA

photo E-713)

3
3

 Young, Supersonic Flight, pp. 2-18; Becker, High-Speed Frontier, pp. 90-93; Richard Hallion, Supersonic Flight:

Breaking the Sound Barrier and Beyond, The Story of the Bell X-1 and the Douglas D-558 (rev. ed.; London and

Washington, DC: Brassey’s, 1997), esp. pp. 35-82; Louis Rotundo, Into the Unknown: The X-1 Story (Washington, DC

and London: The Smithsonian Institution Press, 1994), esp. pp. 11-33; Toward Mach 2: The Douglas D-558 Program, ed.

J. D. Hunley (Washington, DC: NASA SP-4222, 1999), esp. pp. 3-7.

longer period of time, the D-558-1

complemented the XS-1 nicely and

provided comparable data.3

To support the research flights, the

contractors, the AAF (after September

1947, the Air Force), and the NACA sent

teams of pilots and support personnel to

the Muroc Army Air Field starting in

September 1946 to support the XS-1,

and then the Navy joined in to help fly

the D-558. At Muroc, the 44-square-mile

XS-1 Number 2

on Rogers Dry

Lakebed in a

photo that gives

some sense of the

expanse of

natural runway

provided by the

dry lake. (NASA

photo E49-001)

XS-1 Number 1

in flight with

copy of “Mach

jump” paper

tape data record

of the first

supersonic flight

by Air Force

Capt. Chuck

Yeager. (NASA

photo E-38438

4
Rogers Dry Lakebed provided an enormous

natural landing field, and the clear skies and

sparse population provided an ideal environment for conducting classified flight

research and tracking the aircraft.4

The most immediate and dramatic result

of these twin flight research efforts that

proceeded simultaneously at Muroc was

Air Force pilot Chuck Yeager’s breaking

the sound barrier on 14 October 1947 in

the XS-1, for which feat he garnered the

Collier Trophy the next year in conjunction with John Stack for the NACA and

Larry Bell for his company.5 The flight

dispelled the myth about a sound barrier

4

 Rotundo, Into the Unknown, pp. 96, 123-132; James R. Hansen, Engineer in Charge: A History of the Langley Aeronautical Laboratory, 1917-1958 (Washington, DC: NASA SP-4305, 1987), p. 297. Contrary to what is reported in a

number of sources, the initial NACA contingent did not arrive on 30 Sept. 1946, with Walter C. Williams. Harold H.

Youngblood and George P. Minalga arrived Sunday, 15 Sept., William P. Aiken, sometime in October after Williams’

arrival. Telephonic intvws., Hunley with Youngblood and Aiken, 3 and 4 Feb. 1997.

5

 See especially Rotundo, Into the Unknown, pp. 279, 285.

NACA research

aircraft on the

ramp at the South

Base area of

Edwards Air

Force Base, (left

to right) D-558-2,

D-558-1, X-5, X1, XF-92A, X-4.

(NASA photo

EC-145)

In-flight photo of

the X-3. (NASA

photo E-17348)

5
and undoubtedly did much to gain credit

for flight research, resulting in the small

contingent of NACA engineers, pilots,

and support people at Muroc becoming a

permanent facility of the NACA and later

NASA.

All of this was extremely important,

but even more important than the

record and the glory that went with it

were the data that the NACA garnered

from the flight research not only with

the several X-1 and D-558-1 aircraft,

but also with the Douglas D-558-2,

the Bell X-2, the Douglas X-3 “flying

stiletto,” the Northrop X-4

semitailless, the Bell X-5 variablesweep, and the Convair XF-92A deltawinged aircraft. Not all of these airplanes were successful in a conventional

sense, even as research airplanes. But all

of them provided important data for

either validating or correcting information from wind tunnels and designing

future airplanes ranging from the Century series of fighter aircraft to today’s

commercial transports, which still fly in

the transonic speed range and feature the

movable horizontal stabilizer demonstrated on the X-1 and D-558s. Even the

ill-fated X-2, of which Dick Hallion has

written, “its research was nil,” and the

X-3, which he has dubbed “NACA’s

glamorous hangar queen,” 6 nevertheless

contributed to our understanding of the

insidious problem of coupling dynamics.

Furthermore, the Air Force-NACA X-2

program featured the first simulator used

for the various functions of flight-test

planning, pilot training, extraction of

aerodynamic derivatives, and analysis

of flight data.7

6

 Hallion, On the Frontier: Flight Research at Dryden, 1946-1981 (Washington, DC: NASA SP-4303, 1984), pp. 78,

59 respectively. In both cases, Hallion’s characterizations are justifiable in some degree.

7

 On the coupling and the computer simulation, Richard E. Day, Coupling Dynamics in Aircraft: A Historical Perspective (Dryden Flight Research Center, CA: NASA SP-532, 1997), esp. pp. 8-15, 34-36. On the value of the X-2 and X-3,

see also Ad Inexplorata: The Evolution of Flight Testing at Edwards Air Force Base (Edwards AFB, CA: Air Force

Flight Test Center History Office, 1996), pp. 14, 16. For the other research results, see especially Walter C. Williams

and Hubert M. Drake, “The Research Airplane: Past, Present, and Future,” Aeronautical Engineering Review (Jan.

1958): 36-41; Becker, High-Speed Frontier, pp. 42, 95-97; Hallion, On the Frontier, 59-62. In writing this account, I

have benefited greatly from comments made to me over the years by long-time Dryden research engineer Ed Saltzman.

X-2 in flight.

(NASA photo

E-2822)

6
Before this account discusses some of the

other highlights of flight research at what

became NASA’s Dryden Flight Research

Center, perhaps it should explain the

differences and similarities between

flight research and flight test. Both

involve highly trained, highly skilled

pilots and sometimes exotic or cuttingedge aircraft, although flight research can

use quite old aircraft modified for

particular kinds of research. There is no

hard and fast dividing line separating

flight research from flight test in practice, but flight research, unlike flight test

in most applications, is oblivious to the

particular aircraft employed so long as

that airplane can provide the required

flight conditions. On the other hand,

flight test, as the name implies, often

involves testing specific prototype or

early production aircraft (somewhat later

production aircraft in the case of operational flight-testing) to see if they fulfill

the requirements of a particular contract

and/or the needs of the user. In addition,

however, flight testing—at least in the

Air Force—involves flying aircraft that

may be quite old to try to improve them

and to develop their systems. For example, the Air Force Flight Test Center

recently began testing the F-22, a brand

new airplane, while at the same time it

continued to test the F-15 and its systems

even though various models of F-15s had

been in the inventory for more than two

decades.

In partial contrast to flight test, flight

research sought and seeks fundamental

understanding of all aspects of aeronautics, and in achieving that understanding,

its practitioners may fly experimental

aircraft like the early X-planes and the

D-558s or armed service discards like

early production models of the F-15s,

F-16s, and F-18s researchers at Dryden

are modifying and flying today. They

may even fly comparatively new aircraft

like the F-100 in its early days; here,

however, the purpose is not to test them

against contract standards but to understand problems they may be exhibiting in

operational flight and learn of ways to

correct them—a goal very similar to that

This NACA

High-Speed

Flight Station

photograph of

the Century

Series fighters

in formation

flight was

taken in 1957

(clockwise

from left —

F-104, F-101,

F-102, F-100).

(NASA photo

E-2952)

7
of flight testing in its effort to improve existing

aircraft. I should note in this connection that

the Air Force Flight Test Center (as it is called

today) and Dryden (under a variety of

previous names) have often cooperated in

flight research missions, with Air Force and

NACA/NASA pilots flying together. So

clearly flight test organizations in these cases

participate in flight research just as research

pilots sometimes engage in flight tests.8 It

should also be added that although many

researchers at what is today Dryden might be

quick, if asked, to point out the differences

between flight test and flight research, many

of them, including Milt in the account below,

often used the two terms as if they were

interchangeable.

To return to specific flight research projects

at Dryden, on 20 November 1953, with

NACA pilot Scott Crossfield in the pilot’s

seat, the D-558-2 exceeded Mach 2 in a

slight dive, and on 27 September 1956,

Air Force Capt. Mel Apt exceeded

Mach 3 in the X-2 before losing control

of the aircraft due to inertial coupling

and plunging to his death.9 With the

then-contemporary interest in space

flight, clearly there was a need at this

point for research into hypersonic

speeds (above Mach 5) and attendant

problems of aerodynamic heating, flight

above the atmosphere, and techniques

for reentry. In early 1954, therefore, the

NACA’s Research Airplane Projects

8

 The account that comes closest to what I have said above is Lane Wallace’s Flights of Discovery: 50 Years at the NASA

Dryden Flight Research Center (Washington, DC: NASA SP-4309, 1996), pp. 4-8. On flight test per se, see Ad

Inexplorata, esp. pp. 12-13. For a useful history of both flight testing and flight research, see Richard P. Hallion, “Flight

Testing and Flight Research: From the Age of the Tower Jumper to the Age of the Astronaut,” in Flight Test Techniques,

AGARD Conference Proceedings No. 452 (copies of papers presented at the Flight Mechanics Panel Symposium,

Edwards AFB, CA, 17-20 Oct. 1988), pp. 24-1 to 24-13. Finally, for an early discussion of flight research (despite its

title) see Hubert M. Drake, “Aerodynamic Testing Using Special Aircraft,” AIAA Aerodynamic Testing Conference,

Washington, DC, Mar. 9-10, 1964, pp. 178-188. In writing and refining the above two paragraphs, I have greatly

benefited from AFFTC Historian Jim Young’s insightful comments about flight test, especially as it is practiced today at

Edwards AFB, as well as from comments by Ed Saltzman. A point Ed offered that I did not incorporate in the narrative is

that in the obliviousness of flight research to the specific aircraft used, it has more in common with wind-tunnel research

than with flight test.

9

 Hallion, On the Frontier, pp. 308, 316.

The D-558-2

Number 2 is

launched from

the P2B-1 in this

1956 NACA

High-Speed

Flight Station

photograph.

This is the same

airplane that

Scott Crossfield

had flown to

Mach 2.005 in

1953. (NASA

photo E-2478)

8
Panel began discussion of a new

research airplane that became the X-15.

Developed under an Air Force contract with

North American Aviation, Inc., and flown

from 1959 to 1968, the X-15 set unofficial

world speed and altitude records of 4,520

miles per hour (Mach 6.7) and 354,200 feet

(67 miles).10

Much more importantly, however, the joint Air

Force-Navy-NASA-North American

program investigated all aspects of piloted

hypersonic flight. Yielding over 765

research reports, the 199-flight program

“returned benchmark hypersonic data for

aircraft performance, stability and control,

materials, shock interaction, hypersonic

turbulent boundary layer, skin friction,

reaction control jets, aerodynamic heating,

and heat transfer,”11 as well as energy

management. These data contributed to

the development of the Mercury,

11 Kenneth W. Iliff and Mary F. Shafer, Space Shuttle Hypersonic Aerodynamic and Aerothermodynamic Flight Research and the Comparison to Ground Test Results (Washington, DC: NASA Technical Memorandum 4499,1993), p. 2,

for quotation and see also their “A Comparison of Hypersonic Flight and Prediction Results,” AIAA-93-0311, paper

delivered at the 31st Aerospace Sciences Meeting & Exhibit, Jan. 11-14, 1993, in Reno, NV.

10 See esp. ibid., pp. 101-129, 333, 336, and Wendell H. Stillwell, X-15 Research Results (Washington, DC: NASA SP60, 1965), p. vi and passim.

The X-15 ship

Number 3 (56-6672)

is seen here on the

lakebed at the

Edwards Air Force

Base, California.

Ship Number 3 made

65 flights during the

program, attaining a

top speed of Mach

5.65 and a maximum

altitude of 354,200

feet. (NASA photo

E-7896)

In this photo the

Number 1 XB70A (62-0001) is

viewed from

above in cruise

configuration

with the wing

tips drooped for

improved controllability.

(NASA photo

EC68-2131)

9
Gemini, and Apollo piloted spaceflight

programs as well as the later Space

Shuttle program.12

Overlapping the X-15 program in time,

the XB-70 also performed significant

high-speed flight research. The XB-70

was the world’s largest experimental

aircraft. It was capable of flight at

speeds of three times the speed of

sound (roughly 2,000 miles per hour) at

altitudes of 70,000 feet. It was used to

collect in-flight information for use in

the design of future supersonic aircraft,

both military and civilian.

The more specific major objectives of

the XB-70 flight research program were

to study the airplane’s stability and

handling characteristics, to evaluate its

response to atmospheric turbulence, and

to determine the aerodynamic and

propulsion performance. In addition,

there were secondary objectives to

measure the noise and friction associated with airflow over the airplane and

to determine the levels and extent of the

engine noise during takeoff, landing,

and ground operations. The first flight

of the XB-70 was made on 21 September 1964. The Number two XB-70 was

destroyed in a mid-air collision on 8

June 1966. Program management of the

NASA-USAF research effort was

assigned to NASA in March 1967. The

final flight was flown on 4 February

1969. The program did provide a great

deal of data that could be applied to a

future supersonic transport or a large,

supersonic military aircraft. It also

yielded data on flight dynamics, sonic

booms, and handling qualities.13

Another important high-speed flight

research program involved the

Lockheed YF-12 “Blackbird,” precursor

12 See, e.g., John V. Becker, “The X-15 Program in Retrospect,” 3rd Eugen Sänger Memorial Lecture, Bonn, Germany,

Dec. 4-5, 1968, copy in the NASA Dryden Historical Reference Collection; Milton O. Thompson, At the Edge of Space:

The X-15 Flight Program (Washington, DC, and London: Smithsonian Institution Press, 1992); and the sources cited

above.

13 Hallion, On the Frontier, pp. 185-188; and see, e.g., P. L Lasagna and T. W. Putnam, “Engine Exhaust Noise during

Ground Operation of the XB-70 Airplane” (Washington, DC: NASA TN D-7043, 1971) and C. H. Wolowicz and R. B.

Yancey, Comparisons of Predictions of the XB-70-1 Longitudinal Stability and Control Derivatives with Flight Results

for Six Flight Conditions (Washington, DC: NASA TM X-2881, 1973).

A YF-12A in

flight. (NASA

photo EC72-

3150)

10
14 On the YF-12, see esp. Berwin M. Kock, “Overview of the NASA YF-12 Program,” YF-12 Experiments Symposium,

Vol. 1, (3 vols.; Washington, DC: NASA CP-2054, 1978) plus more specialized papers in the volume; Robert D. Quinn

and Frank V. Olinger, “Flight Temperatures and Thermal Simulation Requirements,” NASA YF-12 Flight Loads Program

(Washington, DC: NASA TM X-3061, 1974), pp. 145-183; and Hallion, On the Frontier, pp. 196-199, 349-356.

of the SR-71 reconnaissance airplane

that flew at Dryden during the 1990s.

Three YF-12s flew at Edwards in a joint

NASA-AF research program between

1969 and 1979. The aircraft studied the

thermal, structural, and aerodynamic

effects of sustained, high-altitude, Mach 3

flight. They also studied propulsion, air

flow and wind gusts, jet wake dispersion,

engine stalls, boundary-layer noise, and

much else. The 125 research reports the

program produced contained vast

amounts of information used in designing

or improving other supersonic aircraft,

including the SR-71. Among other things,

engineers at Dryden developed a central

airborne performance analyzer to monitor

YF-12 flight parameters. It became the

forerunner of the on-board diagnostic

system used on the Space Shuttle.14

Not all of Dryden’s flight research has

concerned high-speed flight. One crucial

flight research project that certainly had

implications for high-speed flight but was

not restricted to that regime was the F-8

Digital Fly-By-Wire project. Dryden

engineers replaced all purely mechanical

linkages to flight-control surfaces (rudders, ailerons, elevators, and flaps) in an

F-8C with electronic ones controlled by a

digital flight-control system. Although

there had been previous analog flight

control systems, this was not only the

first digital system but also the first

electronic system without a conventional mechanical backup, using an

analog backup instead. Flown in the

1970s and into the mid-1980s, the F-8

DFBW first used the Apollo computer

developed by Draper Lab and then the

IBM AP-101 later employed on the

Shuttle. Flying this system without a

mechanical backup was important in

giving industry the confidence to

develop its own digital systems since

flown on the F-18, F-16, F-117, B-2,

F-22, and commercial airliners like the

This photo shows the F-8

Digital-Fly-By Wire aircraft

in flight. The project involving this aircraft contributed

significantly to the flight

control system on the space

shuttles by testing and getting

the bugs out of the IBM AP101 used on the shuttles and

by helping the Dryden Flight

Research Center to develop a

pilot-induced oscillation

(PIO) suppression filter that

reduced the likelihood of

pilots overcontrolling the

shuttles on landings and

thereby creating excursions

from the intended landing

path. (NASA photo EC77-

6988)

11
Boeing 777—not to mention the X-29

and X-31 research aircraft. Some of these

aircraft would be uncontrollable without

DFBW technology, which is not only

lighter than mechanical systems but

provides more precise and better maneuver control, greater combat survivability,

and for commercial airliners, a smoother

ride.15

While the F-8 Digital Fly-By-Wire

project was still ongoing, Dryden

hosted the Approach and Landing Tests

for the Space Shuttle in 1977. These

involved testing the 747 Shuttle Carrier

Aircraft (SCA) modified to carry the

Shuttle back to its launch location at

the Kennedy Space Center in Florida

(following Shuttle landings on the

Rogers Dry Lakebed), plus flying the

Shuttle prototype Enterprise mated to

the 747, both without and with a crew

on the Shuttle, and then five free flights

of the Enterprise after it separated from

the SCA, including four lakebed landings and one on the regular runway at

Edwards. Flying without a tail-cone

fairing around the dummy main engines

as well as landing on the smaller

runway for the first time, astronaut and

former Dryden research pilot Fred

Haise was keyed up and overcontrolled

the orbiter on the concrete-runway

landing, resulting in a pilot-induced

oscillation. Once he relaxed his controls, the Enterprise landed safely, but

only after some very uneasy moments.16

15 James E. Tomayko, “Digital Fly-by-Wire: A Case of Bidirectional Technology Transfer,” Aerospace Historian (March

1986), pp. 10-18, and Computers Take Flight: A History of NASA’s Pioneering Digital Fly-By-Wire Project (Washington,

DC: NASA SP-2000-4224, 2000); Proceedings of the F-8 Digital Fly-By-Wire and Supercritical Wing First Flight’s 20th

Anniversary Celebration (Edwards, CA: NASA Conference Publication 3256, 1996), Vol. I, esp. pp. 4, 15, 19-20, 34,

46-51, 56; oral history interview, Lane Wallace with Kenneth J. Szalai and Calvin R. Jarvis, Aug. 30, 1995, transcript in

the Dryden Historical Reference Collection. See also Lane Wallace’s account in Flights of Discovery, pp. 111-118.

16 Space Shuttle Orbiter Approach and Landing Test, Final Evaluation Report (Houston, TX: NASA JSC-13864. 1978).

For shorter, less technical descriptions, see Hallion, On the Frontier, pp. 242-250, and Wallace, Flights of Discovery, pp.

134-137.

The Space Shuttle

prototype Enterprise flies free after

being released from

NASA’s 747

Shuttle Carrier

Aircraft (SCA)

during one of five

free flights carried

out at the Dryden

Flight Research

Center, Edwards,

California, as part

of the Shuttle

program’s Approach and Landing

Tests (ALT).

(NASA photo

ECN-8611)

12
This posed a hazard for Shuttle landings

from space, because if a keyed-up pilot

overcontrolled, the results might be more

dangerous. There needed to be a correction to the Shuttle’s flight control system.

So the F-8 DFBW and other aircraft were

pressed into service to find a solution.

Dryden engineers suspected the problem

lay in the roughly 270-millisecond time

delay in the Shuttle’s DFBW flight

control system, so pilots flew the F-8

DFBW research airplane with increasing

time delays to test this belief. When the

experimental time-delay reached 100

milliseconds, Dryden research pilot John

Manke was doing a touch-and-go landingtake-off sequence and entered a severe

pilot-induced oscillation at a high angle of

attack and low speed. As the aircraft

porpoised up and down in increasingly

severe oscillations, hearts stopped in the

control room until fellow F-8 pilot Gary

Krier reminded Manke to turn off the time

delay, allowing him to climb to a safe

altitude. The control room remained in a

stunned silence until Gary keyed up the

mike again and said, “Uh, John, I don’t

think we got any data on that; we’d like

you to run it again.” The ensuing

laughter broke the tension. As a result

of this and 13 other flights in March

and April 1978, Dryden engineers had

the data they needed to suppress control-surface action resulting from

excessive pilot inputs. A suppression

filter reduced the probability of a pilotinduced oscillation without affecting

normal flying qualities, contributing to

the safe landings of the Shuttle ever

since.17

17 On this flight research, see especially Wallace’s intvw. with Szalai and Jarvis; Tomayko, “Digital Fly-by-Wire,” p. 17,

and Computers Take Flight, pp. 113-114; and Wallace, Flights of Discovery, p. 137.

This 1964 NASA

Flight Research

Center photograph shows the

Lunar Landing

Research Vehicle

(LLRV) Number

1 in flight at the

South Base of

Edwards Air

Force Base.

(NASA photo

ECN-506)

13
A much earlier contribution to the

nation’s space effort was the Lunar

Landing Research Vehicle (LLRV). When

Apollo planning was beginning in 1960,

NASA began looking for a simulator to

emulate the descent to the moon’s surface. Three projects developed, but the

most important was the LLRV developed

by Bell Aerosystems in partnership with

the Flight Research Center. Two LLRVs

paved the way for three Lunar Landing

Training Vehicles (LLTVs) supplemented

by the LLRVs, which were converted into

LLTVs.

Ungainly vehicles humorously called

“flying bedsteads,” they simulated the

moon’s reduced gravity on descent by

having a jet engine provide five-sixths of

the thrust needed for them to stay in the

air. A variety of thrusters then handled the

rate of descent and provided control. The

vehicles gave the Apollo astronauts a

quite realistic feel for what it was like to

land on the Moon. Neil Armstrong said

that he never had a comfortable moment

flying the LLTVs, and he crashed one

of the LLRVs after it was converted to

an LLTV, escaping by means of the

ejection system. But he said he could

not have landed on the Moon without

the preparation provided by the

LLTVs.18

Another very important contribution to

the Shuttles and probably to future

spacecraft came from the lifting bodies.

Conceived first by Alfred J. Eggers and

others at the Ames Aeronautical Laboratory (now the Ames Research Center),

Mountain View, California, in the mid1950s, a series of wingless lifting shapes

came to be flown at what later became

Dryden from 1963 to 1975 in a joint

program with the Air Force, other NASA

centers, and both Northrop and Martin on

the industrial side. They included the

M2-F1, M2-F2, M2-F3, HL-10, and

X-24A and B. Flown at comparatively

low cost, these low lift-over-drag vehicles

18 On the LLRVs and LLTVs, see Donald R. Bellman and Gene J. Matranga, Design and Operational Characteristics of

a Lunar-Landing Research Vehicle (Washington, DC: NASA TN D3023, 1965) and Hallion, On the Frontier, pp. 140-

146.

The HL-10

landing on the

lakebed with an

F-104 chase

aircraft. (NASA

photo ECN2367)

14
demonstrated both the viability and versatility of the wingless configurations and their

ability to fly to high altitudes and then to

land precisely with their rocket engines no

longer burning. Their unpowered approaches

and landings showed that the Space Shuttles

need not decrease their payloads by carrying

fuel and engines that would have been required

for conventional, powered landings initially

planned for the Shuttle. The lifting bodies

also prepared the way for the later X-33

and X-38 technology demonstrator programs that feature lifting-body shapes to be

used for, respectively, a potential nextgeneration reusable launch vehicle and a

crew return vehicle from the International

Space Station.19

A very different effort was the F-8

Supercritical Wing flight research project,

19 For the details of this remarkable program, see R. Dale Reed with Darlene Lister, Wingless Flight: The Lifting Body

Story (Washington, DC: NASA SP-4220, 1997); Milton O. Thompson with Curtis Peebles, Flight without Wings: NASA

Lifting Bodies and the Birth of the Space Shuttle (Washington, DC: Smithsonian Institution Press, 1999).

This photo

shows the M2-F3

Lifting Body

being launched

from NASA’s B52 mothership at

the NASA Flight

Research Center.

(NASA photo

EC71-2774)

The X-24B

landing on the

lakebed with an

F-104 safety

chase aircraft.

(NASA photo

EC75-4914)

15
conducted at the Flight Research Center

from 1971 to 1973. This project illustrates an important aspect of flight

research at what is today Dryden

because the design was the work of Dr.

Richard Whitcomb at the Langley

Research Center and resulted from his

insights and wind-tunnel work. Frequently, projects flown at Dryden have

resulted from initiatives elsewhere in

NASA, in the armed services, in industry, or other places. However, researchers often discover things in flight that

were only dimly perceived—or not

perceived at all—in theoretical and

wind-tunnel work, and flight research

also can convince industry to adopt a new

technology when it wouldn’t do so as a

result of wind-tunnel studies alone. In this

case, Larry Loftin, director of aeronautics

at Langley, said, “We’re going to have a

flight demonstration. This thing is so

different from anything we’ve ever done

before that nobody’s going to touch it

with a ten-foot pole without somebody

going out and flying it.”20

In this case, although there was some

discovery resulting from the flight research—e.g., that there was some laminar

flow on the wing that was not predicted,

in addition to the numerous discrepancies

Milt notes in his account below—generally there was good correlation between

wind-tunnel and flight data. The SCW had

increased the transonic efficiency of the

F-8 by as much as 15 percent, equating to

savings of $78 million per year in 1974

dollars for a 280-passenger transport fleet

of 200-passenger airplanes. As a result of

this study, many new transport aircraft

today employ supercritical wings. Moreover, subsequent flight research with

supercritical wings on the F-111 showed

that the concept substantially improved a

fighter aircraft’s maneuverability and

performance.21

A final project that should be mentioned

here is the research with the three-eighthsscale F-15/Spin Research Vehicle. This

was a sub-scale remotely piloted research

vehicle chosen because of the risks

20 Ted Ayers, “The F-8 Supercritical Wing; Harbinger of Today’s Airfoil Shapes,” Proceedings of the F-8 . . .

Supercritical Wing, pp. 69-80, and Richard Whitcomb, “The State of Technology Before the F-8 Supercritical Wing,”

ibid., pp. 81-92, quotation from p. 85.

21 Ayers, “Supercritical Wing,” p. 78; Whitcomb, “State of Technology,” pp. 84, 90; Hallion, On the Frontier, pp. 202-

208; Wallace, Flights of Discovery, pp. 90-92.

The F-8

Supercritical

Wing aircraft in

flight. (NASA

photo EC73-

3468)

16
involved in spin testing a full-scale fighter

aircraft. The remotely piloted research

technique enabled the pilot to interact

with the vehicle as he did in normal flight.

It also allowed the flight envelope to be

expanded more rapidly than conventional

flight research methods permitted for

piloted vehicles. Flight research over an

angle-of-attack range of -20 degrees to

+53 degrees with the 3/8-scale vehicle—

during its first 27 flights through the end

of 1975 in the basic F-15 configuration—

allowed FRC engineers to test the mathematical model of the aircraft in an

angle-of-attack range not previously

examined in flight research. The basic

airplane configuration proved to be

resistant to departure from straight and

level flight, hence to spins. The vehicle

could be flown into a spin using techniques developed in the simulator,

however. Data obtained during the first

27 flights gave researchers a better

understanding of the spin characteristics of the full-scale fighter. Researchers later obtained spin data with the

vehicle in other configurations at angles

of attack as large as –70 degrees and +88

degrees.

There were 36 flights of the 3/8-scale F-15s by

the end of 1978 and 53 flights by mid-July of

1981. These included some in which the

vehicle—redesignated the Spin Research

Vehicle after it was modified from the basic

F-15 configuration—evaluated the effects of

an elongated nose and a wind-tunnel-designed

nose strake (among other modifications) on the

airplane’s stall/spin characteristics. Results of

flight research with these modifications

indicated that the addition of the nose

strake increased the vehicle’s resistance

to departure from the intended flight

path, especially entrance into a spin. Large

differential tail deflections, a tail chute, and

a nose chute all proved effective as spin

recovery techniques, although it was

essential to release the nose chute once it

had deflated in order to prevent an inadvertent reentry into a spin. Overall, remote

piloting with the 3/8th-scale F-15 provided

high-quality data about spin.22

22 Kenneth W. Iliff, “Stall/Spin Results for the Remotely Piloted Spin Research Vehicle,” AIAA Paper No. 80-1563

presented at the AIAA Atmospheric Flight Mechanics Conference, Aug. 11-13, 1980; Kenneth W. Iliff, Richard E.

Maine, and Mary F. Shafer, “Subsonic Stability and Control Derivatives for an Unpowered, Remotely Piloted 3/8-Scale

F-15 Airplane Model Obtained from Flight Test,” (Washington, DC: NASA TN D-8136, 1976).

This photograph

shows NASA’s

3/8th-scale

remotely piloted

research vehicle

landing on

Rogers Dry

Lakebed at

Edwards Air

Force Base,

California, in

1975. (NASA

photo ECN4891)

17
In these and many other projects, what is

today the Dryden Flight Research Center

has shown that while theory, ground

research facilities, and now Computational

Fluid Dynamics are critical for the design of

aircraft and for advancing aeronautics,

flight research is also indispensable. It

serves not only to demonstrate and validate

what ground research facilities have

discovered but also—in the words of Hugh

Dryden—to “separate the real from the

imagined . . .” and to discover in flight

what actually happens as far as instruments

and their interpretation will permit.23 This

essential point is reemphasized in Milt’s

study from his own particular perspective,

but his account also contains a great deal

more that practitioners of flight research

today—and perhaps even ground researchers—would do well to heed.

23 For the quotation, Hugh L. Dryden, “General Background of the X-15 Research-Airplane Project,” in the NACA,

Research-Airplane-Committee Report on Conference on the Progress of the X-15 Project (Langley Field, VA: Compilation of Papers Presented, Oct. 25-26, 1956): xix. Dryden’s comment related specifically to the X-15 but has more general

applicability. On the need for interpretation of data from instruments, see Frederick Suppe’s interesting “The Changing

Nature of Flight and Ground Test Instrumentation and Data: 1940-1969” on the Internet at http://carnap.umd.edu:90/

phil250/FltTest/FltTest1.pdf. Of course, with the use of lasers in a variety of applications today to augment more traditional instrumentation, and with careful calibration of instruments as well as the use of instruments from different

manufacturers in the same general location on an aircraft, there is less room for assumption and interpretation as well as

for theoretical models to bias the understanding of flight research data than otherwise would be the case. But whenever

aeronautical researchers use instruments in an experimental environment, there is always a need to spend a lot of time

understanding what those instruments measure and how they do it to ensure accuracy in using data from them.

Portrait of Dr.

Hugh L. Dryden a

couple of years

after he made the

remark quoted in

the narrative.

(NASA photo E4248)

18
Introduction

The NASA Flight Research Center

(FRC—formerly the NACA High-Speed

Flight Station [and now known as the

NASA Dryden Flight Research Center])

has been involved in experimental and

research flight testing for over 27 years.

FRC’s experience began with the X-1

series of aircraft and extended through all

the manned X-series aircraft, the D-558

series, and most recently the lifting

bodies. Other experience was also gained

with unusual vehicles such as paragliders

and the Lunar Landing Research Vehicles.

FRC has flight tested vehicles with

operating speeds ranging from zero to

4,500 miles per hour and altitude ranges

from ground level to 354,000 feet. Over

5,000 research flights have been made in

over 60 different types of research

aircraft. Only three aircraft and two pilots

have been lost during research testing and

none of these losses were attributable to

negligence or inadequate planning or

preparation.24

This is a remarkable record, especially

considering the extremely hazardous

nature of the testing FRC has been

involved in. FRC has, however, had a

number of accidents and incidents not

involving the loss of an aircraft or a

pilot. Numerous problems have been

encountered in flight that were

unpredicted or unanticipated. This, of

course, is the justification for flighttesting. This document will describe

some typical examples of the kinds of

problems we have encountered. The

intent is to make people aware of the

kinds of problems we have encountered

so that these same mistakes will not be

repeated as they have been so often in

the past.

The kinds of problems that we have

encountered can be categorized into

hardware problems, aerodynamic problems, and what might be called environmental problems. Hardware problems are

those where a component or subsystem

does not perform up to expectations. The

component or subsystem doesn’t function

properly or fails completely. Aerodynamic

problems are those encountered because

the wind-tunnel predictions were not

accurate or were misinterpreted or even

inadequate. Environmental problems are

those that show up only in flight. They

generally result from a lack of foresight or

understanding of the effects of the environment on a subsystem or component, or

the vehicle itself.

Of the three types of problems, the

emphasis will be on aerodynamic- and

environmental-type problems. Two

research aircraft have been selected as the

prime examples, the HL-10 and the X-15.

The HL-10 was an unconventional

configuration with state-of-the-art offthe-shelf subsystems. Its problems, as you

might suspect, were aerodynamic in

nature. The X-15 was a relatively conventional configuration but most of its subsystems were newly developed and many

pushed the state of the art. Its problems

were mainly with subsystems. Both vehicles explored new flight regimes.

Aerodynamic Problems

Figure 1 illustrates an example of a

serious problem encountered on the first

flight of the HL-10—flow separation. The

flow separation occurred at the junction of

the tip fin and the fuselage. It occurred in

flight as the pilot began his practice flare

at altitude. When this occurred, the pilot

essentially lost all pitch and roll control.

Problems

Encountered

in Flight

Research

Milt Thompson

24 Milt did not specify, but presumably he meant Howard Lilly’s crash after takeoff due to compressor disintegration in

the D-558-1 No. 2 on 3 May 1948, which resulted in Lilly’s death; the crash of the M2-F2 without loss of life on 10 May

1967; and Michael Adams’ fatal accident in X-15 No. 3 on 15 Nov. 1967. Although badly damaged, the M2-F2 was not

lost and was rebuilt with a center fin to make it more stable and a more successful research airplane. This list does not

include the deaths of Air Force Maj. Carl Cross and NASA pilot Joe Walker as a result of a mid-air collision between an

XB-70A and an F-104N in 1966 because that did not occur as part of a research flight.

19
He had almost full right aileron in, and he

was rolling slowly to the left. In this case,

we were lucky. The vehicle recovered by

itself since the flow separation also

caused a nose-down pitching moment that

lowered the angle of attack, causing the

flow to reattach. A detailed reassessment

of the wind-tunnel data revealed some

slight evidence of a potential separation

problem at the flight conditions that

produced it; however, a substantial

amount of additional wind-tunnel testing

was required to confirm this and define a

fix.

On that same flight we had longitudinalcontrol-system limit-cycle and sensitivity

problems. The pilot used only one inch of

longitudinal stick deflection from flare

initiation at 300 knots to touchdown at

200 knots. The sensitivity and controlsystem limit-cycle problems were primarily a result of the elevon effectiveness

being higher than anticipated. I say

“anticipated” rather than “predicted”

because the measured effectiveness

compared quite well with that measured

in the small-scale wind tunnel; however,

we had chosen to believe the full-scale

wind-tunnel results. This is an interesting

case since the full-scale wind-tunnel data

were obtained using the actual flight

vehicle as the model, and the Reynolds

number range was from 20 to 40 million.

The small-scale model was a 0.063-scale

model (16 inches long), and the Reynolds

number range was an order of magnitude

lower—2 to 4 million. Flight Reynolds

numbers ranged from 40 to 80 million.25

One might question whether it could be a

compressibility effect, but one wouldn’t

25 Reynolds number, named after Osborne Reynolds, is a non-dimensional parameter equal to the product of the velocity

of, in this case, an airplane passing through a fluid (air in this instance), the density of the fluid, and a representative

length, divided by the fluid’s viscosity. In shorthand, this is the inertial forces divided by the viscous forces for the mass

of air acted upon by the vehicle. Among other uses, it served to compare data from wind-tunnel models with that from

full-sized airplanes or components. The Reynolds number was not determined solely by the viscosity of the air. A large

transport aircraft, for example, would have a much larger Reynolds number when flying through air at a given altitude,

location, and time than would a small model simply because of the difference in size and the amount of air displaced.

Furthermore, the Reynolds number would be much larger at the rear of a flight vehicle than at the front.

Figure 1: Schematic

showing flow

separation in the

original HL-10

design and the

modification to the

HL-10 that solved

the problem. This is

not the exact

illustration Milt

intended for this

paper, but it shows

what he was talking

about. (Original

drawing by Dale

Reed; digital

version by the

Dryden Graphics

Office).

20
expect compressibility effects at 0.4 Mach

number. We did, however, see compressibility effects as low as Mach 0.5 on the X24A. It should be noted that flight-measured

longitudinal stability was higher than

predicted by either the small-scale or fullscale tunnels, whose data agreed quite well.

In the case of aileron characteristics, again

the small-scale and full-scale tunnel

results agreed quite well; however, the

flight-measured results were higher than

either, and again we had control-system

limit-cycle and sensitivity problems

during flight. The predicted subsonic

longitudinal trim was off by approximately four degrees in angle of attack due

to a combination of discrepancies in zerolift pitching moment as well as static

stability and control effectiveness. Discrepancies in longitudinal trim of roughly

this same magnitude were observed in

each of the lifting bodies.

The HL-10 configuration had over 8,000

hours of wind-tunnel testing. One model

that was tested was actually larger than

the flight vehicle—28 feet long as compared to a 20-foot flight vehicle, or 1.4

scale. The actual flight vehicle was tested

in the 40X80-foot tunnel at Ames Research Center. You couldn’t get better

model fidelity, and yet we still saw

discrepancies between the predicted and

flight-measured data.

Aerodynamic discrepancies were not

restricted to the HL-10 configuration. The

HL-10 was simply used as an example.

Each of the other lifting bodies exhibited

similar kinds of discrepancies between

predicted and flight data. The M2-F2

wind-tunnel tests were conducted and

analyzed by another team of experts

including people such as [Alfred J.]

Eggers, [Clarence] Syvertson, [Jack]

Bronson, [Paul F.] Yaggy, and many

others, and yet again, the predictions were

not perfect. The X-24A configuration was

developed and tested by the Martin

Company for the United States Air Force

(USAF). It was a highly optimized and

finely tuned configuration. The X-24A

designers, for example, detected the

potential for a flow separation problem at

the fin-fuselage juncture and tested over

twenty different fin leading-edge configurations before settling on the final leading

edge for the flight vehicle. As meticulous

as these designers were, we still saw some

slight evidence of unpredicted flow

separation.

On the X-24A, we also observed a

discrepancy in aileron yawing-moment

derivative. In terms of the actual numerical value, the discrepancy was small. In

terms of percentage, it was an error by a

minus 100 percent. In terms of vehicle

handling qualities, the discrepancy was

Figure 2: Controllability boundaries

for the X-24B at

Mach 0.95. A

A

 Adapted and simplified from Christopher J. Nagy and Paul W. Kirsten, “Handling Qualities and Stability Derivatives

of the X-24B Research Aircraft” (Edwards AFB, CA: AFFTC-TR-76-8, 1976), p. 56. It is obvious that this was not

precisely the figure Milt had in mind, but it illustrates his point. Note that Nagy and Kirsten comment on p. 54,

“Although modeling of the rocket exhaust conditions was not exact (hence the resulting data was not considered to be

accurate), the results were used as guidelines to evaluate the potential loss of stability with the rocket engine on.”

They added, “The comparison of the handling qualities boundaries before and after the flight-test program exemplifies

the need for an incremental envelope expansion approach to flight test of new aircraft. Boundaries determined by

actual lateral-directional stability were considerably more restrictive than they were predicted to be. Although poweron wind tunnel test did indicate an effect of the rocket engine, tests of this nature are not conducted for most test

programs.”

21
extremely significant since it caused a pilot

induced oscillation (PIO) in flight.

We have seen some unusual power effects in

each of the lifting bodies. These include

longitudinal trim changes of as much as four

degrees in angle of attack. Most recently, we

have observed a loss of directional stability in

the X-24B in the Mach number range from 0.9

to 1.0 and higher. This is illustrated in Figure 2.

These power effects were not due to thrust

misalignments [although some existed]. They

were the result of rocket-plume induced flow

separation over the aft fuselage, fins, and

control surfaces. This phenomenon is apparently peculiar to lifting-body configurations or

non-symmetrical shapes, since it had not been

noted in earlier rocket aircraft or in missiles to

any significant extent.26

Aerodynamic discrepancies are not limited to

lifting-body configurations. We saw a reversal

of sign in yaw due to aileron on the XB-70 as

illustrated in Figure 3. Aileron characteristics

26 Milt seems to be forgetting here that there were rocket-plume effects in the D-558-2 when any other cylinder of the

XLR-8 rocket engine fired in a combination including the top cylinder. These effects were most severe at the highest

Mach number tested—approximately Mach 1.6. The plume effects were small when only the two middle cylinders fired

together in a horizontal plane. See Chester W. Wolowicz and Herman A. Rediess, “Effects of Jet Exhausts on FlightDetermined Stability Characteristics of the Douglas D-558-II Research Airplane” (Washington, DC: NACA RM

H57G09, 1957), esp. pp. 16-17. There apparently were also plume effects on rockets such as the Saturn V.

Figure 3: Variations

of XB-70-1 flightbased and predicted aileron

yawing-moment

control derivative

with Mach number

in hypothetical

climbout profile. B

B

 Taken from Chester H. Wolowicz, Larry W. Strutz, Glenn B. Gilyard, and Neil W. Matheny, “Preliminary Flight Evaluation of the

Stability and Control Derivatives and Dynamic Characteristics of the Unaugmented XB-70-1 Airplane Including Comparisons with

Predictions” (Washington, DC: NASA TN D-4578, 1968), p. 64. This may not have been the precise figure Milt had in mind, but it

illustrates his point, showing that the predicted aileron yawing-moment control derivative was positive (proverse), whereas the flightbased values were negative (adverse) from a Mach number of about 0.90 through the supersonic range.

Figure 4: Calculated decrement/

increment of lift-todrag ratio resulting

from the difference

between predicted

and measured base

pressure coefficients in the XB70. Only the lift-todrag ratio increment in the shaded

regions is used for

range-increment

calculations. C

C

 Taken from Edwin J. Saltzman, Sheryll A. Goecke, and Chris Pembo, “Base Pressure Measurements on the XB-70 Airplane

at Mach numbers from 0.4 to 3.0” (Washington, DC: NASA TM X-1612, 1968), p. 31. Again, this may not have been the

exact figure Milt intended to use, but it makes his point. Notice that there was a favorable increment in lift-to-drag ratio at

22
of delta-wing aircraft have historically

been hard to predict since the days of the

XF-92, one of the first delta-wing aircraft.

A discrepancy in aileron characteristics

may not seem too significant, and yet an

aircraft (a B-58) was lost during the

flight-test program because of this

particular error in prediction.

The B-70 drag discrepancy shown in

Figure 4 resulted in a 50-percent reduction in predicted range. This is an excellent example of a discrepancy that

couldn’t be completely resolved even

after the fact. It was suspected that the

discrepancy was primarily due to flexibility or aeroelastic effects. After the flight

program was completed, a new model

was constructed and tested in an attempt

to get better correlation between windtunnel and flight data. The best correlation that could be obtained—even knowing the answer in advance—was 10

percent on overall drag, and that still

means a big error in overall flight range.

Aerodynamic discrepancies have not

disappeared with time. During tests of the

F-8 Supercritical Wing in the 1971-72

time frame, we saw numerous discrepancies between wind-tunnel and flight data

even at the optimized design cruise

condition of 0.99 Mach number. Figure 5

shows comparisons of wind-tunnel and

flight data for some of the aerodynamic

derivatives where significant discrepancies occurred. Admittedly, the design

Mach number region is extremely hard to

work in. Yet the 50- to 100-percent errors

in such basic stability derivatives as the

one for sideslip could hardly be considered acceptable accuracies. At other than

design cruise condition, a large discrepancy was observed in aileron effectiveness and smaller but still significant

discrepancies in the pitching-moment

coefficients.

Here again, we had a master of the craft,

[Richard] Whitcomb, conducting the

wind-tunnel tests and analyzing the

results before the fact. Admittedly, the

airfoil concept was somewhat revolutionary; however, Whitcomb had essentially

unlimited access to any wind-tunnel

facility he needed and should therefore

cruise speeds above Mach 2.5 but that at low supersonic speeds near Mach 1.2 there was the very unfavorable decrement

Milt talks about. Thus, even though ground researchers had overestimated base drag at cruise speeds, their underestimate at

low supersonic climbout speeds seriously reduced the aircraft’s range.

Figure 5: Lateraldirectional derivatives as a function

of angle of attack in

the F-8 Supercritical Wing

aircraft. D

D Taken from Neil W. Matheny and Donald H. Gatlin, “Flight Evaluation of the Transonic Stability and Control Characteristics of an Airplane Incorporating a Supercritical Wing” (Edwards, CA: NASA Technical Paper 1167, 1978), pp. 42, 43, 46.

Once more, this may not be the precise illustration Milt intended to use, but it shows roughly the level of discrepancy

between wind-tunnel and flight data that he had in mind and does so for some of the derivatives he mentions.

23
not have any good excuse other than the

fact that the wind tunnels still have some

obvious shortcomings.27

More recently, discrepancies in very basic

stability characteristics have been observed

in the newest and latest aircraft. The

[Y]F-16 and [Y]F-17 showed a substantial

difference between predicted and flightmeasured longitudinal stability throughout a

major portion of the usable angle-of-attack

envelope.28 The B-1 exhibited much more

adverse yaw due to roll control on its first

flight than had been predicted. This discrepancy showed up on a configuration and at a

flight condition that should have been

highly predictable.

In summation, we just haven’t seen evidence to prove that wind-tunnel predictions

are improving that much in accuracy or

that we have gotten that much smarter

in anticipating all the potential aerodynamic problems.29

Environmental Problems

To turn to the subject of environmental

problems, I would like to review some that we

experienced with the X-15. There have been

numerous reports published and many papers

given on the results of the X-15 flight program,

but nothing has been published that summarized all the problems we had. We went back

into the records to try to identify all the various

problems. Before discussing them, however,

we must recognize that the X-15 was quite an

advanced aircraft for its time, except in terms

of its configuration. This was pretty conven27 Note that Whitcomb discussed some of the preliminary differences between wind-tunnel and flight data in his “Comments on Wind-Tunnel-Flight Correlations for the F-8 Supercritical Wing Configuration,” in Supercritical Wing Technology: A Progress Report on Flight Evaluations (Washington, DC: NASA SP-301, 1972), pp. 111-120, a report that was

still classified when Milt was writing this document.

29 If Milt were writing today, he would no doubt add the results of Computational Fluid Dynamics (CFD) to his comments,

since the results of CFD have also failed to anticipate many potential aerodynamic problems in vehicles that have used it as a

design tool. On the other hand, many people would argue that wind-tunnel predictions have improved significantly, partly as

a result of comparing previous predictions with the actual results of flight research, partly from other sources.

28 The YF-16 and YF-17 were in a very close competition for an Air Force contract, which the YF-16 won in January 1975,

and this led to the production F-16As—a fact that Milt could not have known at the time of his writing this document. The

YF-17 later led to the Navy/Marine Corps F/A-18. See the Air Force Flight Test Center History Office’s Ad Inexplorata: The

Evolution of Flight Testing at Edwards Air Force Base (Edwards AFB, CA: AFFTC/HO, 1996), pp. 27-28.

tional as can be seen in Figure 6 with the

possible exception of the upper and lower

vertical tails, which were wedge-shaped. The

aircraft had a unique structure for dealing with

aerodynamic heating, and it featured many

Figure 6: Threeview and cutaway

drawings of the

X-15. E (See

page 25)

24
new systems that were required to fly to the

limits of the flight envelope. These included

the reaction control system, the inertial system,

the LR-99 rocket engine with throttling, the

skid [landing] gear, auxiliary power units, sidearm controller, the ball nose to provide air data,

and the MH-96 Flight Control System (a rate

command system with adaptive gain [that

appeared only in the X-15 Number 3]).

As you might suspect and as will be discussed

later, our major problems were with the

systems rather than with configuration

aerodynamics. In most areas, the aerodynamics

were pretty much as predicted. There was

good correlation between wind-tunnel and

flight data throughout the entire Mach range.

The only significant difference was in base

drag, which was 50 percent greater than

predicted. Again, a characteristic historically

hard to predict. The lift-to-drag ratio (L/D),

however, was higher than predicted—4.5 as

compared to 4.2—which indicates that there

were compensating factors not evident in the

wind-tunnel data. Ground-effect and geardown L/D were also inaccurately predicted.

One other important bit of data obtained during

the X-15 flight program was aerodynamic

heating data, which revealed that actual heat

transfer rates were substantially lower than

predicted by theory.30

Figure 7 addresses the X-15 program and

some of the problems encountered. It

E

 This was taken from Wendell H. Stillwell, X-15 Research Results with a Selected Bibliography (Washington, DC: NASA

SP-60, 1965), p. 3.

30 Another inaccurate prediction stemmed from the theoretical presumption that the boundary layer (the thin layer of air close to the

surface of an aircraft) would be highly stable at hypersonic speeds because of heat flow away from it. This presumption fostered the

belief that hypersonic aircraft would enjoy laminar (smooth) airflow over their surfaces. Because of this, many designers computed

performance and heating for the hopeful case of laminar flow. At Mach 6, even wind-tunnel extrapolations indicated extensive

laminar flow. However, flight data from the X-15 showed that only the leading edges of the airfoils exhibited laminar flow and that

turbulent flow occurred over the entire fuselage. Small surface irregularities, which produced turbulent flow at transonic and supersonic speeds, did so equally at speeds of Mach 6. Thus, designers had to abandon their hopeful expectations. On this matter, see John

V. Becker, “The X-15 Program in Retrospect,” 3rd Eugen Sänger Memorial Lecture, Bonn, Germany, Dec. 4-5, 1968, pp. 8-9; Albert

L. Braslow, “Analysis of Boundary-Layer Transition on X-15-2 Research Airplane” (Washington, DC: NASA TN D-3487, 1966).

Figure 7: X-15

Flight Program,

found in the Milt

Thompson Collection of the Dryden

Historical Reference Collection.

Note that M=Mach

number, k=1,000,

Max q=maximum

dynamic pressure,

and

hmax=maximum

altitude.

25
indicates some milestones in the program

and correlates them with the flight

numbers. There was a total of 199 flights

made with the three aircraft. In the middle

of the figure, the various phases of the

program are shown. On the bottom of the

figure, some of the problem areas are

listed. The bars and dashed lines indicate

where the problems occurred during the

program. The solid bar indicates continuing significant problems. The dashed lines

indicate continuing minor problems, and

the asterisks represent unique problems.

The Number Two aircraft was severely

damaged on its 31st flight—the 74th X-15

flight of the program as a whole—and

was subsequently rebuilt and modified to

achieve higher performance.31 It began

flying again shortly after the halfway

point in the program as shown in the test

phase part of the figure.

31 On the 9 Nov. 1962 flight, Jack McKay could not get the XLR-99 engine to advance its throttle setting beyond 30 percent and had

to make an emergency landing at Mud Lake under X-15 mission rules. He was unable to complete his jettison of propellants after

Figure 8: Time

history of the flare

and touchdown of

X-15-1 on its first

flight. F

F

This was taken from Thomas W. Finch and Gene J. Matranga, “Launch, Low-Speed, and Landing Characteristics Determined from

the First Flight of the North American X-15 Research Airplane” (Washington, DC: NASA TM-195, 1959), Fig. 13 on p. 26. This

probably is the figure Milt had in mind to illustrate his point. On pp. 9-10, Finch and Matranga state:

From [the] figure it is obvious that a severe pitching oscillation was induced near the end of the flap cycle. Reduced

longitudinal trim was required as the flaps were being deflected, and the pilot added further airplane nose-down trim to

avoid flaring too high. Apparently the oscillation became more severe because of the control input at about 18 seconds

before touchdown. From this point, the pilot was not able to anticipate the oscillation accurately, which may have been

aggravated by the fact that the control surface was rate-limited to 15° per second. . . . The transient in pitch covered an

angle-of-attack range from -1° to 13°, with the amplitude as high as ±5°.

26
The envelope expansion to design speed,

altitude, and dynamic pressure concluded

with the 53rd flight—rather early in the

program. It didn’t take many flights to

achieve these design conditions when one

considers that 30 of the first flights were

made with the interim engine—the LR-11

[two of which flew on each flight], which

limited the maximum performance to

about Mach 3. Once the LR-99 engine

was available, the flight envelope was

rapidly expanded—roughly half a Mach

number at a time to the design speed of

Mach 6, and 30,000 feet at a time to the

design altitude of 250,000 feet. After

achieving the design conditions, we began

exploring the total flight envelope and

continued to expand the altitude envelope,

finally achieving an altitude of 354,[200]

feet. We had the total impulse available to

go even higher; however, the reentry was

becoming somewhat critical. We also

began exposing the aircraft to greater heat

loads, going to high Mach numbers at

lower and lower altitudes. We also began

carrying piggyback experiments on the

aircraft before the 80th flight and from the

130th flight on. That’s essentially all that

the X-15s were used for after that point

since we had completed the basic aircraft

flight-test program.32

Control-System Problems

The first major flight problem we had was

with the control system, and this occurred

on the first flight. The pilot got into a PIO

during the landing flare. Very simply, the

PIO was due to the limitation of the

horizontal stabilizer to 15 degrees per

second of surface rate and the pilot was

asking for more than 15 degrees per

second as illustrated in Figure 8. The

airplane was almost lost on the first flight

as a result of this.

The PIO was a surprise because the

simulation used to define the maximum

control surface rate requirement did not

adequately stimulate the pilot to get his

own personal gain up. In the real environment on the first flight, his gain was way

up. He was really flying the airplane. Our

experience has verified that the pilot

generally demands the maximum control

surface rates for a given vehicle in the

period just prior to touchdown, at least for

unpowered landings.

In retrospect, this isn’t hard to understand.

Just prior to touchdown, the pilot is trying

to control the flight path to within onehalf a degree or so to make a good

landing, five feet per second or less. An

unpowered landing, in our opinion, is one

of the most demanding tasks required of a

pilot and a flight-control system. The

problem is that you can’t adequately

simulate it. Visual simulators don’t have

the necessary resolution near the ground,

and even sophisticated flight simulators

such as variable stability aircraft can’t

seem to get the pilot’s personal gain up

sufficiently to thoroughly assess a potential PIO problem in landing. A PIO

problem may not be evident until the first

real unpowered landing is made. Even

with a successful first landing one can’t

be sure the problem doesn’t exist, since

we have found that individual pilot gain

varies considerably and another pilot may

induce a PIO. The control system of the

X-15 was modified after the first flight to

increase the horizontal control surface rate

from 15 degrees per second to 25 degrees

per second.

shutting down the engine, and the excess weight caused him to be high on airspeed. He touched down at 296 miles per hour rather

than the normal 230. The result subjected the main gear to both a rebound and a high aerodynamic load, causing the left landing gear

to collapse, and eventually the aircraft flipped over on its back, injuring McKay and causing the Number Two aircraft to be rebuilt

and modified. See Milton O. Thompson, At the Edge of Space: The X-15 Flight Program (Washington and London: Smithsonian

Institution Press, 1992), pp. 227-230 and his further discussion of this flight below in this study.

32 On the other hand, it could be argued that the hypersonic aircraft itself was the primary experiment from flight 1 to

flight 199, even when it was carrying piggyback experiments.

27
The next major control-system problem

didn’t show up until the 23rd flight. The

problem was a structural resonance

problem wherein the Stability Augmentation System (SAS) was responding to the

vibration of the structure on which the

SAS box was mounted. This self-sustaining control-system problem almost shook

the airplane apart during an entry from

169,000 feet. We subsequently added a

notch filter to eliminate this problem. It

surprised us because we did not conduct a

structural resonance test. The X-15 SAS

was one of the first high-gain, highauthority systems capable of responding

to structural frequencies. Since that

occurrence, we always conduct resonance

tests of an aircraft with SAS on to look

for such problems.

Structural Problems

We also had basic structural problems. On

the 4th flight, one of the thrust chambers

exploded during engine start, causing

engine damage and a fire. The pilot shut

down all the thrust chambers and jettisoned fuel before making an emergency

landing on Rosamond Dry Lake. He was

unable to jettison all the propellant

because of the steep nose-down attitude.

As a result, the aircraft broke behind the

cockpit on nose-gear touchdown.

The aircraft designers had failed to

anticipate the nose-down jettison

problem. The aircraft were subsequently beefed up to handle this problem.

Landing-Gear Problems

Landing-gear problems plagued us

throughout the X-15 flight program.

The landing gear failed on the first

landing. The landing gear was reworked

and performed satisfactorily until the

74th flight. On that flight, after launch,

the engine would only develop 30

percent thrust. The pilot was told to

shut down the engine, jettison propellants, and make an emergency landing

Figure 9: Typical

X-15 landing using

wing flaps. Nosegear touchdown at a

time interval

between initial

main-gear contact

and nose-gear

contact of 1.35

seconds (flight 1-

30-51 on June 27,

1962). Taken from

Richard B. Noll,

Calvin R. Jarvis,

Chris Pembo, and

Betty J. Scott,

“Aerodynamic and

Control-System

Contributions to the

X-15 Airplane

Landing Gear

Loads” (Washington, DC: NASA TN

D-2090, 1963), p.

26.

at the launch lake. Again, the pilot was

unable to jettison all the propellants

and, to compound the problem, the

landing flaps did not extend when

selected. The main gear failed shortly

after touchdown and subsequently, the

nose gear failed and the aircraft ended

up on its back.

This gear failure resulted primarily

from the high-speed touchdown due to

the flap failure, and the high gross

weight. Touchdown speed was almost

300 miles per hour. At main-gear

touchdown, with skid-type gear, the

nose tends to slam down rather rapidly.

28
As the nose starts to pitch down, the

SAS applies nose-up elevator to counteract the nose-down pitching moment.

The airload at this high speed, resulting

from the extreme deflection of the horizontal stabilizer located immediately

above the main landing gear, plus the

airload due to the negative three-point

aircraft attitude, added to the normal

rebound load from nose-gear impact, was

sufficient to break the main landing gear.

A typical time history of loads on the

main landing gear is shown in Figure 9.

The air load problem due to SAS response

was not fully appreciated in the initial

design. A squat switch was later included

to deactivate the SAS on main-gear

touchdown. The squat switch worked

quite well, but as the airplanes gained

additional weight during the program due

to added instrumentation, add-on experiments, and required modifications,

additional fixes were required. The pilots

were first asked to push forward on the

stick at touchdown to relieve the air loads

on the main landing gear. Later, a stick

pusher and a third skid were added to

prevent landing-gear failure. We were still

having gear problems when the program

ended after nearly 200 flights.

Aerodynamic Heating Problems

We had a number of problems associated with aerodynamic heating. They

began showing up as we intentionally

subjected the airplanes to high heating

rates and temperatures. We had two

windshields shatter, becoming completely opaque as shown in Figure 10,

and four that cracked during flight. The

shattering was due to failure of the

glass itself at the high temperatures. An

inappropriate choice of material was the

cause. The cracking was due to distortion of the window frame at high

temperatures. The support structure for

the windshield glass was finally redesigned.

We had a problem with canopy seals.

When the cabin was pressurized, the

canopy leading edge deflected up just

enough to allow the air to get to the

canopy seal. At speeds above Mach 3, the

air was hot enough to burn the seal,

resulting in the loss of cabin pressure. The

fix for this was to add a lip over the front of

the canopy leading edge that prevented the

air from impinging on the canopy seal.

We had a problem with local heating on the

wing leading edges. Expansion gaps in the

wing leading edge were designed to allow

for the expansion due to aero[dynamic]

heating. These gaps, however, triggered

turbulent flow, which caused a hot spot

directly behind the gaps. This caused the

wing skin behind the gap to expand and pop

the rivets holding the skin to the leading

edge. Gap covers were added to eliminate

this problem, but it persisted.

Aerodynamic heating also caused problems

with the landing gear. The first problem was

due to distortion of the nose-gear door. As the

airplane got hot, the nose gear door tended to

bow, opening a gap between the rear lip of the

door and the fuselage skin behind the door.

This allowed ram air to enter the nose gear

compartment. The hot air cut through electrical

wiring and tubing like an acetylene torch. The

nose-gear door and its supporting structure

were finally modified to eliminate this problem.

Another landing-gear problem due to aerodynamic heating resulted in the nose-gear-scoop

Figure 10: A

noteworthy scar

from the X-15’s

first flight to Mach

6 was this cracked

outer panel on the

right side of the

windshield.

(NASA photo

E-7508)

29
door opening at a Mach number of 5.3. The

nose-gear-scoop door was a small door

designed to assist the extension of the nose

gear aerodynamically. When it opened, it

scooped ram air into the nose-gear compartment. At that speed, the air became hot

enough to burn the tires off the nose wheels.

The scoop door released due to distortion of

the uplock linkage system under heating

loads. The uplock system required a

complete redesign.

The scoop door opened another time at

Mach 4.3 on the modified Number Two

X-15 because of a different problem with

the uplock system. Again, the nose-wheel

tires burned and when the pilot extended the

landing gear just prior to touchdown, the

nose gear extended very slowly. A nosegear-up landing was barely averted because

we had a sharp chase pilot who called the

X-15 pilot to hold off until the nose gear

was fully locked. The Number Two aircraft,

which had been modified and rebuilt after a

gear failure that resulted in a roll-over, thus

had other gear problems attributed to aero

[dynamic] heating.

The nose gear extended in flight at Mach

4.3 due to insufficient allowance for

additional structural expansion in the

landing-gear deployment cable system. The

fuselage had been lengthened, but additional compensation for fuselage expansion

had not been included in the landing-gearcable release system. In another incident,

the right-hand main landing gear deployed in flight at Mach 4.5 when the

uplock hook broke as a result of the

bowing of the main-landing-gear strut.

The main landing gear on the modified

Number Two aircraft had been lengthened

to accommodate the supersonic combustion ramjet engine. The additional bending of the longer strut due to differential

heating on the outer and inner portions of

the strut had not been adequately compensated for, and the resulting deflection

in bending of the strut caused the uplock

hook to fail in tension.

Auxiliary-Power-Unit (APU) Problems

We had APU problems during the early

altitude-buildup phase of the program. No

one had thought of pressurizing the APU

gearbox cases. The lubricating oil was

vaporizing at high altitude, and APUs were

failing because of inadequate lubrication.

During a climbout on an altitude flight, the

184th flight,33 one APU shut down because

of an electrical transient that caused an

electrical overload. When the first APU

shut down, the electrical load shifted

automatically to the second APU. The

second APU should have accommodated

the additional load, but because it was

heavily loaded as a result of increased

power demands over the years, it also shut

down. The shutdown of both APUs

resulted in a complete loss of hydraulic

and electrical power as the aircraft was

climbing through 100,000 feet. The

aircraft virtually disappeared. The control

room lost radar tracking, telemetry, and

voice communications. The pilot lost the

engine, all electrically driven instruments,

and all control except for the manual

reaction-control system operated by

cables. He managed to get one APU

restarted to regain hydraulic pressure for

the aerodynamic flight-control system and

successfully reentered the atmosphere

from an estimated 160,000 feet altitude

with no stability augmentation system and

only a couple of instruments—a “g” meter

and his barometric instruments.

MH-96 Problems

The Minneapolis-Honeywell flightcontrol system [MH-96] was fairly

advanced for its time. It was a commandaugmentation-type control system with

adaptive gain scheduling and various

33 This section is moved from a separate heading in the original typescript entitled “Other Problems.” The original said

this was the 154th flight, but as an anonymous reviewer of this publication correctly pointed out, it was the 184th flight,

the 5 being an apparent typographical error.

30
autopilot modes. It was installed in only

the Number Three X-15 for evaluation.

At launch on the first flight, the stability

augmentation portion of the system

completely disengaged due to an electrical transient. The pilot managed to

reengage it, but as you can imagine, it was

quite a shock. Minor problems cropped up

during the next 37 flights of the system.

On the 39th flight [125th flight in the

overall program], shown by an asterisk [in

Figure 7], the system—you might say—

went berserk. The horizontal stabilizers

used for both pitch and roll control started

limit cycling as a result of excessive gain

through a deflection of ±10 degrees.

During this limit cycling, which occurred

at Mach 5.5,34 the aircraft was essentially

out of control in pitch and roll and was

being oscillated by the motion of the

horizontal stabilizers in both pitch and roll

as shown in Figure 11.

It was quite a ride. Luckily the gain

finally came down and the oscillation

stopped. It took a while to find out the

reason for this problem. It turned out that

the system had fooled itself into believing

it was at a flight condition where maximum system gain was required. The

aircraft had been trimmed in a steady-state

4g [acceleration equal to four times the

force of gravity] and the pilot was not

making any inputs. Normal gain-reducing

stimuli for the flight-control system were

totally missing, and the system slowly

drifted to maximum gain. When the pilot

finally made a control input, the system

went unstable. The electronics were

saturated. This same problem later contributed to the structural breakup and loss

of the number three airplane. We did not

implement any specific fix for this problem after the first occurrence.

A cure for the problem was, however,

discovered after some intensive simulator

investigation, and we elected to continue

to fly the system without modification.

This particular problem could not be

duplicated on the Iron Bird simulator after

its initial occurrence. One of the reasons,

we later found, that it could not be duplicated was that the hydraulic pumps

supplying pressure to the Iron Bird

34 It was actually Mach=5.35. Milt probably was giving just a ballpark figure.

Figure 11: X-15

No. 3 on flight 39

(3-39-62), 13

January 1965—the

125th flight of the

X-15 program—in

which the airplane

became uncontrollable in pitch and

roll for a short

time.G

G

The dashed line indicates that the roll and pitch rates exceeded the recorded limits and had to be estimated. Taken from

Euclid C. Holleman, “Control Experiences of the X-15 Pertinent to Lifting Entry,” in Progress of the X-15 Research Airplane

Program (Washington, DC: NASA SP-90, 1965), p. 72. This appears to have been the figure Milt had in mind for this point

in his narrative.

31
simulator could not equal the output of

the aircraft pumps. On the simulator, we

could not physically rate-limit the control

surfaces.

Another reason we could not duplicate the

problem on the simulator was that we

could not cause the system gain to hang

up at maximum gain as it did in flight. It

was finally concluded that the simulator

gain would not hang up because of all the

extraneous electrical noise in the simulation electronics. The electrical noise in the

aircraft system was substantially lower.

We finally managed to duplicate the

problem on the simulator by physically

pinning the system gain at its maximum. Once the problem was understood, the cure was obvious. The pilot

had simply to reduce the gain manually

and the limit cycle would cease. The

pilot flying the Number Three aircraft

on its final flight was aware of this

potential problem and the required

action should it develop. However, for

unknown reasons, he did not take the

proper action.

Fatigue Problems

Even though the total flight program

included only 199 flights and only an

average of 66 flights per airplane, we had

what we considered a couple of fatigue

problems. The 66-flight average in reality

probably involved 200 to 300 system

cycles when you include ground checkouts and aborts. The first fatigue problem

was a rupture of the casing of the engine

turbo-pump. The second was a rupture of

a main bulkhead in one of the propellant

tanks.

X-15 Program Results

With regard to environmental-type

problems, the X-15 program has definitely convinced us of the desirability of a

buildup-type test program when you have

a lot of new systems that you are exposing

to flight for the first time. If we had gone

to the design speed and altitude on the

first flight and had encountered all of the

heating problems and the other subsystem

problems simultaneously, we probably

would have lost the aircraft. Regardless of

all the problems we had, we did make a

lot of successful flights.

The pilots, because they were designed

into all systems, saved many missions and

the aircraft itself on numerous occasions.

Problems notwithstanding, Dr. [Hugh]

Dryden [Director of the NACA and

Deputy Administrator of NASA] referred

to the X-15 flight program as the most

successful research airplane program in

the history of aircraft.

Control-System Problems in General

Flight-control systems are becoming more

and more an integral element of new

aircraft. Even now with the current

generation of aircraft, the control-system

design has in most instances been factored

in to some extent before the configuration

is finalized. It is therefore no longer

practical to allow the aerodynamicist, the

propulsion-systems engineer, and the

structures people to design an aircraft as

they have in the past since now flightcontrol technology has so much to offer.

In a control-configured vehicle or activecontrols-technology vehicle of the future,

the control system will be factored into

the initial design as early and extensively

as the vehicle’s aerodynamic, structural,

performance, propulsion, stability-andcontrol, and other disciplines to achieve

the optimum vehicle. The trend is obvious. Because of this trend, the flightcontrol system assumes much greater

importance in the flight testing of a new

vehicle. These new systems tend to use

higher gains and authorities. They are thus

more susceptible to such things as structural resonance, limit cycling, and surface-rate limiting.

Extensive ground testing of these systems

is required to assure that the control

system itself won’t destroy the aircraft in

flight. The structural-resonance and limit32
cycle problems encountered in flight on

the X-15 are excellent examples of the

seriousness of the problem. It is essential,

for example, that the flight-control system

be active and monitored in all its various

operating modes during ground-vibration

testing. Other ground testing to define

resonance and limit-cycle boundaries is

also a mandatory requirement on any

Flight Research Center flight vehicle.

Ideally, the control surfaces should be

unloaded and then loaded during these tests

to simulate the total hinge-moment environment that any particular surface can expect

to be exposed to. For example, during the

checkout of the M2-F2 for its first flight,

some lead weights were placed on the upper

flaps. The flight-control system was active

at the time, and immediately, a controlsurface oscillation of ±1 degree began. That

particular problem resulted from a slight

deflection of the power-actuator support

structure under load. The support structure

had to be beefed up to eliminate the problem. On another occasion, during limitcycle testing of the HL-10 flight-control

system, a limit cycle was intentionally

induced at approximately 12 cycles per

second. When the limit-cycle stimulation

was terminated, the limit cycle continued.

The stability-augmentation system was then

disengaged in an attempt to stop the limit

cycle, but it still persisted. Shutting off

hydraulic power to the flight-control system

finally stopped the limit cycle. That particular problem was a result of the servoactuator and mechanical-feedback linkage

dynamics.

The individual and combined effects of all

other subsystem operations on the flightcontrol system should be examined as

well as start-up transients of each subsystem. Fly-by-wire control systems will

require additional pre-flight testing to

ensure that no spurious inputs can get into

the system. Lightning-strike problems

have yet to be defined.

Even after performing all this ground

testing, researchers should anticipate and

make provisions for handling potential

problems in flight. For high-gain, highauthority stability-augmentation systems,

it is essential in our opinion that manualgain-changing capability be provided the

pilot during the flight-research program.

The HL-10 limit-cycle problem discussed

earlier, resulting as it did from higherthan-predicted control effectiveness, is a

good example of the need for such

capability in the cockpit. If through

simulation, other potential handlingquality or control problems are identified,

provisions should be made to vary the

questionable parameter. The M2-F2

required a rudder-aileron interconnect to

achieve adequate roll power throughout

the flight envelope. The lateral-directional

handling qualities were extremely sensitive to interconnect ratio as a function of

angle of attack and dynamic pressure.

Because of this and the fact that we could

possibly have had some variations in

predicted control effectiveness and thus,

effective interconnect ratio, we provided

the pilot an adjustable interconnect

control.

Any potential PIO problem observed in

simulation dictates consideration of a

means of reducing stick gearing in

conventional control systems or system

gain in command-augmentation-type

control systems. Stick-gearing-change

provisions, however, are not easily

provided in conventional control

systems. We did not provide this capability in the HL-10, although we knew

from simulation that a control-sensitivity problem might be encountered. As

described earlier, the problem did arise,

and only the skill of the pilot prevented

a potential disaster. In commandaugmentation or fly-by-wire control

systems, effective gearing-change

capability is relatively easy to implement. Thus, there should be no good

excuse for a serious or prolonged PIO

problem in an aircraft with such a

system. Yet they have occurred on first

flights, indicating that someone didn’t

face up to the facts.

33
To reiterate, any PIO potential revealed

through simulation should be taken quite

seriously since we and others have been

caught short so many times in the past.

The absolute PIO potential is extremely

hard to predict even with the most sophisticated moving-base or flight simulator.

One of the major unknowns, as mentioned

previously, is the pilot’s own system gain.

No simulator will stimulate the pilot to

anything approaching a real-life first

flight. On a first flight, the pilot’s personal

gain may be an order of magnitude

greater than anything observed in simulation. To further complicate the problem,

individual pilot gains vary substantially.

One pilot, or even a series of pilots, may

successfully fly a vehicle without a hint of

a problem. Then, all of a sudden a pilot

appears who can compete with the

stability augmentation system in response.

We have seen dramatic evidence of this.

Six research pilots successfully flew the

M2-F1. The seventh pilot, an experienced

test pilot, got into a divergent PIO immediately after takeoff on two successive

attempts to fly the vehicle. The resultant

slow rolls to a landing left even the most

seasoned pilots in the world speechless.

The same pilot, flying a different liftingbody vehicle, was actually able to compete, in response, with the stability

augmentation systems at one cycle per

second.

Command augmentation systems, mentioned earlier, are becoming quite popular.

They are showing up in more and more of

the newer aircraft. One might question

whether they are really needed in some

cases. Command augmentation systems

generally do provide good control characteristics and are quite pleasant to fly. They

do not, however, conform to MIL Specs

[military specifications] in all respects,

and they do have a number of insidious

characteristics. These systems tend to

mask many of the cues the pilot normally

relies on to give information or warning.

For example, a high-gain commandaugmentation-type system tends to

eliminate transients or trim changes due to

gear or flap extension, or center-ofgravity changes. This may not seem

significant, and yet these trim changes in

the past have informed the pilot that the

gear and/or flaps did indeed move when

the appropriate lever was moved or that

the center of gravity was not where you

wanted it. A subtle thing—yet somewhat

disconcerting when you don’t have these

cues.

These same control systems tend to

provide invariant response throughout the

flight envelope. This again would appear

to be highly desirable; however, the

variable response of the older control

system warned the pilot of an approaching

low-speed stall or overspeed just through

feel alone. These new systems feel

completely solid up to and sometimes

over the brink of disaster, and thus

artificial stall-warning systems are

generally required. Speed stability is also

generally lacking in an aircraft equipped

with this type of system, and unless it is

artificially provided, the pilot must

continually monitor airspeed.

Normal dihedral effect and ground effect

are also masked by systems such as these.

A paper discussing many of these insidious characteristics was presented at an

AGARD ([NATO] Advisory Group for

Aerospace Research and Development)

Flight Mechanics Panel meeting in 1967,

sometime prior to the introduction of the

F-111 into operational squadrons.35 Yet at

least one aircraft accident resulted from

each of the insidious characteristics.

Inadvertent high- and low-speed stalls

35 Milt apparently was thinking of his paper with James R. Welsh, “Flight Test Experience with Adaptive Control

Systems,” presented at the AGARD Guidance and Control and Flight Mechanics Panels, Sept. 3-5, 1968, Oslo, Norway,

which was a year later than he remembered and also a year later than when the first F-111As entered service with the

U.S. Air Force, although only in limited numbers.

34
resulted in F-111 aircraft losses. A malfunction in the automatic fuel transfer

system of one F-111 allowed the center of

gravity to move well aft of the normal

flight range and finally sufficiently far aft

to cause the aircraft to diverge. The pilot

of that aircraft was completely unaware of

the problem until the aircraft diverged.

The inadvertent stall-spin problems in other

aircraft equipped with these systems is still

much in evidence. For some reason, the

word is not getting around, and everyone

has to learn the hard way. It’s almost as bad

as the old T-tail problem. The artificial cues

being provided to warn the pilots of impending stall are in many instances completely inadequate, as are many of the stall

inhibitor devices. There is still much more

work to be done in these areas. These

command aug[mentation] systems also in

most cases have to be deactivated in a spin

since they tend to apply improper spin

recovery control.

Adaptive-gain flight-control systems

potentially have problems maintaining

programmed gains. Two different systems

with which we are familiar have had

histories of excessive gain problems and

also too low a gain at times because of

turbulence or other external stimuli that

have not been adequately compensated

for in the initial design.

Control Configured Vehicles (CCV) will

without question be major contenders in

the next generation of military and

possibly commercial aircraft. The first

step has already been taken in the YF-16.

The concept is completely feasible;

however, these control systems must have

the predicted control moments and power

and cannot be marginal on surface rates

or hinge moments.

And, finally, automatic control systems

are not infallible. Automatic flight-control

systems are only as good as the people

who designed them. If the designers have

not anticipated all the possible situations

or flight conditions that the pilot and

aircraft can get into, trouble can result.

Admittedly, it is usually easier to make

the desired or necessary changes through

electronics, but it is still surprising to

realize how many changes are made

during a flight-test program on some of

the newer systems. Twenty to fifty

changes in the flight-control-system

configuration are not uncommon in these

newer systems. The changes required are

in many instances minor changes or

tweaking to optimize the system. We are

aware, however, of some major changes

that were required such as gearing

changes as high as fifty percent of the

original value. The fact that major

changes such as these are required is quite

disturbing since these aircraft are not

exploring new flight regions. Thus, the

predicted aerodynamic data should be

good as far as basic stability and control

derivatives are concerned, and these are

the primary requirements for the design of

a flight-control system. The reason for

such drastic changes being required is

therefore not clear. Somewhere, somehow,

something is being overlooked or not

being considered in the design process.

Finally, the primary message on these new

control systems is to shake, rattle, and roll

them thoroughly before flight and then

expect problems in flight and provide the

necessary system-adjustment capability to

alleviate the problem if it does occur.36

First Flight Preparation

The aerospace industry has had much

more first-flight experience than either

NACA-NASA or the military. It has been

only recently that the government has

been directly involved in preparing a

36 Despite the fact that this section, like the rest of this document, was written about 1973 or 1974, much of what Milt

says is still applicable, although in many cases pilots have adjusted their flying styles to adapt to the circumstances

imposed by control systems, such as those in trim.

35
vehicle for and participating in first

flights, and we may therefore not be the

most qualified to define the best approach.

The current problem, however, is that

very few new aircraft have been designed

and flown in the past fifteen years, and

therefore even industry has had little

recent experience. Many aerospace

companies were essentially without a real

flight-test organization during the lean

years, and some did not have company

test pilots. Thus, with the renewal of

aerospace activity, many companies had

to put a flight-test team together from

scratch. That can spell trouble.

We at the NASA FRC have been fortunate

in being in a position to actually conduct a

number of first flights over the past ten

years. We made first flights on all of the

lifting bodies (the M2-F1, M2-F2, M2-F3,

HL-10, the X-24A, and the X-24B) as

well as the F-8 Supercritical Wing, the

F-8 Digital Fly-By-Wire, the F-111

Transonic Aircraft Technology, and most

recently the F-15 Remotely Piloted

Research Vehicle. We have gained some

appreciation for all of the concerns that go

along with making a first flight and have

developed our own procedures and

ground rules for qualifying a vehicle for

flight. These are by no means all-inclusive, since in many disciplines we depend

on the designer and builder for the

necessary confidence to proceed.

For example, we depend heavily on the

contractor to provide an adequate structure and functional systems. We have in

many cases done the conceptual design

but have never attempted to do the detail

design since we are not designers—a fact

that others in government don’t always

admit. We do monitor and analyze the

design, do proof loading of critical

portions of the structure, and do functional testing of all the subsystems once a

vehicle is delivered. We also do all the

other normal pre-flight things such as taxi

tests and all-up rehearsals of the first

flight with all systems operating and with

all personnel at their appropriate stations.

One of the most important things we do is

simulation. We have learned from experience that extensive simulation is the key

to success in flying a new configuration

or vehicle. We analyze all the wind-tunnel

data and then start with the best fairing of

all the data. Once we complete that

evaluation, we then begin looking at the

worst cases. We intentionally vary each

and every derivative over a wide range to

determine the sensitivity of the vehicle’s

flying qualities to that particular derivative. The range of variability we investigate is much broader than the scatter of

the wind-tunnel data. From experience,

we expect—or I should say we are not

surprised by—discrepancies of 25 to 30

percent in predicted derivatives. In

practice, we vary them as much as 100 to

200 percent. In the case of dynamic or

rotary derivatives, which are hard to

measure both in the wind tunnel and in

flight, we may vary them even more.

Based on experience, we even vary

combinations of various derivatives to

look for the worst possible cases. Any

potential handling-quality problems

exposed in this type of investigation are

thoroughly explored to determine

possible fixes, recovery techniques,

and/or, if necessary, ways to avoid the

problem area. The low-angle-of-attack

PIO problem identified during early

simulations of the M2-F2 was a classic

example. We spent many hours evaluating the problem and determined that we

had two effective recovery techniques.

One was to reduce the rudder-aileron

interconnect ratio and the other was

simply to pull up and increase angle of

attack. We could not easily avoid the

area since we had to go to low angles of

attack to pick up the necessary airspeed

for landing flare. We did, however, have

landing rockets as a backup in case we

did have to increase angle of attack

from the desired pre-flare condition. In

flight, both of these recovery techniques were validated by necessity. The

value of this type of simulator investigation cannot be over-emphasized. We

also feel that it is essential that the pilot

36
participate extensively in the engineering simulation. We at FRC rely heavily

on the pilot for a successful flight

program. The importance of the pilot is

critical, particularly in simulation, the

design or evaluation of the control

system, and first-flight preparation.

Pilots generally have a much broader and

more objective view of the overall picture

[than do other participants in flight

research]. We are fortunate to have

exceptionally well qualified and experienced pilots, all of whom have made first

flights. They are all well suited to serve

as either a project pilot on a new vehicle

or a member of the Flight Readiness

Review Board, and in this manner, we get

double the input. A pilot with first-flight

experience is invaluable. Unfortunately,

there aren’t too many active pilots who

have first-flight experience because of the

limited number of new aircraft that have

been produced in the last fifteen years.

Following the extensive engineering

simulation, we proceed into the procedural and flight-planning simulation

phases. In these phases, we develop the

first-flight plan and then look at every

imaginable failure, malfunction, or

emergency. In developing the control

system and preparing for the first flight

of the M2-F2, we spent at least a year

on the simulator, and I as pilot averaged

two to three hours a day in the simulator cockpit.

We have not normally used anything other

than a fixed-base simulator; however, we

were sufficiently concerned about the low

angle-of-attack PIO problem in the M2-

F2 to also investigate it in-flight with a

variable-stability aircraft. We have

resorted to variable-stability aircraft and

more sophisticated simulators on a few

other occasions; however, the simple

fixed-base simulator has generally been

adequate. We thoroughly exercise the

flight-control system to establish limitcycle boundaries and ensure that we are

free of structural resonance problems as

discussed earlier.

Another very important part of our preflight preparation is a Flight Readiness

Review (FRR). An FRR team is designated at least six months prior to a

scheduled first flight. The team is

generally composed of members of each

of the disciplines involved (aerodynamics, stability and control, performance,

etc.) as well as subsystem experts,

instrumentation experts, and a pilot. The

chairman of the FRR team is at least a

senior division-director-level individual

with a broad test background. The FRR

team members are not associated in any

way with the project team and act as

devil’s advocates. The FRR team has

unlimited access to any data, can monitor any tests, question any project team

member, and make any recommendations on pre-flight preparation. In

essence, it has carte blanche to examine

the program. The FRR reports directly

to the Center Director, and prior to

flight, it submits an oral and written

report. The FRR is an extremely effective means of ensuring a safe first flight.

First Flight and Envelope Expansion

Our general philosophy on first flights is

that once the aircraft is airborne, we

immediately begin worrying about how to

get it back on the ground again safely.

Data maneuvers are of secondary importance. The main area of interest during the

first flight involves the approach and

landing tasks. Various potential failure or

backup control modes are evaluated in the

approach-and-landing configuration, as

are other possible subsystem malfunctions. This emphasis on approach and

landing during the first flight is easily

justified, since on each and every succeeding flight an approach and landing

must be made. Subsystem malfunctions

will ultimately occur during the test

program, and it is thus wise to assess

these potential malfunctions in a controlled manner as early as possible.

Our philosophy on envelope expansion

is not unique. We select the most benign

37
Mach number expansion corridor and

concentrate first on validating stability,

control, and handling-quality characteristics. We update the simulator following each flight [to incorporate what we

learned on that flight that we did not

know before]. Once the Mach number

envelope has been explored, the remainder of the flight envelope is

somewhat systematically expanded in

terms of angle of attack, dynamic

pressure, etc. In the case of the X-15,

we began expanding the altitude and

angle-of-attack envelope about halfway

through the Mach envelope expansion.

Once the design Mach and altitude had

been achieved, we continued to expand

the altitude envelope and simultaneously began expanding the dynamic

pressure and aerodynamic heating

boundaries. As previously indicated,

this type of envelope expansion is

typical of the general approach used

throughout the aircraft industry and is a

time-proven way to test aircraft.

In such an incremental buildup of flight

research, we can, so to speak, poke our

noses into a new area and if we encounter a problem, we can immediately back

out of that area and into a known safeflight region where we have flown

before. When we do encounter a

problem in flight, we come back down,

analyze the data, update the simulator,

and then try to determine a fix for the

problem before we again probe into the

problem area. The longitudinal-sensitivity problem we observed on the first

flight of the HL-10 is a good example.

Flight data confirmed that we had more

control effectiveness than we anticipated. We made a change in the control-system gearing before the next

flight and eliminated the problem.

During the buildup test program in the

X-15, we were fortunate to encounter

our environmental problems one by

one. We burned the canopy seal at

Mach 3.3, well below the design Mach

number of 6.0. We encountered the

nose-gear-door problem at Mach 5. We

saw the first indications of the wingleading-edge and windshield problems

at 5.2 Mach number. The first nose-gear

scoop door opened at 5.5 Mach number.

If we had gone to Mach 6 on the first

flight, we would probably have had all

of these things happen within seconds

of one another. Also, each problem

would have been more severe than we

actually experienced because we would

have had more exposure time.

Remotely Piloted Research Vehicles

(RPRVs)

The Flight Research Center has developed a remotely piloted research

technique that was first applied in the

testing of an advanced lifting entry

configuration, the Hyper III. The

technique illustrated in Figure 12

includes basically a ground cockpit, an

uplink for command control signals,

and a telemetered downlink that closes

the control loop through the pilot’s

instruments and controls. The cockpit

has all the conventional instruments and

controls normally found in an aircraft,

and the pilot thus has complete instrument-flight-rules flight capability. In

addition, a forward-looking television

mounted in the flight model provides

the ground pilot an out-the-window

view for additional reference. A highspeed computer is included in the

control loop to provide or exactly

duplicate any flight control system

augmentation or automatic control

mode. This allows for a relatively

simple and inexpensive on-board

control system.

This technique has recently been

applied to spin testing. A 3/8th-scale

model of the F-15 has been tested

throughout the achievable angle-ofattack range and intentionally spun

using several different control modes.

As of this time, the model had not been

departed or spun using the primary

control mode with the operational stall38
inhibitor system in the loop. We are still

trying, however.

The flight program has, in our opinion,

been an outstanding success even

though we have damaged and finally

lost the first flight vehicle. We anticipated losses and originally had three

vehicles constructed to ensure completing the planned flight program. In the

fourteen flights that have been accomplished,37 we have thoroughly documented the stability and control characteristics of the model from plus 40

degrees to minus 20 degrees angle of

attack. We have spun the vehicle

upright and inverted and departed the

vehicle at several different g-levels. We

have validated two different spin modes

predicted in the Langley spin studies

and confirmed the proposed recovery

techniques. We have not as yet compared model data with full-scale airplane data, since the full-scale flight

data has so far been unobtainable. The

3/8-scale-model data has so far compared quite favorably with wind-tunnel

predictions. The simulator developed

during the flight program is probably the

first good spin simulator ever implemented. Flight results have confirmed the

validity of the simulator, and the simulator can and has been used to investigate

and develop new spin entry techniques.38

37 Readers who skipped the background section may like to know that the vehicle completed 27 flights by 1975, 53 by

mid-July 1981.

Figure 12: The

remotely augmented vehicle

concept. 

H

 H

This was taken from Dwain A. Deets and John W. Edwards, “A Remotely Augmented Vehicle Approach to Flight Testing RPV

Control Systems,” paper presented at the AIAA RPV Technology Symposium, Tucson, AZ, 12-14 Nov. 1974 (also published at

Edwards, CA, as NASA TM X-56029, Nov. 1974), p. 17.

38 Kenneth W. Iliff, “Stall/Spin Flight Results for the Remotely Piloted Spin Research Vehicle,” paper presented at the

AIAA Atmospheric Flight Mechanics Conference, Danvers, MA, 11-13 Aug. 1980 (AIAA Paper 80-1563) gave the

results of flight research with the 3/8-scale F-15, later redesignated the Spin Research Vehicle, most of the way through

its flights. Among other findings it reported were: “the basic airplane configuration was found to be departure and spin

resistant. When control authority was increased, the model could be spun using several techniques developed with the

simulator.” Also: “The acquisition of high quality steady spin data for this vehicle was made possible by the remotely

piloted technique.”

39
Much more remains to be done in the F15 spin program, but whether it will be

completed is questionable because of

other higher-priority flight-program

commitments. Higher-Mach-number

departures and spins should be investigated. The model should be modified to

provide more representative inertias,

since the present inertias are higher than

that required for dynamic scaling. The

inertias are properly ratioed, however.

Additional stability-and-control derivatives should be obtained during the actual

spin. Wind-tunnel data is lacking at the

extreme angle of attack investigated in

flight. And finally, wind-tunnel, small

scale-model, 3/8th-scale model, and fullscale data should be compared to completely validate the RPRV technique. All

of this, we feel, is essential to ensure that

in the future accurate simulations can

completely predict departure and spin

techniques as well as recovery techniques.

The current FRC position on RPRVs is

that they can be used effectively to

provide meaningful and accurate data.

They can be cost effective, and they can

save time and potentially even reduce the

amount of full-scale testing required. The

RPRV technique is extremely attractive

for high-risk-type testing such as spin

testing, testing of new structural concepts,

testing of flutter-suppression systems, etc.

A number of such programs have already

been proposed, and we anticipate many

more to materialize. Our problem now is

to maintain some reasonable balance

between unpiloted and piloted flight

programs.39

There is a definite role for RPRV-type

testing based on what we now know. The

RPRV approach is, however, by no means

a panacea for flight testing. RPRV tests

may still have to be supplemented by

piloted testing, and thus it may not always

be most cost effective overall to go the

RPRV route. We still have a lot to do in

developing the technique and reducing the

potential loss rate. It will also be some

time before the reliability of the on-board

pilot can be reproduced in RPRVs.

Flight-Test Errors

The remarkable safety record mentioned

in the introduction does not imply that we

have been without fault. The M2-F2

landing accident is a good example of

poor judgment on our part. The M2-F2 at

best had marginal lateral-directional

handling characteristics. The pilot initiated a serious PIO inadvertently on the

first flight of the vehicle. Another PIO

occurred on a later flight during a datagathering maneuver involving another

experienced research pilot. Following

the second occurrence, we should have

quit flying the aircraft and gone back to

the wind tunnel to look for a fix. We

chose instead to continue flying the

vehicle without modification. A third

pilot, who had previously flown the

vehicle, encountered a PIO on final

approach on the [six]teenth flight. He

successfully recovered, but as a result

of the PIO, he was forced to attempt a

landing on an unmarked portion of the

lakebed.

Depth perception on the lakebed is

extremely poor without known reference

marks. To further complicate the problem,

a rescue helicopter was operating in the

immediate area of the modified landing

site. This distracted the pilot because of a

39 Following the flight research with the 3/8th-scale F-15/Spin Research Vehicle, the Center flew research programs

with the Mini-Sniffer, the Oblique-Wing Research Aircraft, Drones for Aerodynamic and Structural Testing, and the

Highly Maneuverable Aircraft Technology , among other remotely piloted vehicles. This range of vehicles showed that

although Milt’s comments in 1973 or 1974 were accurate as far as they went, sometimes—as he suggests below—subscale vehicles could be more expensive and time-consuming than full-scale programs because of the need to develop

miniature systems to accommodate the smaller spaces in the vehicles. In other cases, however, RPRV operations could

be cost effective, especially if flights were planned for high data output.

40
concern about a possible collision with

the helicopter. The chase aircraft were

also forced out of position during the

PIO and were not in position to call out

height above the lakebed. The result was

a gear-up landing. The pilot suffered the

permanent loss of vision in one eye. The

vehicle was rebuilt with a center vertical

fin after additional wind-tunnel tests

predicted a significant improvement in

flying characteristics. The flight research

with the modified vehicle proved to be

completely uneventful, since the flying

qualities were quite good, as predicted.

We have had a number of problems in

the past during towing operations

involving unconventional flight vehicles.

Some of these were due to our overall

inexperience with aerial towing. We

reinvented many of the problems well

known to glider and sailplane people

even though we had an experienced

sailplane pilot as Center Director. We

also used poor judgement when we

decided because of cost to do our own

towing. None of our pilots had any real

towplane experience and only a couple

had ever been exposed in any way to

towing operations. As one might expect,

we had several problems and one serious

accident, luckily without any pilot injury.

The events leading up to the accident

started with the acquisition of an L-19,

which we modified for towing the

paraglider. The pilot selected to do the

towing had never flown an L-19, and I

was elected to check him out. Because of

various schedules, only one day was

available for checkout. I rode through

two flights with the newly selected tow

pilot and undoubtedly overrode the

controls, particularly the rudder, during

landing approaches. I was then called

away to a meeting and decided on the

spot to let him take it alone. On his first

landing, he ground looped and severely

bent one main-gear strut.40

The strut was replaced and a towing

flight was scheduled the following day.

On the morning of the scheduled flight,

the pilot who had been checked out

was rescheduled for a higher priority

flight. Another pilot was selected to do

the towing, and I gave him a quick

checkout in the L-19 prior to the actual

towing flight. After takeoff and upon

reaching the edge of the lakebed, the

tow-plane pilot, as instructed, began a

turn to stay close to the lakeshore in

case of a tow-line failure. His rate of

turn was excessive, and within seconds, the tow-line was hanging slack

between us. Since we were only 300

feet or so high, the only recourse was

to release and attempt a landing in the

sagebrush. The vehicle was extensively

damaged in the landing attempt.

Following this accident, we reverted

back to using professional tow pilots.

We finally gave up towing altogether

after two hair-raising incidents while

towing the M2-F1.

The loss of the Number Three X-15

could be attributed to some extent to a

faulty experiment that we developed

and flew on the aircraft. The total

experiment did not undergo a complete

environmental check, although a

component of the experiment, the drive

motor that caused a problem, had

successfully passed all environmental

checks and had been used in other

piggy-back experiments carried on the

aircraft. The motor began arcing at

approximately 80,000 feet altitude on

the way up to a planned maximum

altitude of 250,000 feet. The experiment was supposedly isolated from all

primary aircraft systems, and yet it

caused faulty guidance-computer and

flight-control-system operation. This is

another potential problem to be assessed with command augmentation

and fly-by-wire control systems. The

faulty experiment cannot be completely

40 Milt added here, “We found out later that that particular pilot had never flown a tailwheel airplane.” The pilot in

question wrote beside these words, “Not true. I was the pilot involved. I flew the T-6 210 hours in pilot training.”

41
excused of partial blame for the ultimate loss of the aircraft.41

Another example of questionable judgment

on our part involved the maximum speed

flight of the Number Two X-15 to a Mach

number of [almost] 7. The flight was made

to demonstrate the capability of the X-15 to

carry a supersonic combustion ramjet

(scramjet) engine to Mach numbers

approaching 8. Figure 13 shows the X15A-2 with a dummy ramjet on the lower

stub ventral [, eyelid, drop tanks, and

ablative coating42 for what turned out to be

the Mach 6.7 flight]. In building up to fly

this combined configuration, we first

made a flight with empty tanks to

demonstrate tank jettison capability.

41 Here Milt inserted a comment, “Show time history of X-15 #3.” In lieu thereof, perhaps the section of the accident

report for the aircraft quoted in his At the Edge of Space, p. 263, will better indicate the problems that caused the pilot,

Michael Adams, to lose his life in the accident:

The only unusual problem during the ascent portion of the flight was an electrical disturbance that started at an

altitude of 90,000 feet and that effected [sic] the telemetered signal, the altitude and velocity computer associated with the inertial platform and the reaction controls that operate automatically in conjunction with the MH96 adaptive control system. Although the pilot always had adequate displays and backup controls, the condition

created a distraction and degraded the normal controls. As the aircraft approached the peak altitude of 266,000

feet, it began a slow turn to the right at a rate of about 0.5 degrees per second. The rate was checked by the MH96 system which operated normally for a brief period so that at peak altitude, the aircraft was 15 degrees off

heading. Then the pilot, apparently mistaking a roll indicator for a sideslip (heading) indicator[,] drove the

airplane further off in heading by using the manual reaction controls. Thus the aircraft was turned 90 degrees to

the flight path as the aerodynamic forces became significant with decreasing altitude. The aircraft continued to

veer and entered what appeared to be a classical spin at an altitude of about 230,000 feet and a Mach number of

about 5.0.

Some combination of pilot action, the stability augmentation system, and the inherent aircraft stability caused

the aircraft to recover from the spin at an altitude of about 120,000 feet and a Mach number of about 4.7. As the

aircraft recovered from the spin, however, a control system oscillation developed and quickly became selfsustaining. At this time the airplane was descending at a rate of about 160,000 feet per minute and dynamic

pressure was increasing at nearly 100 pounds per square foot each second. There was a corresponding rapid

increase in the g forces associated with the oscillation, and structural limits were exceeded. The airplane broke

into many pieces while still at high altitude probably in excess of 60,000 feet, and fell to earth northeast of

Johannesburg, California.

The pilot, probably incapacitated by the high g forces[,] did not escape from the cockpit and was killed on

ground impact. The accident board concluded that the accident was precipitated when the pilot allowed the

aircraft to deviate in heading and subsequently drove it to such an extreme deviation that there was a complete

loss of control. The board believes that these pilot actions were the result of some combination of display

misinterpretation, distraction, and possible vertigo. The board further concludes that the destruction of the

aircraft was the result of a sustained control system oscillation driven by the MH-96 adaptive control system

that caused the divergent aircraft oscillations and aerodynamic loads in excess of the structural limits. The

electronic disturbance was attributed to the use in one of the scientific experiments of a motor that was unsuited

to very high altitude environments.

Milt said he did not believe that there was any pilot error and that he thought the accident board agreed with him. He did

think that the vertigo contributed to the accident (pp. 263-264).

42 When the X-15A-2 was rebuilt from the Number Two X-15 following its landing accident, it gained an elongated

fuselage and a small internal tank within the plane. Because the ablative coating put on the aircraft to protect it from

severe heating on the higher-speed flights would char and let off residue at very high velocities, North American had

placed an eyelid above the left cockpit window. The pilot would keep it closed until just before the approach and

landing, using the right window for visibility during launch and most of the remainder of the flight. Above Mach 6,

however, the residue coming from the charred ablator would cover the exposed window and restrict visibility. Hence the

need to open the eyelid for approach and landing.

42
Next we made a flight with full tanks to

demonstrate proper fuel and lox [liquid

oxygen] transfer. The flight was aborted

shortly after launch when there was no

indication of flow from one of the tanks.

The next full-tank flight was successful

to a Mach number of 6.3. The next flight

was with the eyelid and dummy ramjet

to Mach 4.75.

At this point in the program there was a

strong desire to put it all together and

go for an all-out flight. The argument

was that we had adequately demonstrated each of the configuration

changes. We compromised for another

flight with the ablator, eyelid and

dummy ramjet. This flight raised the

speed to Mach 5. We examined the

airplane after that flight and saw some

indications of localized charring but

nothing that we considered significant.

We simply made local repairs to the

ablator and put it all together for a

flight to Mach 7.0. Figure 14 shows

[some of] the results.

A shock wave off the dummy ramjet

interacting with the boundary layer

caused severe localized heating that

burned off all the ablator and burned

through the basic Inconel ventral fin

Figure 13: Photo

showing X-15A-2

with ablative

coating, drop tanks,

and dummy ramjet.

(NASA photo

ECN 1889)

Figure 14: Result of severe heating from the Mach 6.7

flight in the X-15A-2. (NASA photo E-17525)

43
structure. We almost lost the airplane.

We took too big a step. We essentially

went from Mach 5 to almost Mach 7 in

one step with the dummy ramjet. The

moral is that even though we supposedly checked out each individual

configuration, we should have put them

all together and again worked up

incrementally in Mach number. We

would thus have appreciated the

significance of the shock-impingement

heating problem. As mentioned in the

introduction, we have been in the flightresearch business over twenty-seven

years. Many of the people who worked

on the X-1s are still here, and yet we still

occasionally get caught short. We seldom

get caught on the same problem, but it

seems that we never run out of new

problems.

Conclusions

Irrespective of the fact that our new

generation of aircraft (F-14, F-15, F-16,

F-17 [precursor of the F/A-18], and B1) is not probing new frontiers, we are

still seeing discrepancies between

wind-tunnel and flight data as mentioned earlier. The X-1 achieved a

Mach number of 2.5 over twenty years

ago and we are still operating within

that Mach region with most of our new

aircraft. The aerodynamic discrepancies or problems we are currently

seeing are not as dramatic as the loss of

directional control and consequent

tumbling that Chuck Yeager encountered in the X-1.43

We don’t expect surprises such as roll

coupling or aileron reversal, but we are

impressed for example with the unexpected

43 This is an apparent reference to Yeager’s flight in the X-1A on 12 December 1953. Bell engineers had warned him

before the flight that the aircraft might go out of control at speeds above Mach 2.3, but Yeager flew the X-1A to Mach

2.44 (1,612 miles per hour) despite the warnings, which proved correct. He shut off the rocket engine, but the aircraft

became violently unstable, going into something like an oscillatory spin with frequent roll reversals. He was thrown

about the cockpit as the X-1A lost altitude, falling some 50,000 feet (from an altitude of about 76,000 feet). Semiconscious, Yeager brought the decelerating aircraft into a normal spin, recovering to level flight at about 25,000 feet.

Subsequently, he landed on Rogers Dry Lakebed. A pilot with lesser skills and instincts would probably have perished.

See Hallion, On the Frontier, pp. 292, 308, and Hallion, Supersonic Flight: The Story of the Bell X-1 and Douglas D558 (New York: MacMillan, 1972), p.174.

high-angle-of-attack capability of the F-14.

This was not anticipated or at least not

advertised. We still are occasionally disappointed in actual airplane performance. We

still see substantial discrepancies between

predicted and actual basic stability derivatives on occasion, which means we haven’t

improved our predictive capabilities or

techniques substantially in the last twenty

years regardless of any new or improved

ground facilities.

The many discrepancies between windtunnel or predicted and flight data discussed in this report undoubtedly give the

impression that we are extremely critical

and/or in opposition to wind-tunnel

testing. That is absolutely not the case.

Wind tunnels have provided extremely

good data for many years and are continuing to do so. As mentioned earlier, the

wind-tunnel predictions of the X-15

aerodynamics were extremely good.

There have been numerous aircraft

developed that have had no aerodynamic

discrepancies whatsoever. We at FRC

resort to the wind-tunnel people continuously in support of our flight research,

and they have bailed us out, so to speak,

on many occasions. We always request

wind-tunnel support whenever we make

other than a minor configuration change.

We are completely dependent on windtunnel predictions in many things that we

do, such as air launch. We depend entirely

on these predictions to assure that we

have no collisions during separation.

We know, too, that the wind-tunnel people

are not blind to their own limitations. We

have supported, at their request, a number

of combined wind-tunnel/flight-research

correlation tests designed to improve their

44
predictive capability in areas where they

know they have problems simulating the

total flight environment.

One might ask, if we are not critical of

the wind-tunnel results or people, what

are we critical of? The answer is that we

are critical of the system. The system is

not closing the loop. When discrepancies

are noted between wind tunnel and flight,

they are seldom examined in sufficient

detail to pin down the actual reason for

the discrepancies. The tendency is to

dismiss the discrepancy with an excuse

that the tares were wrong, or that the

model was not representative, or that

propulsion effects were not duplicated,

and so on. There is little enthusiasm to go

back and prove it. Both wind-tunnel and

flight people are more enthusiastic about

moving on to the next program [than

about investigating the problem on an

existing or completed program]. As a

result, we see reoccurring problems

continuously.44 Our track record hasn’t

improved that much except in such

catastrophic problem areas as roll coupling and the T-tail deep stall problems.

Significant aerodynamic discrepancies

still show up as do performance- and

control-related deficiencies.

One of NASA’s primary responsibilities is to close the loop from the windtunnel to flight, and yet in many

instances this has never been adequately done. An honest attempt to do

this was made with the XB-70 in trying

to explain the large discrepancy between predicted and actual performance. New models were constructed

and a new series of wind-tunnel tests

was conducted. These tests showed

somewhat better agreement with flight

results and yet a 10 percent discrepancy

still existed; this still would result in a

significant range discrepancy. No

further attempts were made to improve

the correlation. A more serious effort to

correlate wind-tunnel and flight data is

currently underway with the F-111

Transonic Aircraft Technology program.45 Hopefully this will be carried

through to its ultimate conclusion.46

This, of course, does not provide the

correlation we need at higher Mach

numbers. We need additional validation

of wind-tunnel predictions in the

hypersonic speed region since the only

good flight data is from the X-15. We

have successfully flown some smallscale vehicles at hypersonic speeds, but

the flight data obtained was minimal

and compromised by lack of accurate

air data. As far as environmental

problems are concerned, we don’t

anticipate any significant new problems

in the near future since there are no

current plans for higher performance

aircraft.

A potentially serious problem for future

aircraft designers is emerging as the trend

toward contracted wind-tunnel operation

44 At this point in the text, Milt intended to insert a table listing some of the recurring problems, the aircraft involved, and

the period of development of the particular aircraft. Unfortunately, he apparently did not leave behind such a table, and I

am not competent to construct it.

45 Following a great deal of wind-tunnel testing at NASA’s Langley Research Center and by General Dynamics, the

Flight Research Center began flight research with an F-111 on 1 November 1973. The program continued until the late

1970s and was resumed in a second phase in the mid-1980s. See Richard P. Hallion, On the Frontier: Flight Research at

Dryden, 1946-1981 (Washington, DC: NASA SP-4303, 1984), pp. 207-209.

46 In a later document in the files he left behind, Milt wrote, “All objectives of the TACT program have been met in FY

1978. In general, flight data has validated the improvements to aerodynamic characteristics as predicted by wind-tunnel

and calculated data.” Draft for “Annual Report of Research and Technology Accomplishments and Applications, FY

1978, Hugh L. Dryden Flight Research Center,” p. 3, held in the Dryden Historical Reference Collection.

45
increases. We certainly aren’t going to

produce experts like [Richard] Whitcomb,

[Robert T.] Jones, [John] Becker, [Eugene]

Love, [Alfred J.] Eggers, [Clarence]

Syvertson, [probably Robert W.] Rainey[,]

and on and on[,] with contracted windtunnel operations. The aerospace contractor will be pretty much on his own in

assessing the quality of his data. This is

not to say that an individual contractor

cannot do an excellent job on his own.

Individual contractors cannot, however,

have access to all the data that NASA

[does] because of proprietary problems.

They thus would be hampered in developing equivalent experts. NASA has in the

past provided continuity and good advice

to many contractors in the development of

new configurations. NASA has also had

the luxury of looking at far-future and

high-risk concepts. This is a luxury

industry could not afford. Thus, this trend

toward contractor operation of all wind

tunnels should be halted and hopefully

reversed.

To make the future look even more bleak,

there is pressure from some sources to

eliminate all government aeronautical

research and development. In our opinion,

that could mean complete disaster for the

U.S. aerospace industry in the international sale of aircraft and would, as a

result, significantly affect the country’s

balance of payments. We feel the present

cost of government aeronautical research

and development is a very minimal and

essential subsidy to our aerospace industry. We in the flight-research business do

feel that we will still be in business for

some time to come because of the many

other potential new problem areas alluded

to earlier.

The Future

Flight research in the next five to ten years

doesn’t look as though it will be very

exciting. As of now, there are no big

advances planned in terms of flightenvelope expansion for future aircraft.

There are no serious efforts to design a

triple-sonic fighter. Even more disturbing

is that there is no real enthusiasm for a

hypersonic research aircraft. True, we

have flown the X-15 to hypersonic

speeds, but the X-15 did not address

many of the critical disciplines such as

structures, propulsion, etc. Kelly

Johnson was pretty much on his own in

designing the superb YF-12 and SR-71

aircraft. We are currently flying two of

these aircraft in an attempt to determine

why they fly as well as they do. So far

we have seen several discrepancies

between theoretical and actual data.

The boundary layer conditions, for

example, are significantly different

from what one would predict.

Considering the number of aerodynamic,

propulsion, and performance discrepancies we have observed, it is obvious that a

lot of good, sound engineering judgment

was applied in the design of the aircraft.

It is again only after the fact that we are

capable of explaining why. It is discouraging to realize that we have to resort to

operational aircraft to obtain data to

update theoretical and wind-tunnel

predictive capability.

We feel we critically need a new series of

research aircraft to stimulate new aircraft

development. The early series of research

aircraft stimulated a wide variety of new

supersonic operational aircraft in the

1950s. The swept-wing F-100 was based

on the success of the D-558-2. The

straight-wing F-104 was based on the X-1

and X-3 successes. The F-102 was based

on the X-4 and XF-92 results.47 And

finally, the F-111 stemmed from the

marginally successful X-5 results. Subsequent to that extremely stimulating

period, there have been no real[ly]

imaginative developments in aircraft

configurations.

47 Milt had said XF-91. As an anonymous reviewer commented, this should be the XF-92. “The XF-92 was a delta wing.

The XF-91 was a reverse taper prototype.”

46
The current prototype fighter program

is a step in the right direction, as are

the Highly Maneuverable Aircraft

Technology (HiMAT) and Advanced

Fighter Technology Integration (AFTI)

programs. However, none of these are

real challenging. They are all still

focused primarily on the transonic

flight regime, as were the F-14 and F15. To us this is indicative of shortsightedness. We feel that the Vietnamese conflict convinced too many

advanced planners that all future

combat would take place at transonic

speeds. They have not acknowledged

the fact that this was inevitable at the

time because of the thrust-to-weight

ratios of the aircraft involved. The new

prototype fighters are capable of

supersonic combat because of their

high thrust-to-weight ratios and can

virtually eat up the highly touted and

transonically optimized F-14 and F-15,

with properly developed tactics.

There are those doomsayers who say that

we will never have supersonic fighter

combat. That is ridiculous. Give a fighter

pilot an aircraft capable of supersonic

combat and he’ll find a way to use that

advantage, just as he did with the early jet

aircraft against the best propeller-driven

aircraft. You quickly learn not to fight in the

opponent’s best arena. If one believes that

philosophy, then the Spad is still the best

fighter ever conceived.48 In the case of the

F-14 and F-15, the philosophy is that if the

Foxbat49 comes down to my piece of the

sky, I’ll eat him up. If he doesn’t, I’ll shoot

him down from below. If that philosophy

holds true, the Navy should resurrect the

old B-52s, hang a hundred or so Phoenix

missiles on them and have a fleet of flying

battleships. Better yet, rebuild some

dirigibles and have worldwide air superiority.

48 There were several models of Spads, but undoubtedly Milt is referring to the Spad XIII built in France at the end of

World War I and flown by the French, Italian, and Belgian forces as well as the American Expeditionary Force. Many

famous pilots flew it, including Captain “Eddie” Rickenbacker. The fighter served well into the 1920s in seven countries.

49 The Foxbat was the NATO reporting name of the Soviet MiG-25, which could climb to over 123,000 feet.

If the U.S. is to retain air superiority, we

must begin developing aircraft that can

go up and stick their noses up the

Foxbat’s tailpipe. We have the technology

in hand to build a true triple-sonic fighter

that can maneuver aggressively in the

Foxbat’s arena. We can give our pilots the

perch for the first time since World War I.

We lost the perch in World War II and

haven’t regained it since. It’s time we did.

47
Index

Adams, Michael, 19 n., 42 n.

Advanced Fighter Technology Integration (AFTI)

 program, 47

Aerodynamic heating, 29-30, 43-44

Aiken, William P., 5 n.

Air Force, 3, 7, 9-10, 14, 21

Air Force Flight Test Center, 7, 8, 14

Altitude and equipment problems, 30, 41-42

Ames Aeronautical Laboratory, 14

Ames Research Center, 14, 21

AP-101 flight control computer, 11

Apollo, 10

Computer, 11

Apt, Mel, 8

Army Air Forces, 2-3

B-1, 44

B-2, 11

B-29, 2, 3 ill.

B-50, 2

B-52, 47

B-58, 23

Becker, John, 46

Bell Aircraft, 2

Bell Aerosystems, 14

Bell, Larry, 5

Boeing 747, 12

Boeing 777, 12

Boundary layer, 25 n.

Bronson, Jack, 21

Central airborne performance analyzer, 11

Century series of fighters, 6, 7 ill

Compressibility, 1, 20-21

Computational fluid dynamics, 18, 24 n.

Control, see flight control

Coupling dynamics, 6, 8, 44

Cross, Carl, 19 n.

Crossfield, Scott, 8

D-558-1 Skystreak, 2, 3 ill.-4, 5 ill., 6, 19 n.

D-558-2 Skyrocket, 5 ill., 6, 8, 22 n., 46

Delta-wing aircraft, 23, see also XF-92A, F-102

Design, 36, 46

Digital fly-by-wire, 11-12

Douglas Aircraft, 2

Drag, 25, and see lift-to-drag ratio

Draper Laboratory, 11

Dryden Flight Research Center, Hugh L., 1, 7, passim

Dryden, Hugh L., 18 ill., 32

Edwards Air Force Base, 5 ill.

South Base, 5 ill., 13

Eggers, Alfred J., 14, 21, 46

Envelope expansion, 37-38

F-8 Digital Fly-By-Wire (DFBW) project, 11-13, 11 ill.

F-8 Supercritical Wing project, 15-16, 16 ill.

F-14, 44, 47

F-15, 7, 44, 47

F-15, three-eighths-scale, 16-17, 36, 38-40

F-16, 6, 11, 24, 35, 44

F-18, 7, 11, 44

F-22, 7, 11

F-100, 7 ill., 46

F-101, 7 ill.

F-102, 7 ill., 46

F-104, 7 ill., 15 ill., 19 n., 46

F-111, 16, 34-35, 36, 45, 46

F-117, 11

First flight preparation, 35-37

Flight control, 19-22, 27-28, 30-32, 32-35, 38-39 and see

 F-8 DFBW

Flight Readiness Review, 37

Flight research, discussed, 7-8, 16, 19, 36-38, passim

Flight Research Center (FRC), 1, 16, 19, passim

Flight test, discussed, 7-8, 19, 32, 38

Flow separation, 19-20

Gemini, 10

Ground effects, 25

Ground-vibration testing, 33

Haise, Fred, 12

Hallion, Richard P., 6

Handling qualities, 21-22, 36

Highly Maneuverable Aircraft Technology (HiMAT)

 vehicle, 47

High-Speed Flight Station, 1, 19

HL-10, 14 ill., 19-21, 33, 38

Hyatt, Abraham, 2

Hyper III, 38

Hypersonic speeds, 8, 25 n., 27 n., 30, 38, 42-44, 45

Inertial coupling, see coupling dynamics

Iron Bird, 31-32

Johnson, Kelly, 46

Jones, Robert T., 46

Kotcher, Ezra, 2

Krier, Gary, 13

L-19, 41

Laminar flow, 25 n.

Langley Memorial Aeronautical Laboratory, 2

Langley Research Center, 16

Lifting bodies, 14-15, 21, 33, 36, and see M2-F1, M2-F2,

 M2-F3, HL-10, X-24

Lift-to-drag ratio, 22-23, 25

Lilly, Howard, 19 n.

Loftin, Larry, 16

Love, Eugene, 46

Lunar Landing Research Vehicle (LLRV), 13 ill., 14

48
Lunar Landing Training Vehicles, 14

M2-F1, 14, 34, 41

M2-F2, 14, 19 n., 21, 33, 36, 37, 40-41

M2-F3, 14, 15 ill., 19 n., 41

Manke, John, 13

Martin Company, 14, 21

Me 262A, 1

Mercury, 9

Meteor F.Mk., 1

MiG-25 Foxbat, 47

Minalga, George P., 5 n.

Minneapolis-Honeywell adaptive control system, 25, 30-

 32

Movable horizontal stabilizer, 3, 6

Muroc Army Air Field, 4-5

National Advisory Committee for Aeronautics (NACA),

 1-2, 5

Research Airplane Projects Panel, 8-9

National Aeronautics and Space Administration, 6

Navy Bureau of Aeronautics, 2

North American Aviation, Inc., 9

Northrop, 14

P-38 Lightning, 1, 2 ill.

P-51 Mustang, 1 ill.

Paresev, 41

Pilot-induced oscillation (PIO), 11-13, 22, 27-28, 33-34,

 36, 40-41

Pilots, reliance upon, 37

Pilot saves, 32

PIO filter, 13

Power effects, 21-22

Propellers, 1

Rainey, Robert W. [?], 46

Reaction controls, 25

Reaction Motors, 2

Remotely Piloted Research Vehicles, 38-40, and see F-15,

 three-eighths-scale

Rocket-plume effects, see power effects

Rogers Dry Lake, 4 ill., 5

Root, Eugene L., 2

Rosamond Dry Lake, 28

Rudder-aileron interconnect, 33, 36

Saltzman, Ed, 6 n., 8 n.

Shuttle Carrier Aircraft, see Boeing 747

Simulation, 6, 31-32, 34, 36-37, 39-40

Sound barrier, 2, 5

Space Shuttle, 11, 13, 14-15

Approach and Landing Tests, 12-13

Speed, 1-2, 5, 8, 11

Spin Research Vehicle, 16-17

SR-71 Blackbird, 11, 46

Stability and control, 21, 24, and see flight control

Stability augmentation systems, 28, 33

Stack, John, 2-3, 5

Structural resonance, 28

Supersonic flight, 2, 6, 8, 10-11

Syvertson, Clarence, 21, 46

T-6, 41 n.

Three-eighths-scale F-15, see F-15

Towing, 41

Transonic speed range, 2

Transport aircraft, 6

Trim, 21

Walker, Joseph, 19 n.

Whitcomb, Dr. Richard, 16, 23-24, 46

Williams, Walter C., 5 n.

Wind tunnels, 6, 16, 20-21, 23-25, 36, 39-40, 44-46

World War II, 1

Wright, Orville and Wilbur, 1

XB-70A, 9 ill., 10, 19 n., 22-23

XF-92A, 5 ill., 6, 23, 46

XS-1, 2, 3 ill.-4

XLR-11 rocket engine, 2, 27

XLR-99 rocket engine, 25, 27

X-1, 5 ill., 6, 44, 46

X-2, 6 ill.

X-3, 5 ill., 6, 8, 46

X-4, 5 ill., 6, 46

X-5, 5 ill., 6, 46

X-15, 8-10, 9 ill., 19, 24-32, 24 ill., 29 ill., 38, 41-44

Auxiliary-Power-Unit Problems, 30

Fatigue problems, 32

Fuel-jettison problems, 28

Landing gear problems, 28-29

X-24A, 14, 21-22

X-24B, 14, 15 ill.

X-29, 12

X-31, 12

X-33, 15

X-38, 15

Yagy, Paul F., 21

Yaw reversal, 22

Yeager, Charles (Chuck), 4-5, 44

YF-12 Blackbird, 10-11, 10 ill., 46

YF-16, 24, 35

YF-17, 24, 44

Young, James, 8 n.

Youngblood, Harold H., 5 n.

49
About the Author

Milton O. Thompson was born in Crookston, Minn., on 4 May 1926. He began flying with the U.S. Navy as a

pilot trainee at the age of 19 and served with the Navy for six years, including duty in China and Japan at the

end of World War II. At the conclusion of his period of service with the Navy, he entered the University of

Washington and graduated with a bachelor of science degree in engineering in 1953.

After two years of employment with the Boeing Aircraft Co. as a flight test engineer, Milt became an engineer

at the High-Speed Flight Station, predecessor of NASA’s Dryden Flight Research Center. He joined the

Station in 1956 and was assigned to the pilots’ office two years later, serving as a research pilot until 1968.

Associated with many significant aerospace projects in that period, Milt is perhaps best known as one of the

initiators of and pilots in the lifting-body program and as an X-15 pilot. He was the first pilot to fly a lifting

body—the lightweight M2-F1—and later piloted the heavyweight M2-F2. He flew the X-15 14 times, reaching a maximum speed of 3,723 miles per hour and a peak altitude of 214,100 feet.

In 1968 Milt moved on to become the director of Research Projects. He followed that assignment with

appointment in 1975 as Chief Engineer. He remained in that position until 1993, the year of his death. During

the 1970s, he was a member of NASA’s Space Transportation System Technology Steering Committee, in

which capacity he was successful in applying the knowledge from the lifting-body and X-15 programs to the

Shuttle design. That is, he led the effort to design the orbiters for power-off landings rather than increase their

weight with air-breathing engines for airliner-type landings. His work on this committee earned him NASA’s

highest award, the Distinguished Service Medal—but one of numerous awards he received.

In the year before he died, Milt published At the Edge of Space: The X-15 Flight Program (Washington, DC,

and London: Smithsonian Institution Press, 1992). After his death, Curtis Peebles edited and completed his

Flight without Wings: NASA Lifting Bodies and the Birth of the Space Shuttle (Washington, DC: Smithsonian

Institution Press, 1999), of which Curtis is listed as co-author. In addition, Milt wrote or co-authored about 20

technical reports.

Monographs in Aerospace History

Launius, Roger D., and Gillette, Aaron K. Compilers. The Space Shuttle: An Annotated Bibliography. (Monographs in Aerospace History, No. 1, 1992).

Launius, Roger D., and Hunley, J.D. Compilers. An Annotated Bibliography of the Apollo Program. (Monographs in Aerospace History, No. 2, 1994).

Launius, Roger D. Apollo: A Retrospective Analysis. (Monographs in Aerospace History, No. 3, 1994).

Hansen, James R. Enchanted Rendezvous: John C. Houbolt and the Genesis of the Lunar-Orbit Rendezvous

Concept. (Monographs in Aerospace History, No. 4, 1995).

Gorn, Michael H. Hugh L. Dryden’s Career in Aviation and Space. (Monographs in Aerospace History,

No. 5, 1996).

Powers, Sheryll Goecke. Women in Flight Research at the Dryden Flight Research Center, 1946-1995

(Monographs in Aerospace History, No. 6, 1997).

Portree, David S.F. and Trevino, Robert C. Compilers. Walking to Olympus: A Chronology of Extravehicular

Activity (EVA). (Monographs in Aerospace History, No. 7, 1997).

50
Logsdon, John M. Moderator. The Legislative Origins of the National Aeronautics and Space Act of 1958:

Proceedings of an Oral History Workshop (Monographs in Aerospace History, No. 8, 1998).

Rumerman, Judy A. Compiler. U.S. Human Spaceflight: A Record of Achievement, 1961-1998 (Monographs in

Aerospace History, No. 9, 1998).

Portree, David S.F. NASA’s Origins and the Dawn of the Space Age (Monographs in Aerospace History, No. 10,

1998).

Logsdon, John M. Together in Orbit: The Origins of International Cooperation in the Space Station Program

(Monographs in Aerospace History, No. 11, 1998).

Phillips, W. Hewitt. Journey in Aeronautical Research: A Career at NASA Langley Research Center (Monographs in Aerospace History, No. 12, 1998).

Braslow, Albert L. A History of Suction-Type Laminar-Flow Control with Emphasis on Flight Research (Monographs in Aerospace History, No. 13, 1999).

Logsdon, John M. Moderator. Managing the Moon Program: Lessons Learned from Project Apollo (Monographs in Aerospace History, No. 14, 1999).

Perminov, V.G. The Difficult Road to Mars: A Brief History of Mars Exploration in the Soviet Union (Monographs in Aerospace History, No. 15, 1999).

Tucker, Tom. Touchdown: The Development of Propulsion Controlled Aircraft at NASA Dryden (Monographs in

Aerospace History, No. 16, 1999).

Maisel, Martin D.; Demo J. Giulianetti; and Daniel C. Dugan. The History of the XV-15 Tilt Rotor Research

Aircraft: From Concept to Flight. (Monographs in Aerospace History #17, NASA SP-2000-4517, 2000).

Jenkins, Dennis R. Hypersonics Before the Shuttle: A History of the X-15 Research Airplane. (Monographs

 in Aerospace History #18, 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. (Monographs in Aerospace History #19, NASA SP-2000-4519).

Waltman, Gene L. Black Magic and Gremlins: Analog Flight Simulations at NASA’s Flight Research Center.

(Monographs in Aerospace History #20, NASA SP-2000-4520).

Portree, David S.F. Humans to Mars: Fifty Years of Mission Planning, 1950-2000. (Monographs in Aerospace

History #21, NASA SP-2002-4521).

Those monographs still in print are available free of charge from the NASA History Division, Code ZH, NASA

Headquarters, Washington, DC 20546. Please enclosed a self-addressed 9x12" envelope stamped for 15 ounces

for these items.

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