Roger D. Launius and Dennis R. Jenkins

 Reentry and 

Recovery 

from Space
Roger D. Launius and Dennis R. Jenkins

 Reentry and 

 Recovery

from Space
 Library of Congress Cataloging-in-Publication Data

 2012022456

F as ro el t yb eh S 

I tn re en :t tskoob ro e pg. 

aF :x ( 5 )202 

ISB N 978-0-16-091064-7

9 780160 910647

90000

F ro as le b y t eh S pu eri tn e dn e tn fo D co mu e tn s, .U S. G vo er mn e tn P ri tn i gn Office

I tn er en t: koob st ro e. opg . vog P noh e: lot l ree f ( 668 ) 215 - 0081 ; DC a re a ( 202 ) 215 - 0081

Fa :x ( 202 ) 215 - 4012 Mail: St po IDCC, W a hs i gn t no , D C 20402 - 1000

ISBN 978-0-16-091064-7
Acknowledgments v

Introduction vii

Chapter 1: Creating Ballistic Reentry .......................................................... 1

Early Ideas About Spaceplanes 2

Sänger, Bredt, and the Silverbird 7

The First Ballistic Missiles 9

The Problem of Reentry 13

Nuclear Warheads and the Blunt-Body Theory 16

Thermal Protection Systems 26

Better Warheads 28

The CORONA Satellite Reconnaissance Program 31

Chapter 2: Human Space Flight and the Problem of Reentry ..................... 35

The X-Planes 37

Round Two X-Planes 39

Faget and Nonlifting Reentry 44

Man in Space Soonest 48

Project Mercury 50

Project Gemini 62

Project Apollo 69

Chapter 3: Landing Under Canopies ............................................................ 81

Pioneering Recovery: Project Mercury 83

Recovery from Long-Duration Earth-Orbital Flight: Project Gemini 90

Pursuit of the Gemini Paraglider 94

Three Canopies Into the Pacific: Project Apollo 117

Chapter 4: Spaceplane Fantasies ................................................................ 125

Choices 127

Lifting Bodies 131

Air Force–Sponsored Research 137

Piloted Lifting Bodies 142

M2-F2 144

M2-F3 147

HL-10 148

iii
Air Force Efforts 152

X-24A 153

X-24B 155

Orbital Lifting Bodies 159

BoMi and Brass Bell 160

Rocket-Bomber 164

The Stillborn Round Three X-Plane 169

X-20 Dyna-Soar 171

The 1970 Hypersonic Vehicles Report: 

 Setting the Stage for the Future 182

Chapter 5: Spaceplane Reality .................................................................... 185

Phase B Studies 189

Reusable Surface Insulation 195

Phase C/D Developments 198

Orbiter Contract Award 200

Thermal Protection System 207

Reinforced Carbon-Carbon 208

Reusable Surface Insulation Tiles 210

Space Shuttle Entry 217

Chapter 6: Back to the “Ablative” Future .................................................... 221

The Rise (and Fall) of the X-30 National Aero-Space Plane 221

The X-33/VenturestarTM Program 225

X-34 and the Continuing Challenge of Reentry 234

X-38 Recovery Parafoil 236

Bringing Deep Space Probes Back to Earth: Genesis and Stardust 244

The Orion Capsule as the Next Generation Piloted Vehicle 253

A New President Changes the Paradigm 264

NASA’s Commercial Crew and Cargo Program 270

The Air Force and the X-37B 272

Summary 275

Chapter 7: Summary .................................................................................... 279

About the Authors 287

Index 289

iv
As with any published work, the authors of this book inevitably incurred debts 

to patrons, colleagues, and fellow travelers. In writing Coming Home: Reentry 

and Recovery from Space, we certainly relied on the help of many people. Most 

importantly, this project was made possible through a grant from the NASA 

Aeronautics Mission Directorate, where Tony Springer has been instrumental 

in sponsoring a large number of important historical studies. We are indebted 

to the many peer reviewers who offered important suggestions for improving 

the manuscript. We hope that they will be pleased with the final product.

The authors would like to acknowledge several individuals who aided in the 

preparation of this book. Only through the assistance of these key people were 

we able to assemble the materials for this volume in some semblance of order. 

For their many contributions to this project, we especially wish to thank Jane 

Odom and her staff of archivists at the NASA History Program Office for helping to track down information and correct inconsistencies. We would also like 

to thank Steve Dick, Bill Barry, Steve Garber, and Nadine Andreassen at NASA; 

the staffs of the NASA Headquarters Library and the Scientific and Technical 

Information Program who provided assistance in locating materials; Marilyn 

Graskowiak and her staff at the National Air and Space Museum (NASM) 

Archives; and the many archivists and scholars throughout NASA and other 

organizations whose work was used to complete this manuscript. Archivists at 

the National Archives and Records Administration were also instrumental in 

making this study possible.

Patricia Graboske, head of publications at NASM, provided important 

guidance for this project. Thanks to Tom Watters, Bruce Campbell, John 

Grant, and Jim Zimbelman of NASM’s Center for Earth and Planetary Studies 

for information on science missions. We also thank Karen McNamara for 

information about the Genesis and Stardust missions.

In addition to these individuals, we wish to acknowledge the following people 

who aided in a variety of ways: Debbora Battaglia, David Brandt, William E. 

Burrows, Erik Conway, Bob Craddock, Tom D. Crouch, Gen. John R. Dailey, 

David H. DeVorkin, Robert Farquhar, Jens Feeley, James Rodger Fleming, 

James Garvin, Lori B. Garver, G. Michael Green, Barton C. Hacker, James R. 

Hansen, Wes Huntress, Peter Jakab, Violet Jones-Bruce, Sylvia K. Kraemer, 

John Krige, Alan M. Ladwig, W. Henry Lambright, Jennifer Levasseur, John M. 

Logsdon, W. Patrick McCray, Howard E. McCurdy, Jonathan C. McDowell, 

v
Coming Home

Ted Maxwell, Valerie Neal, Allan A. Needell, Michael J. Neufeld, Frederick 

I. Ordway III, Scott Pace, Robert Poole, Alan Stern, Harley Thronson, and 

Margaret Weitekamp. All of these people would disagree with some of this 

book’s observations, but such is the boon and bane of scholarly inquiry.

vi
Saturday morning, February 1, 2003, dawned as most other weekend days in 

America in the early 21st century. People went about their morning routines 

and perhaps turned on their radios or televisions to catch the latest news. 

Those that did so learned of a tragic story that captured the Nation’s attention throughout that day as well as the days that followed. At 8:51:14 eastern 

standard time (EST), during what most people believed to be a routine landing 

of the Space Shuttle Columbia at the Kennedy Space Center, FL, the vehicle 

started breaking apart. By 8:58:19 EST, a large-scale wing deformation had 

taken place, and a fraction of second later Mission Control at the Johnson 

Space Center in Houston, TX, lost all telemetry. Thereafter, the main vehicle 

disintegrated at 9:00:18 EST, based on video imagery. In the process, a crew 

of seven astronauts lost their lives; debris from the Shuttle scattered across 

the American South, from the Pacific coast to the Gulf states; the National 

Aeronautics and Space Administration (NASA) grounded the Shuttle fleet 

pending an investigation; and as the Nation grieved, many in technical communities worked to determine what had taken place and, most importantly, 

why it had happened.

The subsequent Columbia Accident Investigation Board (CAIB), under the 

leadership of Retired Admiral Harold W. Gehman, Jr., identified the technical 

cause of the accident as taking place during launch when the External Tank’s 

foam insulation struck the leading edge of the left wing and damaged the 

thermal protection system. Equally important, the CAIB explored the nontechnical managerial, organizational, and cultural issues that had allowed the 

technical problems that led to the wing breech to go unchecked for the life of 

the Shuttle program. At a fundamental level, those larger issues emerged as 

more significant than the technical factors leading to the Columbia accident.

The tragic loss of Columbia brought to the fore one of the most difficult 

issues NASA has had to deal with since the beginning of space flight: how its 

space systems operate while transiting the atmosphere as they return to Earth. 

The technologies for the reentry and recovery from space might change over 

time, but the challenge remains one of the most important and vexing in the 

rigorous efforts to bring spacecraft and their crews and cargo home successfully. 

Returning to Earth after a flight into space is a fundamental challenge, and 

contributions from the NASA Aeronautics Research Mission Directorate in 

aerodynamics, thermal protection, guidance and control, stability, propulsion, 

vii
Coming Home

and landing systems have proven critical to the success of the human space 

flight and other space programs. Without this base of fundamental and applied 

research, the capability to fly into space would not exist.

Accordingly, this study represents a means of highlighting the myriad of 

technological developments that made possible the safe reentry and return 

from space and the landing on Earth. This story extends back at least to the 

work of Walter Hohmann and Eugen Sänger in Germany in the 1920s and 

involved numerous aerospace engineers at the National Advisory Committee 

for Aeronautics (NACA)/NASA Langley and the Lewis (now the John H. 

Glenn Research Center at Lewis Field) and Ames Research Centers. For example, researchers such as H. Julian Allen and Alfred J. Eggers, Jr., at Ames pioneered blunt-body reentry techniques and ablative thermal protection systems 

in the 1950s, while Francis M. Rogallo at Langley developed creative parasail 

concepts that informed the development of the recovery systems of numerous 

reentry vehicles.

The chapters that follow relate in a chronological manner the way in which 

NASA has approached the challenge of reentering the atmosphere after a space 

mission and the technologies associated with safely dealing with the friction 

of this encounter and the methods used for landing safely on Earth. The first 

chapter explores the conceptual efforts to understand the nature of flight to 

and from space and the major developments in the technologies of reentry and 

landing that took place before the beginning of the space age in 1957.

While most proposals for satellites prior to 1957 avoided the difficult problem 

of reentry, the work done on solving the practical problems of space flight made 

it increasingly obvious that returning to Earth represented a major step in flight. 

Beginning with theoretical work, experimentation followed to help understand 

the heating of reentry during ballistic missile development in the 1950s.

Three reentry approaches dominated thinking at the time. The first was a 

heat-sink concept that sought to move quickly from space through the upper 

atmosphere. The heat sink, usually a large mass of copper or beryllium, would 

simply absorb the heat as the object plummeted through the atmosphere. 

However, superheating proved a serious problem as range and speed grew, and 

engineers soon realized that heat sinks were unacceptable for orbital reentry. 

The second approach, championed by Wernher von Braun and his rocket team 

in Huntsville, AL—and championed by post–World War II German officer 

Walter R. Dornberger at Bell Aircraft—called for circulating a fluid through 

the spacecraft’s skin to soak up the heat of reentry. Von Braun’s grandiose 

vision foresaw astronauts returning from wheeled space stations aboard huge 

spaceplanes, but when challenged to develop actual hardware, he realized that 

there was no way for the heat to be absorbed without killing the occupants. For 

orbital flight, both of these concepts gave way to Allen and Eggers’s blunt-body 

viii
Introduction

concept, which fundamentally shaped the course of space flight research and 

provided the basis for all successful reentry vehicles. At the same time, John V. 

Becker and others at the NACA’s Langley Research Center were championing the X-15, a winged research airplane intended to demonstrate the ability 

to fly back from near-space to normal runway landings. Ironically, the X-15 

ultimately used a variation of Allen’s blunt-body theory.

The X-planes of the 1950s, especially the X-15, proved critical to advancing reentry and recovery technologies for space flight and are the subject of 

the second chapter in this study. They advanced knowledge of the blunt-body 

reentry design with an ablative shield to deal with the heat generated by atmospheric friction. This approach proved significant for piloted missions, but 

it has also been used successfully in reconnaissance, warhead, and scientific 

reentry from the 1950s to the present day. Additionally, the question of what 

materials to use to protect the spacecraft during blunt-body reentry emerged, 

and research on metallic, ceramic, and ablative heat shields prompted the 

decision to employ ablative technology. All of these decisions required tradeoffs, and the story of the process whereby these decisions were made and 

implemented offers an object lesson for current engineers involved in making 

difficult technical choices.

Chapter 2 also investigates the methods of landing once a spacecraft reaches 

subsonic speeds. Once the orbital energy is converted and the heat of reentry 

dissipated, the spacecraft must still be landed gently in the ocean or on land. 

Virtually all of the early concepts for human space flight involve spaceplanes 

that flew on wings to a runway landing; Sänger’s antipodal bomber of the 1940s 

did so as did von Braun’s popular concepts. However, these proved impractical for launch vehicles available during the 1950s, and capsule concepts that 

returned to Earth via parachute proliferated largely because they represented 

the “art of the possible” at the time.

Chapter 3 tells the story of reentry from space and landing on Earth from 

the beginning of the space age through the end of the Apollo program. During 

that period, NASA and other agencies concerned with the subject developed 

capsules with blunt-body ablative heat shields and recovery systems that relied 

on parachutes. The Department of Defense (DOD) tested this reentry concept publicly with Project SCORE (Signal Communication by Orbiting Relay 

Equipment) in 1958 and employed it throughout the CORONA satellite reconnaissance program of the 1960s, snatching in midair return capsules containing unprocessed surveillance footage dangling beneath parachutes. With the 

Mercury program, astronauts rode a blunt-body capsule with an ablative heat 

shield to a water landing, where the Navy rescued them. Project Gemini eventually used a similar approach, but NASA engineers experimented with a Rogallo 

wing and a proposed landing at the Flight Research Center (now Dryden Flight 

ix
Coming Home

Research Center) on skids similar to those employed on the X-15. When the 

Rogallo wing failed to make the rapid progress required, NASA returned to the 

parachute concept used in Mercury and essentially used the same approach in 

Apollo, although with greatly improved ablative heat shields.

At the same time, the DOD pursued a spaceplane concept with the X-20 

Dyna-Soar orbital vehicle that would have replaced the ablative heat shield with 

a reusable metallic heat shield and a lifting reentry that allowed the pilot to fly 

the vehicle to a runway landing. This is also the general approach pursued by the 

DOD with its Aerothermodynamic Elastic Structural Systems Environmental 

Tests (ASSET) and Martin X-23A Precision Reentry Including Maneuvering 

reEntry (PRIME) vehicles. NASA and DOD also experimented with lifting 

body concepts. Engineers were able to make both of those approaches to reentry and landing work, making tradeoffs on various other capabilities in the 

process. The eventual direction of these programs was influenced more by 

technological choices than by obvious decisions.

Even as Apollo was reaching fruition in the late 1960s, NASA made the 

decision to abandon blunt-body capsules with ablative heat shields and recovery systems that relied on parachutes for its human space flight program. 

Instead, as shown in chapters 4 and 5, it chose to build the Space Shuttle, 

a winged reusable vehicle that still had a blunt-body configuration but used 

a new ceramic tile and reinforced carbon-carbon for its thermal protection 

system. Parachutes were also jettisoned in favor of a delta-wing aerodynamic 

concept that allowed runway landings. Despite many challenges and the loss of 

one vehicle and its crew due to a failure with the thermal protection system, this 

approach has worked relatively effectively since first flown in 1981. Although 

NASA engineers debated the necessity of including jet engines on the Shuttle, 

it employed the unpowered landing concept demonstrated by the X-15 and 

lifting body programs at the Flight Research Center during the 1960s. These 

chapters lay out that effort and what it has meant for returning from space 

and landing on Earth.

The concluding chapter explores efforts to develop new reentry and landing concepts in the 1990s and beyond. During this period, a series of ideas 

emerged on reentry and landing concepts, including the return of a metallic 

heat shield for the National Aero-Space Plane and the X-33, the Roton rotary 

rocket, the DC-X powered landing concept, and the Crew Exploration Vehicle 

(CEV) of the Constellation program between 2005 and 2009. In every case, 

these projects proved too technologically difficult and the funding was too 

sparse for success. Even the CEV, a program that returns to a capsule concept 

with a blunt-body ablative heat shield and parachutes (or perhaps a Rogallo 

wing) to return to Earth (or, perhaps, the ocean), proved a challenge for engineers. The recovery of scientific sample return missions to Earth, both with 

x
Introduction

the loss of Genesis and the successful return of Stardust, suggests that these 

issues are not exclusive to the human space flight community. As this work 

is completed, NASA has embarked on the Commercial Crew Development 

(CCDev) program in which four firms are competing for funding to complete 

work on their vehicles:

•	

•	

•	

•	

Blue Origin, Kent, WA—a biconic capsule that could be launched 

on an Atlas rocket.

Sierra Nevada Corporation, Louisville, CO—Dream Chaser lifting body, which could be deployed from the Virgin Galactic 

WhiteKnightTwo carrier aircraft for flight tests.

Space Exploration Technologies (SpaceX), Hawthorne, CA—

Dragon capsule spacecraft; also a partial lifting body concept to be 

launched on the Falcon 9 heavy lifter.

The Boeing Company, Houston, TX—a 7-person spacecraft, including both personnel and cargo configurations designed to be launched 

by several different rockets, and to be reusable up to 10 times.

These new ideas and a broad set of actions stimulated through the CCDev 

program suggest that reentry and recovery from space remains an unsettled issue 

in space flight. This book’s concluding chapter suggests that our understanding 

of the longstanding complexities associated with returning to Earth safely has 

benefited from changes in technology and deeper knowledge of the process; 

however, these issues are still hotly debated and disagreement remains about 

how best to accomplish these challenging tasks. Engineers have had success with 

several different approaches to resolving the challenges of reentry and landing. 

Discovering the optimal, most elegant solutions requires diligence and creativity.

This history seeks to tell this complex story in a compelling, sophisticated, 

and technically sound manner for an audience that understands little about 

the evolution of flight technology. Bits and pieces of this history exist in other 

publications, but often overlooked is the critical role these concepts played in 

making a safe return to Earth possible. Moreover, the challenges, mysteries, and 

outcomes that these programs’ members wrestled with offer object lessons in how 

earlier generations of engineers sought optimal solutions and made tradeoffs. 

With the CCDev program—a multiphase program intended to stimulate the 

development of privately operated crew vehicles to low-Earth orbit currently 

underway—NASA is returning to a capsule concept for space flight. This may 

prove a significant development, and this history could help enlighten the NASA 

team about past efforts and the lessons learned from those efforts.

xi
Robert H. Goddard working on one of his rockets in the 1930s. NASA.

xii
CHAPTER 1

The atmosphere surrounding Earth protects and supports life on the planet, 

but it also makes space flight more difficult. It is said, only half-jokingly, that 

getting to orbit is halfway to anywhere because of the considerable amount of 

energy necessary to go beyond the gravity well of this planet.1

 Generally overlooked, however, is just how difficult it is to return from orbit. All of the energy 

expended to get to orbit dissipates on the way back to Earth, usually in the form 

of extreme heating. In addition to the aerodynamic concerns of high-speed 

flight, there are serious thermodynamic issues with a 17,500-miles-per-hour 

(mph) plunge through Earth’s atmosphere.2

 Unlike the relatively long gestation periods of airplanes and launch vehicles, the technology needed to survive 

reentry matured quickly, largely in response to national security concerns. The 

warheads developed for ballistic missiles during the Cold War led directly to 

the capsules that first allowed humans to venture into space. Of course, it is 

important to recognize that the requirements for human entry are significantly 

 1. Basic histories of rocketry include Roger D. Launius and Dennis R. Jenkins, eds.,

1
Coming Home

more restrictive than those for warheads. Accordingly, the Mercury capsule 

could not simply be a warhead shell reconfigured for human occupancy. The 

heat shield required for the longer, lower g entries drove the development of 

several new technologies. Despite this, the technologies for ballistic reentry 

paved the way for human space flight and reentry.3

After the Columbia accident in 2003, it became popular to ask: Why does 

the Space Shuttle have wings? The reentry of spacecraft using ablative heat 

shields and parachutes seemingly worked well, so why abandon it? Many 

pointed to the presumably superior safety record of the Mercury, Gemini, and 

Apollo capsules and questioned the apparent change in course of using wings 

protected from melting by fragile thermal ceramic tiles on the Space Shuttle.4

However, it stands to reason that the capsules were the anomaly created 

because the schedule of the space race did not allow the creation of launch 

vehicles large enough to loft anything more sophisticated. In contrast, reusable, usually winged space vehicles—often discussed since Robert Goddard, 

Konstantin Eduardovich Tsiolkovskiy, and Hermann Oberth began writing 

about space flight during the early part of the 20th century—dominated thinking on the subject until the 1950s and the dawn of the space age. From the 

beginning, the spaceplane featured prominently in plans, although it should 

be noted that many (perhaps most) of the early concepts were actually longrange suborbital aircraft and not truly spacecraft. In part because of the dawning realization of the technology involved, realistic visions of true spacefaring 

vehicles did not begin to appear until the early 1950s.

Early Ideas About Spaceplanes

For many years after scientists and engineers began planning for space flight 

in the 20th century, the dominant vision of how to achieve this was via orbital 

spaceplanes that could be launched like rockets or perhaps even like an airplane 

 3. 

2
Creating Ballistic Reentry

from a runway. Those early space flight engineers envisioned the spaceplanes 

flying into orbit, undertaking their missions, and returning to land on Earth like 

airplanes at an airport. These winged orbital vehicles represented an extension 

of concepts associated with airplanes, including reusability; precise, controlled 

horizontal landings; acceleration to speeds sufficient to escape Earth’s gravity 

well (approximately 7.5 miles per second); and orbital velocities of 17,500 mph. 

Some concepts of these vehicles imagined they would be launched vertically 

like rockets while others speculated they would not be; and some concepts 

envisioned wings while others described lifting bodies in which the vehicle’s 

shape, rather than wings, would provide lift during reentry and landing.

The objective of airplanelike space operations dominated thinking 

about space flight from the beginning and had a profound influence on the 

development of reentry and recovery technology. Indeed, this dominant 

paradigm for space flight drove many thinkers to advocate reusable winged 

vehicles until the 1950s, when technological experimentation and the crushing 

pressure of the Cold War prompted its abandonment in favor of more readily 

achievable ballistic technologies. Nevertheless, until the recent past, NASA 

has pursued this concept as its holy grail, jettisoning the ballistic model used 

successfully in the Mercury, Gemini, and Apollo programs in favor of the 

reusable Space Shuttle that pointed efforts toward the spaceplane conception 

so long desired.5

At the same time, the goal of all these studies was to achieve (1) pinpoint 

recovery and (2) reusability, and wings were only one way to achieve these ends. 

These goals were temporarily set aside during the space race when it became 

obvious that a capsule shape with one-time use and a water recovery could 

accomplish the human-in-space goal quicker. However, there is a powerful 

psychological appeal to incorporating wings on spacecraft because it represents 

a tried-and-true method of achieving flight demonstrated from the Wright 

brothers’ era to the present. The primary objective of achieving an operational, 

low-cost space-access system drove this trend toward spaceplanes. But lowcost space flight has not developed as expected, and that may be an underlying reason that the nonreusable, water recovery concepts reemerged with the 

Constellation effort in the first decade of the 21st century. It is possible that 

when a true need for high-volume space-access flights arises, the soft-landing, 

reusable spacecraft concept will again emerge. Will this be achieved with wings 

 5. Roger D. Launius, “Between a Rocket and a Hard Place: The Challenge of Space Access,” in W. 

3
Coming Home

or landing rockets, with gliding parachutes or rotors, or through a separate 

approach? Only time will tell.

As a schoolteacher in Moscow, Konstantin Tsiolkovskiy began writing about 

rockets and space travel in 1898, when he submitted his “Investigating Space 

with Rocket Devices” article to the Russian Nauchnoye Obozreniye (Science 

Review) journal. This article presented years of calculations that laid out many 

of the principles of modern space flight and opened the door to future writings 

on the subject. There followed a series of increasingly sophisticated studies on 

the technical aspects of space flight. In the 1920s and 1930s, Tsiolkovskiy was 

especially productive, publishing 10 major works, clarifying the nature of bodies 

in orbit, developing scientific principles behind reaction vehicles, designing orbital 

space stations, and promoting interplanetary travel. The most important of these 

works was his Plan of Space Exploration, in which he described rocket-propelled 

airplanes with wings—a forerunner of the Space Shuttle. This led Tsiolkovskiy 

to deal with the aerodynamics of reentry in an exceptionally cursory manner. 

He also expanded the scope of studies on many principles commonly used in 

rockets today, which include specific impulse to gauge engine performance, 

multistage boosters, fuel mixtures such as liquid hydrogen and liquid oxygen, the 

problems and possibilities inherent in microgravity, the promise of solar power, 

and spacesuits for extravehicular activity. Significantly, Tsiolkovskiy never had 

the resources—perhaps not even the inclination—to experiment with rockets 

himself. During Tsiolkovskiy’s lifetime, much of this work went unnoticed 

outside Russia.6

In the West, one of the first popular spaceplane concepts was Robert 

Goddard’s turbine rocket ship, which appeared in both Scientific American

and Popular Science in 1935. Famous for his successful efforts to build liquidfueled rockets—the first such launch taking place in 1926, with the succession 

of progressively more sophisticated rockets following until his death in 1945—

Goddard designed his vehicle with elliptical wings and a unique dual-mode 

propulsion system. When outside the atmosphere, the vehicle used a liquidfueled rocket engine, as would be expected of a vehicle designed by the person 

who literally invented the technology. However, when the vehicle was flying 

within the atmosphere, two turbines moved into the exhaust stream of the 

rocket engine, driving conventional propellers affixed to either wing via long 

drive shafts. It was obvious from both the description and the limited artwork 

that Goddard was not an aircraft designer, and the concept had several major 

 6. Roger D. Launius, 

4
Creating Ballistic Reentry

flaws. Nevertheless, it helped popularize the idea that space vehicles would be 

extensions of normal aircraft, complete with wings, tails, and landing gear.7

Meanwhile, others were putting form to the theoretical concept of spaceplanes. The Viennese scientist Max Valier believed that rocket engines would 

eventually replace the internal combustion engine and lead, by natural evolution, to winged spaceplanes that would travel back and forth to Earth orbit 

and the planets.8

 In early 1928, Valier convinced Fritz A. H. von Opel, heir 

to the Adam Opel automobile empire, that the rocket could power a car as a 

publicity stunt to increase sales. The pair soon moved on to what Valier really 

wanted to do: build a rocket-powered aircraft. Using two 44-pound-force 

black-powder rockets attached to a 14-foot-long glider, Friedrich Stamer made 

the first piloted rocket-powered flight on June 11, 1928. Unfortunately, one 

of the rockets exploded during the second flight and fire consumed the glider. 

This did not stop the experiments, however, and on September 30, 1929, von 

Opel flew a 16-foot-long glider for almost 2 miles in front of a large crowd in 

Frankfurt am Main, Germany. A hard landing destroyed the glider and injured 

von Opel. The concept of rocket-powered aircraft became firmly established, if 

not well demonstrated. Later, during World War II, all combatants developed 

rocket-power-assisted technologies, usually called jet-assisted at the time, for 

aircraft to take off from short runways.9

At a fundamental level, the desire to build a spaceplane was inextricably 

linked to the development of hypersonic flight technology, which was in its 

infancy prior to World War II. Indeed, many viewed even hypersonic flight as 

 7. Robert H. Goddard, “A New Turbine Rocket Plane for the Upper Atmosphere,” 

5
Coming Home

difficult and impractical, realizing as they did that the characteristics exhibited by objects moving at Mach 5 or greater were exceptionally, perhaps even 

unfathomably, complex. They knew that when an aircraft reached hypersonic 

speed, the amount of pressure it generated increased to 25 times the atmospheric pressure.10

Researchers also found that as shock waves at those speeds grew in intensity, 

their angles bent as they passed between the shock wave and the surface of the 

vehicle. This region of multiple shock waves, called the shock layer because it 

was very small, could interact with the boundary layer. The body of the aircraft 

then experienced significant turbulence, and the air around the aircraft became 

a swarming jumble of hot gases whose intense heat transferred to the aircraft. 

Once the shock waves have conformed to the windward shape of the vehicle, 

however, the performance and stability are well described by simple Newtonian 

flow theory. Even though stability and controllability were easily achievable, 

heat transfer remained complex and difficult to characterize and control.11

The interaction between the shock layer and the boundary layer affected 

drag and caused unusual thermal reactions, contributing significantly to aerodynamic heating, a product of the friction of the high-speed airflow over the 

vehicle’s surface. Because of these complex interactions, the behavior and characteristics of hypersonic airflow proved extremely difficult to analyze. Only 

Goddard, Sänger, von Braun, and a handful of other space flight advocates and 

experimenters worked seriously to understand hypersonic flight and, mostly, 

to learn how to achieve the speeds necessary to reach Earth’s orbit.12

None of this had much to do with the reality of getting to orbit. It was, 

however, critical to the process of reaching into space; without achieving that 

goal, reaching reentry and recovery from space flight was a moot issue.

 10.

6
Creating Ballistic Reentry

Sänger, Bredt, and the Silverbird

Perhaps Sänger made the most interesting of the early studies into orbital 

winged space flight. Inspired by Oberth’s book By Rocket into Planetary Space, 

Sänger changed his university studies from civil engineering to aeronautics. 

While he was a doctoral candidate at the Viennese Polytechnic Institute in 

1929, Sänger conceptualized a reusable, rocket-powered spaceplane with 

straight wings. However, like many of the other early spaceplane concepts, 

most of the Sänger designs were suborbital aircraft, not true spacecraft. In late 

1933, Sänger privately published Techniques of Rocket Flight, which contained 

a detailed description of the vehicle that was called the “Silverbird.”13 In collaboration with mathematician Dr. Irene Bredt, whom he later married, Sänger 

continued to refine this design for the next 30 years.14

Propelled by a liquid-fueled rocket engine, the Silverbird was capable of 

reaching 6,600 mph at an altitude of 100 miles. Due to the relatively low 

speed, this was not an orbital vehicle, despite the high altitude it theoretically 

could achieve.15 Instead, the Silverbird used a suborbital trajectory to deliver 

a payload halfway around the world using a technique that included a series 

of semiballistic skips off the atmosphere that Sänger called “dynamic soaring.” 

As the vehicle followed a ballistic path from space into the upper atmosphere, 

each skip reduced the forward speed slightly, and, consequently, each skip was 

a little lower and shorter. Sänger believed this strategy would increase range 

 13.

7
Coming Home

The Sänger Silverbird in wind tunnel testing in the United States after World War II. NASA.

and significantly reduce the reentry16 thermal load by allowing heat to radiate 

into space after each skip.17 Despite this hypothesis, Sänger did little actual 

investigation into reentry conditions, concentrating more on the propulsive 

aspects of the vehicle. Of course, even if he had wanted to study reentry effects, 

little theoretical and no practical basis existed to do so.18

Accordingly, while there was some understanding of pressures and heating 

in the period before World War II, the majority of studies into the challenges 

of hypersonic flight had to wait until the postwar era. For one, the large V-2 

 16.

8
Creating Ballistic Reentry

rockets developed in Germany during World War II forced the development 

of concept studies with promising results.19

Around the same time, in 1946, the work of German engineers Sänger 

and Bredt reached the U.S. Navy’s Bureau of Aeronautics, changing many 

people’s perspectives about the possibility of hypersonic flight. The two engineers argued that a rocket-powered hypersonic aircraft could be built with 

only minor advances in technology. NACA engineers found this concept quite 

stimulating, and along with studies by German and Russian researchers, this 

prompted engineers such as John Stack and John V. Becker at the Langley 

Memorial Aeronautical Laboratory to begin the lengthy process of gaining 

approvals to explore this flight regime.20

Skip-gliding remained in favor for nearly 20 years, until research at the 

NACA Ames Aeronautical Laboratory in the 1950s showed the technique 

would not reduce heating as effectively as would a direct reentry. While radiating heat into space during skip-gliding proved a highly effective and useful 

method of rejecting heat and has subsequently proven a primary method of 

controlling temperatures in the true space environment, the direct reentry 

approach proved more expeditious for those early efforts. Although peak temperatures during skip-gliding would be slightly lower than those during direct 

reentry, the time the vehicle spent exposed to those temperatures was considerably longer (cumulatively, through all the skips). This proved a problem for 

heat-sink and ablative thermal protection systems, but research soon devised 

hot structures that used radiation or active cooling where the temperatures 

were stabilized and not subjected to total heat-load restrictions.

The First Ballistic Missiles

During the 1920s and 1930s, organizations of rocket enthusiasts emerged in 

several European countries, with the most influential being the German Society 

for Space Travel, or Verein für Raumschiffahrt (VfR). Spurred by the theoretical arguments of Oberth and Valier, the VfR emerged soon after its founding, 

 19.

9
Coming Home

The German Army in World War II preparing the V-2 for a test. NASA.

10
Creating Ballistic Reentry

on July 5, 1927, as the world’s leading space-vehicle design group. Specifically 

organized to raise money, it tested Oberth’s rocketry ideas. It soon proved successful in building a base of support in Germany, publishing a magazine and 

scholarly studies, and in constructing and launching small rockets. From the 

beginning, one of the VfR’s strengths, however, was its ability to publicize both 

its activities and the dream of space flight.21

The VfR made good on some of those dreams on February 21, 1931, when it 

launched the liquid-oxygen/methane-liquid-fueled rocket HW-1 near Dessau, 

Germany, to an altitude of approximately 2,000 feet. The organization’s public 

relations arm went into high gear after this mission and emphasized the launch’s 

importance as the first successful European liquid-fueled rocket flight.22 Von 

Braun, then a neophyte learning the principles of rocketry from Oberth and 

Valier, was both enthralled with this flight and impressed with the publicity 

it engendered. Later, he became the quintessential and movingly eloquent 

advocate for the dream of space flight and a leading architect of its technical 

development. He began developing both skills while working with the VfR.23

In the end, the VfR conducted numerous static firings of rocket engines and 

launched several small rockets.

As Germany rearmed in the 1930s, its army collected a host of rocket engineers, many from the VfR, at a military test facility at Peenemünde, Germany, 

on the Baltic coast, and initiated the development of a short-range ballistic 

missile under the tutelage of von Braun.24 By 1943, after numerous frustrations, they had a working vehicle. Joseph Goebbels christened the rocket (called 

Assembly 4, or A-4, at Peenemünde) “Vergeltungswaffe Zwei” (Vengeance 

Weapon Two), or V-2. The missile became operational in September 1944 and 

over 3,500 were launched in aggression, mainly against Antwerp and London.

 21.

11
Coming Home

The designers of the first modern ballistic missile did not have much need to 

worry about reentry effects. The A-4 reached an altitude of only 50 to 60 miles, 

after which it descended on a ballistic trajectory with a maximum velocity of 

approximately Mach 4 and a range of 200 miles. Although this velocity made 

the weapon invulnerable to anti-aircraft guns and fighters, the temperatures it 

experienced while streaking through the thick atmosphere near Earth’s surface 

were well within the heat and strength capability of conventional steel and 

aluminum alloys.25

While deployed too late to alter the outcome of the war, the A-4 not only 

radically changed the concept of weapon delivery but also offered the promise 

of space flight in the not too distant future. It became the basis for all that followed in the Soviet Union and United States. By the end of World War II, the 

popular conception of a missile along the lines of the German A-4—a slender 

cylinder with a pointed nose and sweptback fins—had become the norm. This 

basic shape, used for decades in various books and movies, followed the general 

streamlining trends in aviation that began in the 1920s. The experience with 

the A-4 had shown the shape worked adequately, although few truly appreciated the relatively low velocities achieved by the missile.

During the late 1940s, the United States began developing an increasingly sophisticated family of ballistic missiles. As a direct result of war booty 

captured from the Germans at the end of World War II, the United States 

used the V-2 first stage with a WAC Corporal rocket as a second stage and 

tested hypersonic concepts at the White Sands Proving Ground, NM. The 

V-2/WAC Corporal combination became the first manufactured object to 

achieve hypersonic flight. On February 24, 1949, its upper stage reached a 

maximum velocity of 5,150 mph—more than five times the speed of sound—

and a 244-mile altitude. The vehicle, however, burned up on reentry, and only 

charred remnants remained, demonstrating the need to explore the challenges 

of reentry and recovery from space.26

First launched in 1953, the Redstone was a direct descendant of the German 

A-4, designed largely by the same team from Peenemünde led by von Braun. 

The missile flew about as high and far as the A-4, but it carried a much larger 

payload (6,500 versus 2,500 pounds) and reached a maximum velocity of Mach 

 25.

12
Creating Ballistic Reentry

Bumper 8, on July 24, 1950, was the first launch from Cape Canaveral, FL. NASA.

5.5 during reentry. Improved versions, called Jupiter, that were developed a 

few years later extended the maximum range to 1,500 miles and reached Mach 

15 during reentry from a 390-mile peak altitude. This significant increase in 

performance brought new problems for the engineers as they explored how to 

recover portions of the vehicle.

The Problem of Reentry

Given the technology available during the 1950s—indeed, even 60 years 

later—getting into space has been largely an exercise in brute force. Sufficient 

power, usually in the form of rocket engines, must be available to accelerate the 

13
Coming Home

vehicle to slightly over 17,500 mph to attain a stable, low-Earth orbit. Unlike 

aircraft that can match their rate of climb to the available thrust, conventional 

rockets must always accelerate at greater than 1 g since they rely solely on power, 

not lift, to reach in space. The only exception to this 1 g–plus requirement for 

vertical liftoff is a winged rocket (like the X-15) that can fly trajectories that 

are not vertical and use lift as well as thrust. The exit trajectory is usually very 

steep, so the vehicle is well above the dense atmosphere before it reaches high 

speeds, a tactic that minimizes both drag and aerodynamic heating.

Oddly, reentry went largely unnoticed in most early literature on space 

flight. Although many people had speculated on how to get into space, relatively few were thinking about how to get back. One exception was Walter 

Hohmann, a German civil engineer and member of the VfR, who in 1925 wrote 

Die Erreichbarkeit der Himmelskörper (The Attainability of Celestial Bodies).27

Although the book was primarily concerned with the derivation of optimum 

transfer trajectories from Earth to other planets, Hohmann also thought about 

how to return to Earth, and he was especially concerned about the effects of 

atmospheric heating during reentry.28 Unfortunately, Hohmann lacked any 

practical knowledge of and any meaningful way to investigate the environment 

in which a vehicle would have to operate in returning from space. His work was 

entirely theoretical. Although he failed to come up with specific solutions to the 

problem, Hohmann categorized the problem and theorized several techniques 

that returning spacecraft might use, including variable-geometry wings and 

external insulation. Despite the work by Hohmann and a few others, none of 

the technologies available at the time would have permitted building a reentry 

vehicle, even if a means of launching it had existed.29

Twenty years later, at the end of World War II, it was apparent that space 

flight was not too far off, although it shocked nearly everyone when it arrived 

as soon as it did. One of the first serious postwar studies of space flight was 

the Preliminary Design of an Experimental World-Circling Spaceship by Project 

 27.

14
Creating Ballistic Reentry

RAND at Douglas Aircraft. As with Sänger and Hohmann, RAND attempted 

to explore the entry environment but lacked any real basis for developing a 

theory. Nevertheless, the study concluded the following:

An investigation was made of the possibility of safely landing the 

vehicle without allowing it to enter the atmosphere at such great 

speeds that it would be destroyed by the heat of air resistance. 

It was found that by the use of wings on the small final vehicle, 

the rate of descent could be controlled so that the heat would be 

dissipated by radiation at temperatures the structure could safely 

withstand. The same wings could be used to land the vehicle on 

the surface of the earth.30

It was a credible endorsement of a spaceplane. Of course, it was not that 

simple. Although the RAND engineers expected to be able to control the rate 

of descent by using the wings and radiating the heat while still at high velocity, 

this approach ignored the characteristics of the atmosphere at extreme altitudes, 

which were largely unknown at the time. This is the same concept that was 

used on the X-20 Dyna-Soar glider in the late 1950s and early 1960s and was 

successfully demonstrated on the ASSET vehicle.31

Ultimately, most engineers working on the problems of space flight were 

convinced that it would prove impossible to fly back from orbit, at least in 

the short run.32 Even the legendary Theodore von Kármán was worried. In his 

1954 history of aeronautics, von Kármán observed of a reentry vehicle, “At 

such speeds, probably even in the thinnest of air, the surface would be heated 

beyond the temperature endurable by any known material. This problem of the 

temperature barrier is much more formidable than the problem of the sonic 

barrier.”33 Thus was born the thermal barrier terminology that would define 

much of the aeronautical research during the late 1950s and early 1960s.

 30.

15
Coming Home

Nuclear Warheads and the Blunt-Body Theory

Initially, all objects returned from very high altitudes using a ballistic reentry in 

which the force was parallel to the line of flight and the trajectory was always in 

the form of a parabola. Arrows and artillery projectiles have long used ballistic 

trajectories, and long-range missiles initially followed suit. The peak altitude of 

the early missiles, such as the A-4 and the Redstone, was low enough that the 

vehicle did not reach velocities that produced atmospheric friction sufficient to 

create a serious heating problem. However, as the maximum range of missiles 

grew longer, the peak altitudes and velocities increased and heating became a 

major concern. To attain a range of one-quarter the circumference of Earth, an 

object must reach a velocity of about 23,000 feet per second (ft/sec) (15,500 

mph). The vehicle then has about eight times the kinetic energy required to 

turn ice into steam.34 During reentry, almost all of this kinetic energy converts 

into heat through atmospheric heating.35

The steep trajectories used by the initial medium-range ballistic missiles 

resulted in a reentry in which temperatures exceeded 12,000 degrees Fahrenheit 

(°F) within the stagnation zone immediately in front of the vehicle. This is 

twice as hot as the surface of the Sun. The heat generated outside the boundary 

layer by shock wave compression, which was not in contact with the vehicle, 

dissipated harmlessly into the surrounding air. The heat within the boundary 

layer, and in direct contact with the structure, however, was hot enough to 

melt the vehicle.

At Convair, Karel Bossart and his team of engineers were developing the 

Atlas intercontinental ballistic missile (ICBM), the first of the long-range ballistic missiles. In a radical departure from previous missiles, Bossart decided 

that only the warhead would return to Earth, not the entire missile. Convair 

knew the warhead—usually referred to as a “reentry vehicle” since it sounded 

less sinister—would need protection from the thermal effects of reentry, but 

it had little test-validated theory to allow a design to proceed. Using a digital 

computer, one of the first applications of the device as a design simulation tool, 

Convair engineers began examining various shapes for the warhead.

Since the 1920s, airplanes and the missiles that followed became increasingly streamlined. The goal, of course, was to go faster by creating less drag. 

These lessons seemed particularly applicable to the fastest vehicle yet: the 

 34.

16
Creating Ballistic Reentry

ballistic missile. When Convair engineers fed data into their digital computer, 

they did so with the bias that more streamlined was better. So, unsurprisingly, 

the computer-generated results favored a slender, pointed-nose reentry vehicle 

that looked much like the Bell X-1 that first broke the sound barrier.

Convair and NACA researchers then tested the shape in the few hypersonic 

wind tunnels that existed and used rocket-boosted free-flight models at the 

NACA Pilotless Aircraft Research Station on Wallops Island, VA. These tests 

produced an unexpected result: the warhead would absorb so much heat that 

it would vaporize as it reentered the atmosphere. Suddenly, the earlier experiences with the medium-range warheads began to make sense.

This realization, that the slender aircraft body suited to supersonic flight was 

inappropriate for the hypersonic flight of a ballistic missile, led weapons designers to explore the phenomenon more thoroughly. This yielded the observation 

that a blunt-nose body experienced much less heating than a pointed body, 

which would burn up before reaching Earth’s surface. One of the members 

of a panel charged with the reentry problem was NACA researcher H. Julian 

Allen.36 The 42-year-old Allen joined the NACA in 1935 and had been chief 

of the high-speed research division at NACA’s Ames Aeronautical Laboratory, 

now Ames Research Center, since 1945. Allen enlisted the assistance of Alfred 

J. Eggers, Jr., a 30-year-old aerodynamicist who had joined the NACA directly 

out of school in 1944.

The blunt reentry body theory discovered in 1951 by Allen created a stronger shock wave at the nose of the vehicle and dumped a good deal of the reentry 

heat into the airflow, making less heat available to heat the reentry vehicle 

itself. This finding was so significant and in such sharp contrast with intuitive 

thinking that Allen’s work fundamentally reshaped the course of reentry and 

recovery studies and provided the basis for all successful reentry vehicles since.37

By June 1952, Allen and Eggers had found a theoretical solution to the aerodynamic heating problems of ballistic reentry vehicles. The two researchers deduced 

that about half the heat generated by aerodynamic friction transferred into the 

warhead, quickly exceeding its structural limits. The obvious solution was to deflect 

the heat away from the vehicle. The breakthrough was in how to accomplish this. In 

place of the traditional sleek missile with a sharply pointed nose, the researchers proposed a blunt shape with a rounded bottom. When reentering the atmosphere,the 

 36.

17
Coming Home

H. Julian Allen stands beside the observation window of the 8-by-7-foot test section of the 

NACA Ames Unitary Plan Wind Tunnel. He is best known for his “Blunt Body Theory,” which 

revolutionized the design of ballistic missile reentry shapes. NASA GPN-2000-001778.

18
Creating Ballistic Reentry

rounded body creates a powerful detached shock wave that deflects the airflow and 

its associated heat outward and away from the reentry vehicle. As Allen and Eggers 

observed, not only should pointed bodies be avoided, but the rounded nose should 

have as large a radius as possible. The blunt-body theory was born.38

The NACA researchers briefed members of the Atlas ballistic missile development team and select others in September 1952. A secret NACA research 

memorandum was published for those with appropriate clearances on April 

28, 1953, but it would be 5 years before the concept was revealed to the community at large.39 Interestingly, Allen and Eggers had assumed for the purposes 

of their research that any warhead would probably use a liquid cooling system 

to protect it from the residual heat, although the exact nature of an operational 

thermal protection system was not investigated. The researchers also pointed 

out that the blunt-body theory worked best for lightweight reentry vehicles 

and that heavier vehicles likely would need to return, at least partially, to the 

traditional long, slender shapes once adequate thermal protection systems had 

been developed.40

 38.

19
Coming Home

A nonlifting reentry vehicle is measured by its ballistic coefficient (usually 

designated beta and signified with the Greek symbol “β”), a function of mass, 

size, and drag coefficient.41 Vehicles with a high ballistic coefficient—usually 

long, slender, and with a pointed nose—plunge through the upper atmosphere 

and experience most of their deceleration in the thick lower atmosphere, where 

they get very hot for a brief period. Vehicles with a low ballistic coefficient—

blunt bodies—experience most of their deceleration in the thin upper atmosphere. They take longer to slow down and generate less heat but experience 

this heat over a longer period of time. The total heat load (temperature multiplied by time) is essentially the same for both types of vehicles, but the longer, 

lower-magnitude temperatures of the blunt body were easier for the material 

available at the time to absorb.

Although the NACA showed that a blunt body was desirable, exactly what 

shape this should assume was the subject of vigorous debate. Researchers tested 

spheres, cylinders, and blunted ogives. Allen pioneered much of the practical 

research on the blunt-body theory and conducted it in an innovative free-flight 

tunnel at Ames. The $200,000 tunnel had an 18-foot-long test section that was 

1 foot wide and 2 feet high. Air was forced into the test section from one direction, and a small model was shot from a compressed-air cannon in the opposite 

direction. Seven Schlieren cameras provided shadowgraphs that showed the 

airflow characteristics and shock waves around the model as it passed through 

the test section. Using this device, researchers could simulate speeds of about 

Mach 15, which was considerably greater than the few hypersonic wind tunnels 

or any of the experimental rocket-powered research airplanes.42

Eggers contributed another test facility, which was called the Atmospheric 

Entry Simulator. This was a straight, trumpet-shaped supersonic nozzle 20 

inches in diameter and 20 feet long. A hypervelocity gas gun launched a smallscale model upstream through the nozzle into a settling chamber. While in free 

flight through the contracting nozzle, the model passed through ever-denser 

air, closely approximating the plunge of a warhead through the atmosphere. 

Using a model only 0.36 inch in diameter and weighing only 0.005 pounds, 

Eggers could simulate the aerodynamic heating of an object 3 feet in diameter 

and weighing 5,000 pounds.43

 41.

20
Creating Ballistic Reentry

In addition, testing was conducted using the 11-inch hypersonic wind 

tunnel developed by Becker at Langley and free-flight models at the Pilotless 

Aircraft Research Station on Wallops Island. The U.S. Air Force (USAF) and 

Navy also procured 26 three-stage solid-propellant X-17 research vehicles from 

Lockheed to study reentry techniques up to Mach 15 and 500,000 feet in 

altitude. The first launch was on April 17, 1956, from Cape Canaveral, FL.44

With this basic knowledge, the decision to pursue more research into the 

hypersonic arena came on June 24, 1952, when the NACA Committee on 

Aerodynamics passed a resolution to “increase its program dealing with the 

problems of unmanned and manned flight in the upper stratosphere at altitudes between 12 and 50 miles, and at Mach numbers between 4 and 10.” The 

NACA Executive Committee ratified this decision the following month and 

appointed a study group led by Clinton E. Brown to ascertain the feasibility 

of the project.45

The report prepared by Brown’s committee served as a catalyst for a discussion of hypersonics at the October 1953 meeting of the U.S. Air Force’s 

Scientific Advisory Board (SAB) Aircraft Panel. This panel also provided additional support for hypersonic research. Interestingly, the SAB panel member 

from Langley, Robert R. Gilruth, director of the Pilotless Aircraft Research 

Division and later the director of human space flight for NASA, became 

enthralled with these possibilities and played an important role in negotiating 

a joint project by the Air Force and the NACA.46

As Langley engineer Becker remembered the story in 1968:

By 1954, we had reached a definite conclusion: the exciting potentialities of these rocket-boosted aircraft could not be realized without major advances in technology in all areas of aircraft design. In 

particular, the unprecedented problems of aerodynamic heating 

and high-temperature structures appeared to be so formidable 

that they were viewed as “barriers” to hypersonic flight.... But in 

1954 nearly everyone believed intuitively in the continuing rapid 

increase in flight speeds of aeronautical vehicles. The powerful 

new propulsion systems needed for aircraft flight beyond Mach 

 44.

21
Coming Home

3 were identifiable in the large rocket engines being developed in 

the long-range missile programs. There was virtually unanimous 

support for hypersonic technology development.47

Therefore, despite technological challenges, the successful hypersonic 

research led inevitably to greater understanding of the technology needed for 

safe reentry.

Even though a desirable shape had been determined, the blunt-body theory 

was, in itself, not a deployable concept. Although it was a major breakthrough, 

a blunt body still experienced peak temperatures of over 5,000 °F. In the 1950s, 

the aviation community was using mostly conventional metal alloys, such as 

Monel K (for the Douglas X-2) and Inconel X (for the North American X-15), 

but these still could withstand at most 1,200 °F. Allen’s assumption that an 

adequate implementation of a thermal protection system could be designed 

did indeed place a great enough burden on other researchers and engineers, 

especially materials scientists, to make it a reality.

Fortunately, the medium-range ballistic missiles experienced such a shortduration reentry that the total heat load could be absorbed by large heat sinks. 

These heat-sink materials needed to possess certain characteristics. For instance, 

they needed to be amenable to fabrication into the required shapes. Although, 

given the small number of warheads being deployed, they did not necessarily 

have to be mass producible or inexpensive. Most importantly, they had to 

exhibit a high strength-to-weight ratio at elevated temperatures, have a high 

melting or sublimation point,48 have high thermal conductivity and ductility,49

have a low coefficient of thermal expansion, and be resistant to oxidation. 

Accordingly, researchers at any number of Government installations and their 

contractors began investigating heat-sink materials.50

Those researchers arranged possible materials into several groups according to 

shared characteristics and began systematically testing the limits of each group. 

The first group included heavy ductile metals with high thermal conductivity, 

such as copper, gold, and silver. These materials have relatively low melting 

points but maintain their structural integrity and shape under severe heat loads, 

and they are moldable to almost any shape using conventional machine tools. 

 47.

22
Creating Ballistic Reentry

X-15-3 (56-6672) flies over the Mojave Desert in the 1960s. Ship #3 made 65 flights during 

the program, attaining a top speed of Mach 5.65 and a maximum altitude of 354,200 feet. 

NASA E-USAF-X-15.

The second group was composed of medium-density refractory metals such 

as nickel-chromium-iron stainless steels (such as Inconel) and various cobalt 

alloys developed for jet-engine turbine blades. These metals retain strength at 

elevated temperatures, have a relatively high resistance to oxidation, and possess 

low values of specific heat and thermal conductivity. The third group included 

lightweight metals such as beryllium, with high strength-to-weight ratios, high 

specific heat and thermal conductivity, high resistance to short-time oxidation, 

and reasonably high melting points. However, these metals suffered from poor 

ductility and were difficult to manufacture into complex shapes. The last group 

included semimetals such as carbon (graphite) that have very high sublimation 

points, high thermal conductivity and specific heat, and low density. Most of 

these materials, however, exhibit poor high-temperature oxidation resistance, 

low structural strength, and are difficult to fabricate.51

NACA researchers concentrated on a single member of each group, choosing 

copper, Inconel X,52 beryllium, and graphite because they were commercially 

 51.

23
Coming Home

available and relatively well understood. Researchers manufactured a 1-inchthick slab of each material for a series of tests, and the results were telling. For 

instance, the front face of each slab was heated to 2,000 °F and the temperature 

of the rear face was measured to determine the heat absorption capacity. The 

copper slab had a rear-face temperature of only 190 °F, indicating that the 

material absorbed most of the thermal load. On the other hand, Inconel X 

had a rear-face temperature of 1,600 °F, meaning the metal was ineffective at 

absorbing heat.53 Beryllium and graphite had rear-face temperatures of 840 °F 

and 1,240 °F, respectively.54

Researchers also tested the effect of slab thickness on the maximum surface 

temperature during reentry. Inconel X was excluded from this test since the 

surface of the slab reached its melting temperature regardless of the thickness. 

The amount of heat absorbed per unit weight of material was measured at a 

temperature corresponding to 75 percent of the melting (or sublimation) point. 

In this test, copper absorbed 84 British thermal units (BTU) per pound, beryllium 534 BTU, and graphite 1,980 BTU. This made beryllium and graphite 

greatly superior to copper as heat-sink materials, especially given that copper 

was 24 times heavier than graphite and 6 times heavier than beryllium.55 Given 

the low performance of the early missiles, weight was a primary consideration; 

each pound of warhead weight required approximately 50 pounds in gross liftoff weight (mostly in the form of propellants, but also in the structure needed 

to support the propellants).56

Unfortunately, graphite is subject to rapid oxidation and vaporization at 

high surface temperatures, which often led to structural failures. Eventually, 

protective coatings to isolate the material from the airstream would be developed as part of the Dyna-Soar and Space Shuttle programs, but this was much 

too late to help the original ballistic missile programs.57

 53.

24
Creating Ballistic Reentry

The blunt-body theory led directly to the Mk I58 reentry vehicle developed 

by General Electric for the intermediate-range Thor and early Atlas-B/C59

ballistic missiles. Despite the weight penalty, General Electric selected 1,200 

pounds of copper for the Mk I as a matter of expediency, since the material 

was well understood and readily available. A slightly improved Mk II reentry 

vehicle initially used a similar amount of copper.60

General Electric had decided against beryllium because its low ductility posed severe fabrication difficulties, and the company lacked the time to 

develop new processes to manufacture the metal. The Navy, on the other 

hand, had more time and selected a beryllium heat sink for the initial Mk 1 

warheads61 for its Polaris submarine-launched ballistic missile. Based on the 

success of the Navy program, beryllium soon replaced copper in later versions 

of the Air Force Mk II warhead.

The Air Force Mk II reentry vehicle, and its similar, but unrelated, Navy Mk 

1 counterpart, had a low ballistic coefficient and a large heat sink. Although 

both of these vehicles offered an immediate solution to the reentry problem, 

experience quickly showed that heat-sink warheads could fail in several ways. 

For instance, they could lose their strength because of the high temperatures 

and disintegrate under the air and deceleration loads. Alternately, they could 

fail mechanically due to thermal stresses or fail materially through spalling,62

melting, or sublimation. At the extreme, the warhead could combust due to 

an unstable exothermic reaction of the material with the airstream.

Even when the early warheads worked as designed, they had significant 

defects as weapon delivery systems. For instance, they spent a lot of time in 

the upper atmosphere, trailing a stream of ionized gas from the melting heat 

sink that showed up on radar, which made them very susceptible to interception by antiballistic missiles. In addition, decelerating in the upper atmosphere 

 58.

25
Coming Home

meant that the warhead was traveling relatively slowly when it reached the 

lower atmosphere, making it susceptible to winds and frequently causing it to 

miss its target by miles.

Engineers soon discovered that accuracy could be improved by increasing 

the descent velocity so the reentry vehicle was less affected by winds. They had 

come full-circle as what they now wanted was a vehicle with a high ballistic 

coefficient, or exactly the opposite of the blunt-body concept that had made 

reentry possible in the first place. This, however, increased the heat load past 

the absorption capability of a heat sink of any practical size, so researchers 

began investigating methods to supplement or replace the metallic heat sinks.63

Thermal Protection Systems

In the broadest sense, there are three thermal protection concepts: passive, 

semipassive, and active. The type of protection on any space-venturing vehicle 

or, more precisely, on any given area of a vehicle depends largely on the magnitude and duration of the heat load as well as various operational considerations.

As the name implies, passive thermal protection systems have no moving 

parts. They are the simplest but, until the advent of the Space Shuttle, had the 

least capability. These concepts have fallen into three general categories: heat 

sink, hot structure, and insulated structure. The heat sink absorbs almost all of 

the incident heat and stores it in a large, usually metallic mass. Additional mass 

may be added to increase the heat storage capability, but in general the concept 

is limited to short heat pulses. A hot structure allows the temperature to rise until 

the heat being radiated from the surface is equal to the incident heating, much 

like the heating element of an electric stove. This concept is not limited by the 

duration of the heat pulse but is restricted to the acceptable surface temperature 

of the structural material. The Inconel X hot structure of the X-15 research airplane could withstand temperatures up to about 1,200 °F, which was about the 

maximum temperature for the concept. Insulated structures use an outer shell 

that radiates most incident heat away from an underlying structure protected by 

a layer of some insulating material, usually high-temperature ceramic-fiber batt 

insulation. Both the magnitude and duration of the heat pulse are limited for 

insulated systems, but it allows lower-temperature structural materials to be used.

 63.

26
Creating Ballistic Reentry

There are two basic semipassive concepts. Heat pipes are attractive where 

there is a localized area of high heating with an adjacent area of low heating. 

A working fluid vaporizes in the high-heat area, and the vapor flows naturally 

to a cooler region where it condenses and the heat is rejected. The condensed 

working fluid is returned to the high-heat region by capillary action.

The other semipassive concept is ablation. Ablation is a process in which a 

material (ablator) sacrifices itself to protect the underlying structure. However, 

the ablator is consumed in the process, thus limiting the duration of its operation and generally eliminating reuse. These ablative materials may be chemically 

constructed (usually some form of fiberglass or a spray-on resin-based coating) or made from natural materials (both the Chinese and Soviets/Russians 

used oak wood on some early reentry vehicles). As heat is absorbed by the 

ablator, part of the material decomposes into a gas, which carries heat away 

from the surface and into the boundary layer where the chemical byproducts 

further block convective heating. Near the surface, exposure to higher temperatures causes more complete pyrolysis, thus the surface shows a more complete 

decomposition than the inner layers and forms a char, much like charcoal.64

The vaporization of an ablator is an endothermic reaction; i.e., it needs 

energy input to proceed. A proper ablator will absorb the heat flux incident to 

its surface and vaporize at a rate proportional to the magnitude of the heat flux. 

The heat is carried away with the vaporized material, increasing the effectiveness 

of the thermal protection system. The vaporization leads to a thinning of the 

ablator, resulting in an upper limit on the amount of heat it can absorb. Perhaps 

the most unappreciated advantage of ablation is that the coolant (the gas generated by pyrolysis), after accepting all of the heat it is capable of absorbing, is 

automatically jettisoned. It requires no complex pumps or other mechanical 

devices, saving weight and complexity. In addition, as the ablated material 

is jettisoned, the ensuing heat load is lessened by the continuous reduction 

of mass.65 This was particularly important for the early ballistic missiles and 

human space programs for which weight was critical due to launch vehicles 

with limited capability.66

Three different active cooling concepts have been widely investigated. 

Transpiration and film cooling operate on a principle similar to the ablators; 

 64.

27
Coming Home

coolant ejected from the vehicle surface blocks most of the heat from reaching 

the underlying structure. These concepts use a pump to bring liquid coolant 

from a remote reservoir onto the surface of the vehicle. Transpiration cooling 

ejects the coolant through a porous surface, whereas film cooling uses discrete 

slots to flow coolant along the vehicle parallel to the airflow. The mass penalties associated with expendable coolant, of course, have usually limited these 

concepts to small, high-heat regions. The last concept—one first envisioned by 

Wernher von Braun and pioneered by Bell Aircraft—circulated coolant (water 

or liquid metal) around the hot area then through a heat exchanger or radiator, 

much like the engine-cooling system on an automobile. This water-wall concept allowed almost any heat load to be tolerated, but the coolant and radiator 

systems added considerable weight and were therefore not truly practicable.

Better Warheads

The General Electric Mk I and Mk II heat sink passive reentry warheads provided the United States with an immediate method to field intermediate-range 

Thor and Atlas-B/C ballistic missiles. However, their limitations prevented 

their use on warheads intended for the longer-range ICBMs such as Titan 

and the improved Atlas-D/E/F. Avco Manufacturing Corporation (now part 

of Textron) and General Electric, along with numerous military and NACA 

laboratories, began investigating other types of thermal protection systems for 

these warheads. Initially, many researchers thought the answer could be found 

in the active cooling concepts like those assumed by Allen and Eggers for the 

blunt-body theory, but these proved to be heavy and unreliable, particularly 

when subjected to the decelerations experienced by the ICBM reentry vehicles.

While the U.S. Air Force and its contractors were expending efforts on 

heat sinks, the U.S. Army was taking the lead in investigating ablators. At 

Redstone Arsenal, AL, the Vitro Corporation was using the exhaust from liquid 

rocket engines as a heat source to test a new family of ablators. On August 8, 

1957, a Jupiter-C launched from Cape Canaveral carrying a subscale reentry 

vehicle to an altitude of 600 miles and a range of 1,200 miles and proved the 

feasibility of the ablative-type nose cone during reentry.67 It should be noted, 

however, that this test was only meant to validate the reentry conditions for 

an intermediate-range Jupiter warhead. Whether an ablative ICBM warhead 

would work was still unproven, although in theory the concept could be scaled 

 67.

28
Creating Ballistic Reentry

up for those more advanced systems. The Air Force initially looked upon the 

Army’s claimed success with skepticism.68

The NACA was involved from the beginning in the warhead research, and it 

contributed to both the heat sink and the ablative experiments. Several NACA 

researchers that worked on the early problems of reentry would go on to have 

great influence on the human space flight program. Maxime A. Faget, Paul 

E. Purser, Robert O. Piland, and Robert R. Gilruth would all play important 

roles in future programs. Faget and Purser were also members of the Polaris 

Task Group, which provided advice to the U.S. Navy on the Polaris warheads, 

and the task group worked extensively with Lockheed in developing the Navy 

Mk 1 heat-sink warheads for the Polaris A1 and A2.69

NACA researchers also investigated ablative materials such as Teflon, nylon, 

and fiberglass using new test facilities. At Langley, these included an acid-ammonia rocket facility that could generate a maximum temperature of 4,100 °F and 

a gas velocity of 7,000 ft/sec, and an ethylene-air jet facility capable of yielding 

temperatures of 3,500 °F. Electric arc-jet facilities were built at both Langley 

and the Lewis Research Center, and each facility was capable of generating temperatures up to 12,000 °F. The NACA researchers proved the Army was right: 

ablative coatings were superior to heat sinks.70

General Electric saw the promise of ablators, resulting in the ablative Mk 

III warhead for the intermediate-range Jupiter-C, Thor, and Atlas-D missiles. 

The ablative heat shield consisted of nylon cloth impregnated with a phenolic 

plastic resin. High-temperature ablators were an important discovery for the 

mid-ballistic-coefficient warheads then under development and paved the way 

for the high-ballistic-coefficient warheads used today. Avco entered the fray 

with the ablative Mk 4 for the Atlas-E/F and the even longer range Titan I,71

as well as the ablative Mk 5 for the early Minuteman I. All future warheads 

used ablative heat shields as their ballistic coefficients became higher and their 

velocity during reentry increased.72

On the improved Mk 6 reentry vehicle for the Titan II, the use of a nylon 

phenolic ablator was so effective that significantly reduced bluntness was possible, with a half-angle of only 12.5 degrees. The 10-foot-long Mk 6 was 

the largest warhead developed by the United States, with an entry mass of 

 68.

29
Coming Home

A Titan II intercontinental ballistic missile with an Mk 6 reentry vehicle. USAF.

30
Creating Ballistic Reentry

7,500 pounds. Subsequent advances in nuclear weapon design allowed reentry 

vehicles to become significantly smaller, and the Avco Mk 11 warhead began 

the shift toward a more pointed body with an even higher ballistic coefficient.

Eventually, the reentry vehicles used by the later long-range ballistic missiles 

(such as the Minuteman III, Trident, and Peacekeeper) returned to the long, 

slender, pointed vehicles that had long been associated with space travel. New 

ablative coatings protected these reentry vehicles during the plunge through the 

atmosphere, minimizing the time the vehicle spent in the upper atmosphere and 

lowering its chances of detection and interception as well as significantly reducing 

the accuracy dispersions due to upper-altitude winds. The General Electric Mk 12 

reentry system for the Minuteman III, deployed in 1970, contained three separate 

warheads, each having its own pointed reentry vehicle. The Avco Mk 21 replaced 

the Mk 12 on the Air Force Peacekeeper (MX), and the Navy adopted a version 

called the Mk 5 for the Trident D-5. Typical of the latest warheads, the sharp-nose 

Mk 21 had a fine-weave, pierced-carbon-fabric nose tip followed by a graphiteepoxy body that was covered with an ablative carbon-phenolic heat shield.73

The CORONA Satellite Reconnaissance Program

Even as these activities were taking place, in early 1956 President Dwight D. 

Eisenhower authorized a new satellite reconnaissance project jointly managed 

by the Central Intelligence Agency and the U.S. Air Force. Its intention was 

to place in polar orbit, with all possible speed, a satellite carrying a camera that 

could take photographs of the Soviet Union (as well as other nations), return 

the film to Earth, and provide spatial and visual data to analysts on which to 

build more reliable intelligence estimates for decision makers. The long shadow 

of the successful surprise attack on Pearl Harbor in 1941 had prompted U.S. 

intelligence organizations to expand their reconnaissance efforts to avoid being 

caught unaware again.

Because of the successful effort during the 1950s to develop a reconnaissance satellite—Project CORONA had its first successful flight on August 

18, 1960, after several failures—a new era of intelligence gathering began. 

To disguise its true purpose, it was given the cover name “Discoverer” and 

described as a scientific research program. This highly classified reconnaissance 

 73.

31
Coming Home

effort—that acquired 3,000 feet of film with coverage of over 1,650,000 square 

miles of the former Soviet Union—revolutionized how the United States collected and used foreign intelligence. It ushered in an era in which the intelligence community regularly received imagery that offered a synoptic view of 

much of Earth’s surface. The intelligence community thereafter had both a 

high volume and continuous flow of data from satellite imagery. CORONA 

also was an important milestone from a historical perspective. It was the first 

imaging reconnaissance satellite, the first source of mapping imagery from 

space, the first source of stereo imagery from space, the first space program 

to succeed with multiple reentry vehicles, and the first space reconnaissance 

program to fly 100 missions.74

The CORONA program used a unique film-return capsule, essentially a version of the ICBM reentry vehicle, to reenter the atmosphere and return safely 

to Earth. An ablative heat shield, like those developed earlier for the ICBM 

program, dissipated the energy on the spacecraft as it entered the atmosphere 

and slowed down; then the heat shield was jettisoned and a parachute was 

deployed that enabled an Air Force plane to capture the capsule as it followed 

its trajectory into the Pacific Ocean. From August 1960 to May 1972, more 

than 120 successful CORONA missions provided invaluable intelligence on 

the Soviet Union and other nations.

CORONA was succeeded by a series of evermore sophisticated reconnaissance satellites, and a continuous stream of data began to flow from its 

imagery. However, it was the last system that used film imagery that had to 

be physically returned to Earth for processing and analysis. Overseen by the 

National Reconnaissance Office, the more than 50-year-old satellite reconnaissance effort has been enormously significant. Politicians, intelligence 

professionals, and the general public all view it as critical to the welfare of 

the United States and its continued national sovereignty. It enables the discovery of strategic weapons, military buildups, and troop movements, and 

it provides independent verification of strategic weapons-reduction efforts. 

President Lyndon Baines Johnson did not overestimate the importance of the 

CORONA satellite reconnaissance technology in 1967 when he said, in light 

 74.

32
Creating Ballistic Reentry

During the early years of satellite reconnaissance, spacecraft ejected film canisters that 

reentered the atmosphere and were retrieved by aircraft. This CORONA return canister is being 

recovered by a C-119 snatching its parachute during descent in 1962. USAF. 

of the fact that the United States probably spent between $35 and $40 billion 

(in 1960s dollars) on it, “If nothing else had come of it except the knowledge 

we’ve gained from space photography, it would be worth 10 times what the 

whole program has cost.”75 Accordingly, this mission will continue indefinitely, 

and the United States will invest in numerous future reconnaissance satellite 

programs throughout the 21st century.

At the same time, Allen and Egger’s concept of reentry would soon become 

important for a strikingly different use: NASA’s human space flight program 

initiated with the creation of the Agency on October 1, 1958. The application 

of reentry and recovery technology to the NASA human space flight program 

is the subject of the next chapter.

 75.

33
Walt Disney and Wernher von Braun teamed up in the 1950s to create the expectation that 

space flight was a near-term possibility with spaceplanes. NASA.

34
CHAPTER 2

The limitations of the heat-sink approach to returning to Earth, pioneered for 

warhead reentry, became even more apparent when the method was considered 

for a piloted spacecraft. First, a returning orbital spacecraft would enter at a 

faster velocity than a ballistic missile and would get much hotter, requiring 

more copper or beryllium. Perhaps even more importantly, flying a piloted 

capsule into orbit and back home, and recovering a fragile human pilot, is 

more difficult than was lobbing a warhead at Moscow.

However, there was an even greater problem. The heat sink itself got hot, 

and some of this heat—possibly several hundred degrees—transferred to the 

vehicle. This was a workable issue for a nuclear weapon but prohibitive both for 

the film-return canister used on the first reconnaissance program, CORONA, 

and on a piloted spacecraft.

Although almost all of the early development into reentry was oriented toward 

warheads, a small number of researchers believed the efforts were also applicable 

to future piloted spacecraft. In fact, the X-15 research airplane used a manifestation of the blunt-body concept during its suborbital reentry. True ballistic reentry, 

although acceptable for missile warheads, produced g-loads far too high (80 g’s 

or more) for human survival. However, by using the shallower reentry trajectory 

made possible by the blunt-body theory, the g-loads and peak temperatures could 

be significantly reduced, and the early ablative heat shields being developed for 

warheads provided a solution that made a small human space capsule possible. 

The g-loads (approximately 8 g’s for about 90 seconds) were still a concern, but 

centrifuge studies showed these were within the limits of human tolerance if the 

crew was in a reclining position and the force was applied from front to back 

(eyeballs-in). However, studies of the X-15 and Dyna-Soar cockpits produced 

primarily eyeballs-down g-loads. Both problems required serious efforts to understand reentry for human pilots. This led to piloted capsules designed so crews 

would lie on their backs, facing away from the direction of flight.

These concerns for reentry and recovery from space for a piloted spacecraft 

were far removed from the public discussion of space flight in the 1950s before 

35
Coming Home

Sputnik’s launch on October 4, 1957. For example, in 1951 when Willy Ley, 

a former member of the German VfR and himself a skilled promoter of space 

flight, organized a Space Travel Symposium at the Hayden Planetarium in New 

York City, the speakers hardly mentioned the problems of returning from space. 

Out of this symposium came the celebrated series of Collier’s articles appearing 

between 1952 and 1954 that did much to popularize space flight in President 

Dwight D. Eisenhower’s America. An editorial suggested that space flight was 

possible and was not just science fiction, and that it was inevitable that humanity would venture outward. It framed the exploration of space in the context 

of the Cold War rivalry with the Soviet Union and concluded the following:

Collier’s believes that the time has come for Washington to give 

priority of attention to the matter of space superiority. The rearmament gap between the East and West has been steadily closing. 

In addition, nothing, in our opinion, should be left undone that 

might guarantee the peace of the world. It’s as simple as that.1

There was nothing in any of this to suggest that returning from space was 

in any manner whatsoever difficult or would require extended experimentation. Wernher von Braun led off the Collier’s issue with an impressionistic 

article describing the overall features of an aggressive space flight program. He 

advocated the orbiting of an artificial satellite to learn more about space flight 

followed by the first orbital flights by humans; he also advocated developing a 

reusable spacecraft for travel to and from Earth’s orbit, building a permanently 

inhabited space station, and, finally, human exploration of the Moon and 

planets by spacecraft launched from the space station. Ley and several other 

writers followed with elaborations on various aspects of space flight ranging 

from technological viability to space law to biomedicine.2 The series concluded 

with a special issue of the magazine devoted to Mars in which von Braun and 

others described how to get there and predicted what might be found based 

on recent scientific data.3

 It was all going to be so simple and straightforward. 

Practical study and experimentation would soon prove otherwise.

 1. “What Are We Waiting For?” 

36
Human Space Flight and the Problem of Reentry

The X-Planes

Despite the obvious advances in rocket technology, before humans could venture into space, it would first be necessary to fly much faster than the 500 mph 

possible at the end of World War II. Since the beginning of powered flight, 

wind tunnels had proven to be useful tools, but they had several limitations 

regarding high-speed research. Signifying perhaps the greatest limitation, in the 

1930s it became apparent that the transonic regime could not be adequately 

simulated due to the physical characteristics of the wind tunnel test sections. 

The only alternative was to use real airplanes.

This led directly to what became known as the Round One X-planes, so 

called because the Air Force assigned the aircraft an X (for experimental) designation.4

 On October 14, 1947, Air Force Captain Charles E. Yeager became 

the first human to break the sound barrier in level flight when the Bell X-15

achieved Mach 1.06 at 43,000 feet. It took 6 additional years before NACA

test pilot A. Scott Crossfield exceeded Mach 2 in the Navy/Douglas D-558-2 

Skyrocket on November 20, 1953. Only 3 years later, on September 7, 1956, 

Air Force Captain Milburn G. Apt was killed during his first X-2 flight after 

he reached Mach 3.196 (1,701 mph), becoming the first person to fly at three 

times the speed of sound, albeit briefly.6

The original rationale behind the X-planes was to explore a flight regime 

that the wind tunnels could not simulate. However, by the time the X-1 and 

D-558 actually flew, researchers had figured out how to extend ground test 

facilities into this realm with the slotted-throat wind tunnel, developed by 

Richard Whitcomb at the NACA Langley Memorial Aeronautical Laboratory 

in the early 1950s. Therefore, the real value of the research airplanes lay in the 

comparison of the ground-based techniques with actual flight results to validate theories and wind tunnel results. The fact that the first transonic flights 

showed nothing particularly unexpected—dispelling the myth of the sound 

barrier—was of great relief to the researchers. The fearsome transonic zone had 

 4. Although several X-planes explored the high-speed/altitude flight environment, many others 

37
Coming Home

John V. Becker and the 11-Inch Hypersonic Tunnel at Langley Aeronautical Laboratory in 1950. 

NASA EL-2002-00243.

been reduced to an ordinary engineering problem.7 Although few people were 

thinking about it at the time, the results from these experiments would also 

be instrumental in developing the spaceplanes that had long been discussed.

The next step was to push through the so-called thermal barrier predicted 

by Theodore von Kármán and others. Although not related to specific velocity, 

like the speed of sound, vehicles venturing above Mach 5 (hypersonic velocities) experienced significantly increased heating rates that appeared to present 

a significant problem. Between the two World Wars, hypersonics had been an 

area of theoretical interest to a small group of researchers, but little progress was 

made in defining the possible problems and even less toward solving them. The 

major constraint was propulsion. Engines, even the rudimentary rockets then 

being experimented with, were incapable of propelling any significant object 

to hypersonic velocities. Wind tunnels also lacked the power to generate such 

speeds. Computer power to simulate the environment using models, what is 

now known as computational fluid dynamics, had not yet even been imagined.

Hypersonic research was authorized primarily to support the massive effort 

associated with developing intercontinental missiles.8 One researcher interested 

 7. John V. Becker, 

38
Human Space Flight and the Problem of Reentry

in exploring the new science of hypersonics was John V. Becker, then the 

assistant chief of the Compressibility Research Division at NACA Langley. 

On August 3, 1945, Becker proposed the construction of a new type of supersonic wind tunnel for Mach 7. A number of uncertainties caused Becker to 

suggest building a small pilot tunnel with an 11-by-11-inch test section to 

demonstrate the new concept. In September 1945, Charles H. McLellan and 

a small staff began work on the facility, and the first test, on November 26, 

1947, revealed uniform flow at Mach 6.9, essentially meeting all the original 

intents.9

 Nevertheless, many researchers wanted to build an actual hypersonic 

vehicle to validate the data from the new test facilities. The large rocket engines 

being developed for the missile programs were seen as possible powerplants 

for a hypersonic research vehicle. It was time for the Round Two X-planes.10

Round Two X-Planes

In late August 1954, researchers at NACA Langley released a 41⁄2-page paper 

entitled “NACA Views Concerning a New Research Airplane,” giving a brief 

background on the hypersonic research airplane and attaching a study by Becker 

as a possible solution. The researchers at Langley also proposed a secondary mission to explore space-related concerns, particularly reentry techniques, but most 

other researchers dismissed this because they believed human space flight was 

many decades off. Nevertheless, the paper listed two major areas to be investigated: 

(1) preventing the destruction of the aircraft structure by the direct or indirect 

effect of aerodynamic heating, and (2) achieving stability and control at very 

high altitudes, at very high speeds, and during atmospheric reentry from ballistic 

flightpaths. They were important goals on the way toward human space flight.11

An industry competition in 1955 resulted in the Air Force awarding North 

American Aviation a contract to build three experimental X-15 hypersonic 

research airplanes. The Government-industry team—led by Becker, Hartley 

A. Soulé, and Walter C. Williams from the NACA, and Crossfield, Charles H. 

 9. John V. Becker to the Langley Chief of Research, “Proposal for New Type of Supersonic Wind 

39
Coming Home

Feltz, and Harrison A. “Stormy” Storms, Jr., for North American—would soon 

become the stuff of legend.12 North American had accepted an extraordinarily 

difficult task when the company agreed to develop the new hypersonic research 

airplane. Eventually, some 2,000,000 engineering work-hours and over 4,000 

wind tunnel hours were devoted to finalizing the X-15 configuration.13

Researchers at Langley considered two basic structural design approaches. 

The first was a conventional low-temperature design of aluminum or stainless 

steel protected from the high-temperature hypersonic environment by a layer 

of assumed, but as-yet undeveloped, insulation. The other was an exposed 

hot structure in which the materials and design approach would permit high 

structural temperatures and would not require protection from the induced 

environment. It was similar, in many regards, to the heat-sink solution for 

missile warheads.14

Heating projections made by researchers for various trajectories showed 

that the airplane would need to accommodate an equilibrium temperature 

of 2,000 °F on its lower surface. At the time, no known insulating material 

could meet this requirement. The most likely candidate was the Bell doublewall technique in which a low-temperature structure was protected by a hightemperature outer shell with some insulation in between. This concept would 

later undergo extensive development in support of Dyna-Soar, but in 1954 it 

was in an embryonic stage and, in any case, was not applicable to the critical 

nose and leading-edge regions that would experience much of the heat load.15

It was by no means obvious that the hot-structure approach would prove practical either. The permissible design temperature for the best available materials was 

far below the peak 2,000 °F equilibrium temperature during reentry. It was clear 

that either direct internal cooling or heat absorption into the structure itself would 

be necessary, but each approach was expected to bring a heavy weight penalty.

 12.

40
Human Space Flight and the Problem of Reentry

Shock waves festoon a small-scale model of the X-15 in NASA Langley’s 4-by-4-foot 

Supersonic Pressure Tunnel. NASA L-1962-02577.

Exotic alloys made from the rare earth elements were still laboratory curiosities, so the list of candidate materials narrowed to corrosion-resistant steels 

(CRES) and magnesium-, titanium-, and nickel-base alloys. Although AM-350 

CRES and 6A1-4V titanium had good strength over a wide temperature range, 

the strength of both alloys tended to fall off rapidly above 800 °F. Fortunately, 

various magnesium and nickel alloys exhibited a gradual drop in strength up 

to 1,200 °F, and Inconel X16—for all intents an exotic stainless steel—was 

ultimately chosen for the outer skin of the airplane. In a happy coincidence, 

 16.

41
Coming Home

researchers at Langley discovered that the skin thickness needed to withstand 

the expected aerodynamic stresses was about the same needed to absorb the 

thermal load. This meant that it was possible to solve the structural problem 

for the transient conditions of the hypersonic research aircraft with no serious 

weight penalty for heat absorption.17

Other items developed for the X-15 included one of the first stable platforms (what is today called an inertial measurement unit), auxiliary power 

units, a ball nose to measure local flow direction in the high-temperature 

airstream (usually, incorrectly, called a Q-ball), reaction control rockets for 

attitude control in the space environment, and workable full-pressure suits. 

Various incarnations of all of these would find use on future spacecraft.18

Perhaps even more important was the development of extensive engineering 

and mission-simulation systems. Although crude by current standards, the 

X-15 pioneered the use of simulators not just to train pilots but also to engineer the aircraft, plan the missions, and understand the results. Not surprisingly, given the involvement of Feltz, Storms, and Williams in the subsequent 

Apollo and Space Shuttle programs, the X-15 pointed the way to how America 

would conduct its human space missions. Simulation is one of the enduring 

legacies of the small, black experimental airplanes.

The X-15 flight program encompassed 199 flights between June 1959 and 

October 1968. By July 1962, the X-15 had flown more than Mach 6 and 

above 300,000 feet, exceeding its design goals. NASA test pilot Joseph A. 

Walker would fly the maximum-altitude flight of the program, 354,200 feet, 

on August 22, 1963. This was his second excursion above 62 miles in just over 

a month, making him the first person to venture into space (albeit suborbital 

space) twice. Air Force Major William J. “Pete” Knight flew the maximumspeed mission, 4,520 mph (Mach 6.70) on October 3, 1967.19

 17.

42
Human Space Flight and the Problem of Reentry

Neil Armstrong is seen next to X-15-1 (56-6670) after a research flight. Armstrong made his 

first X-15 flight on November 30, 1960. This was the first X-15 flight to use the ball nose, which 

provided accurate measurement of flow angle (attitude) at supersonic and hypersonic speeds. 

NASA E60-6286.

The X-15s amassed a large body of hypersonic flight data that researchers 

are still using more than 50 years later. One of the largest contributions from 

the program was in conclusively demonstrating that there were no hypersonic facilities barriers. Becker remembered that many aerodynamicists had 

expected that because of strongly interacting flow fields, viscous interactions 

with strong shock waves, and possible real-gas effects, the X-15 would reveal 

large discrepancies between wind tunnel data and actual flight. Instead, the 

X-15 demonstrated that most flight results agreed substantially with lowtemperature, perfect-gas wind tunnel predictions. Data from the X-15, coupled with data from the ASSET and PRIME programs, allowed the design 

of future hypersonic vehicles, including the Space Shuttle, to proceed with 

much greater confidence.20

 20.

43
Coming Home

When Project Mercury began, it rapidly dominated some of the research 

areas that had first interested X-15 planners, such as microgravity studies and 

reaction-control systems pioneered in this program and incorporated wholesale into the Mercury capsule. These reaction controls effectively maintained 

attitude control in space for both the X-15 and the Mercury capsule, and, 

beyond that, the X-15 furnished valuable information on the blending of 

reaction controls with conventional aerodynamic surfaces during exit and 

reentry—a matter of concern to subsequent Space Shuttle development. The 

X-15 clearly demonstrated the ability of pilots to fly rocket-powered aircraft 

out of the atmosphere and back and to make precision landings. The director 

of the Flight Research Center, Paul F. Bikle, saw the X-15 and Mercury as a 

“parallel, two-pronged approach to solving some of the problems of piloted 

space flight. While Mercury was demonstrating humanity’s capability to function effectively in space, the X-15 was demonstrating man’s ability to control 

a high-performance vehicle in a near-space environment...considerable new 

knowledge was obtained on the techniques and problems associated with lifting 

reentry.” They were important steps toward an eventual spaceplane.21

Faget and Nonlifting Reentry

Throughout most of 1957, research into human space flight continued to 

examine winged vehicles. In particular, the NACA cooperated with the Air 

Force on the Manned Glide Rocket Research System, and the Ames laboratory 

looked into a flattop, round-bottom skip-glider. Nevertheless, after Sputnik was 

launched on October 4, 1957, it became obvious that a ballistic capsule was 

both the best and quickest way to get Americans into orbit; the available launch 

vehicles simply could not support the increased weight of a winged vehicle. Even 

so, what would later become Dyna-Soar was being discussed, culminating with 

the Round Three X-plane conference at Ames on October 15, 1958.22

A minority contingent within the NACA, mainly at Langley, continued to 

argue that lifting reentry vehicles would be far superior to a nonlifting capsule. 

In fact, at the NACA Conference on High-Speed Aerodynamics in March 1958, 

Becker presented a concept for a piloted 3,060-pound winged orbital satellite. 

According to Becker, this concept drew more industry reaction—almost all of 

 21.

44
Human Space Flight and the Problem of Reentry

it favorable—than any other paper he had written, including the initial X-15 

study. What ruled out acceptance of his proposal was that the 1,000 pounds of 

extra weight (compared to the capsule design presented by Maxime A. Faget) was 

beyond the capability of the Atlas launch vehicle. If the Titan had been further 

along, the concept would have worked, and Becker believed that the first U.S. 

piloted spacecraft might well have been a landable winged vehicle.23

Although wings brought many benefits during entry and landing, they also 

brought structural weight, and there was no practical, and little theoretical, data 

applicable to the development of a spaceplane. On the other hand, Faget at 

NACA Langley believed the first piloted spacecraft should be a simple nonlifting 

capsule that could take full advantage of the research and production experience 

gained through development of warheads for the ballistic-missile programs. Faget 

thought that much of the aerodynamic and thermodynamic work done by the 

NACA on ballistic-missile warheads was directly applicable to the piloted ballistic capsule. Given a maximum weight of about 2,000 pounds, a vehicle with 

a base diameter of 7 feet would provide sufficient volume for a single occupant 

and the required equipment.24

Faget opined that the capsule concept had certain attractive operational 

aspects including following a ballistic path that minimized requirements for the 

autopilot, guidance, and control systems. In his mind, this not only resulted 

in a significant weight savings but also eliminated the hazard of malfunction. 

The ballistic capsule only had to perform a single maneuver to orient itself 

properly and then fire its retrorockets. Faget pointed out that this did not need 

to be done with a great deal of precision for the vehicle to successfully enter the 

atmosphere, but he did acknowledge the disadvantage of a large-area landing 

by parachute with no corrective control during entry.25

There was, of course, another disadvantage. At an entry angle of –0.8 degree, 

the vehicle would endure approximately 8.5 g’s on the way down (although 

the buildup was expected to be relatively gradual). Faget believed this was well 

within the tolerances of the test pilots expected to inhabit the vehicle, and he 

noted that larger entry angles with even more severe deceleration should be 

tolerable with the proper training and equipment. He readily accepted this as 

a program requirement.26

 23.

45
Coming Home

Researchers investigated four different shapes for the vehicle: a hemisphere, 

a heavily blunted 15-degree cone, a 53-degree cone, and a nearly flat nose with 

a spherical segment. Perhaps the greatest unknown was the effect of laminar 

flow on the vehicle. Limited testing on rocket-boosted, free-flying models at 

Wallops Island confirmed that laminar flow could be expected on at least the 

flat-nose shape. If laminar flow was truly present, the total heat load could be 

reduced by an order of magnitude over the ballistic-missile warheads.27 Faget 

concluded his study by saying: “It appears that, insofar as reentry and recovery 

is concerned, the state of the art is sufficiently advanced so that it is possible 

to proceed confidently with a piloted satellite project based upon the ballistic 

reentry type of vehicle.”28

Even before the first Mercury flight, however, it was recognized that a purely 

ballistic vehicle was not the ideal approach. In addition to g-loads and heating 

during reentry, the poor predictability of the final impact point forced the use 

of a large contingent of ships spread over a wide area of ocean for recovery. 

Ultimately, Mercury was the only U.S. piloted spacecraft to use a purely ballistic reentry without some ability to steer the vehicle.29

Providing a small amount of lift—an aerodynamic force perpendicular to 

the flightpath—reduced the severity of reentry and improved the accuracy 

of recovery. For the capsules, lift could be created by slightly offsetting the 

center of gravity from the vertical axis. The trajectory could then be changed 

by rolling the vehicle about its vertical axis so that the offset lift vector could 

be pointed in any direction. The roll maneuver could be accomplished via a 

reaction wheel (gyroscope) or by using small reaction-control thrusters, which 

is the approach that ultimately was selected. The lift could also be directed 

upward to maintain a small flightpath angle for as long as possible. Once the 

vehicle passed the high-speed heating pulse, it could be banked to turn toward 

the desired recovery point, producing a relatively small cross-range capability.30

A lift-to-drag ratio (L/D) of only 0.2 could cut the maximum deceleration in 

half and stretch the range by 280 miles.31

 27.

46
Human Space Flight and the Problem of Reentry

A hot jet research facility, used extensively in the design and development of the reentry heat 

shield on the Project Mercury spacecraft. The electrically heated arc jet simulates the heating 

encountered by a space vehicle as it returns to Earth’s atmosphere at high velocities. NASA 

EL-2002-00308.

The resulting semiballistic reentries were of longer duration than a pure 

ballistic reentry, but they produced more tolerable g-loads and lower peak 

temperatures. A more complex design was needed for the ablative heat shield 

because the heating was not symmetrical and the duration of the heat pulse 

was longer, driving the development of a nonreceding, charring ablator that 

stayed in place to maintain the outer mold line and aerodynamic characteristics of the vehicle. Gemini and Apollo used semiballistic capsules protected 

by nonreceding, charring ablators and were initially intended to be recovered 

on land. Ultimately, however, both vehicles still required large ocean areas and 

significant recovery forces at the end of their missions.32

Although not a design factor during the early space race, all of these ballistic 

and semiballistic vehicles were designed to survive only a single reentry. The 

only capsule known to have flown twice was Gemini II, which after its initial 

test flight was modified as a prototype Gemini-B for the Manned Orbiting 

Laboratory (MOL). An entirely new heat shield was used for the second flight 

on a developmental Titan IIIC, on November 3, 1966.

 32.

47
Coming Home

Man in Space Soonest

While the NACA was pursuing its studies for a human space flight program, 

the U.S. Air Force proposed the development of a piloted orbital spacecraft 

under the title of “Man In Space Soonest” (MISS).33 Initially discussed before 

the launch of Sputnik 1, in October 1957, the Air Force invited Dr. Edward 

Teller and several other leading members of the scientific/technological elite 

to study the issues of human space flight and make recommendations for the 

future. Teller’s group concluded that the Air Force could place a human in 

orbit within 2 years and urged the department to pursue reasons to undertake 

this effort. Teller understood, however, that there was essentially no military 

reason for the mission and chose not to tie his recommendation to any specific 

rationale, falling back on a basic belief that the first nation to put a man in space 

would gain national prestige and generally advance in science and technology.34

Soon after the new year, Lieutenant General Donald L. Putt, the Air Force 

Deputy Chief of Staff for Development, informed NACA Director Hugh L. 

Dryden of the Air Force’s intention to pursue aggressively “a research vehicle 

program having as its objective the earliest possible manned orbital flight which 

will contribute substantially and essentially to follow-on scientific and military 

space systems.” Putt asked Dryden to collaborate in this effort, with the NACA 

taking on the role of a decidedly junior partner.35 Dryden agreed, but by the 

end of the summer the newly created NASA was leading the U.S. human space 

flight effort, and the Air Force was the junior player.36

Notwithstanding the lack of a clear-cut military purpose, the Air Force 

pressed for MISS throughout the first part of 1958, clearly expecting to become 

the lead agency in any U.S. space program. Specifically, it believed hypersonic 

spaceplanes and lunar bases would serve well the country’s national security 

 33.

48
Human Space Flight and the Problem of Reentry

needs in the coming decades. To help make that a reality, the Air Force requested 

$133 million for the MISS program and secured approval for the effort from 

the Joint Chiefs of Staff.37 Throughout this period, a series of disagreements 

between Air Force and NACA officials aggravated both sides. The difficulties 

reverberated all the way to the White House, prompting a review of the roles 

of the two organizations.38 Dryden, the normally staid and proper director of 

the NACA, complained in July 1958 to the President’s science advisor, James 

R. Killian, about the lack of clarity regarding the role of the Air Force versus 

the NACA. He asserted, “The current objective for a manned satellite program 

is the determination of man’s basic capability in a space environment as a prelude to the human exploration of space and to possible military applications of 

manned satellites. Although it is clear that both the National Aeronautics and 

Space Administration and the Department of Defense should cooperate in the 

conduct of the program, I feel that the responsibility for and the direction of 

the program should rest with NASA.” He urged that the President state a clear 

division between the two organizations on the human space flight mission.39

As historians David N. Spires and Rick W. Sturdevant pointed out, the 

MISS program became derailed within the DOD at essentially the same time 

because of funding concerns and the lack of a clear military mission:

Throughout the spring and summer of 1958 the Air Force’s Air 

Research and Development Command had mounted an aggressive campaign to have ARPA convince administration officials to 

 37.

49
Coming Home

approve its Man-in-Space-Soonest development plan. But ARPA 

[Advanced Research Projects Agency] balked at the high cost, 

technical challenges, and uncertainties surrounding the future 

direction of the civilian space agency.40

Project Mercury

President Eisenhower signed the National Aeronautics and Space Act of 1958 

into law at the end of July and the next month assigned the nation’s human 

space flight mission to NASA. Thereafter, the MISS program was folded into 

what became Project Mercury.41 By early November 1958, DOD had acceded 

to the President’s desire that the human space flight program be a civilian 

effort under the management of NASA. For its part, NASA invited Air Force 

officials to appoint liaison personnel to the Mercury program office at Langley, 

and they did.42

Unlike the spaceplane concept pursued by the Air Force during the early 

planning for MISS, the Mercury spacecraft configuration was primarily determined by heating considerations, and the final shape evolved from an extensive study of how to minimize the effects of reentry heating while providing 

adequate stability.43 The first configuration, designed mostly by Faget, used a 

conical body with a flat heat shield—essentially a beefed-up, piloted Mk II 

 40.

50
Human Space Flight and the Problem of Reentry

NASA undertook tests of full-size Mercury capsules (either launched from the ground by rocket 

power or dropped from airplanes at high altitude) to test the capsule’s dynamic stability and 

aerodynamic heating as well as the effectiveness of the pilot-escape and parachute-recovery 

systems. This test took place in 1958. NASA EL-2000-00282.

warhead. The afterbody was recessed slightly from the perimeter of the heat 

shield to minimize heat transfer to the inhabited part of the spacecraft. Tests at 

Langley showed this configuration was unstable at subsonic speeds and suffered 

various heating problems during reentry, leading to the afterbody becoming 

more squat and extending all the way to the perimeter of the heat shield. The 

flat heat shield was soon discarded because it trapped too much heat, leading 

to a rounded bottom, and NASA researchers decided on a 1.5 ratio between 

the radius of the curve and the diameter of the heat shield. The final obstacle 

was figuring out where to locate the landing parachutes. Eventually, engineers 

decided to add a cylinder on top of the afterbody (the frustum), creating the 

now familiar Mercury shape.44

Mercury is the only U.S. piloted spacecraft that used a purely ballistic 

entry, with conditions dictated primarily by the g-limits for the pilot. The 

 44.

51
Coming Home

anticipated environment included a heat pulse of approximately 8,910 BTU 

per square foot for a spacecraft weighing 2,439 pounds. Temperatures in the 

shock layer immediately ahead of the blunt-body heat shield were expected to 

be in excess of 10,000 °F. To provide adequate protection for the astronaut, 

whose back was only a few inches away from the inner surface of the heat 

shield, engineers set the maximum permissible temperature on the back of 

the heat shield at 150 °F. In addition to handling the expected head load, 

the heat shield needed to withstand the dynamic pressures expected during 

reentry as well as relatively high water-impact forces while the heat shield was 

still hot. The heat shield also needed to withstand the acoustic and vibration 

environments imposed during launch and ascent as well as the hard vacuum 

and low temperatures of the space environment.45

The Air Force, along with Avco Corporation and General Electric, recommended using an ablative heat shield, based on the later Mk II reentry vehicle. 

NASA researchers, however, worried that the relatively recent ablative technology was too immature, and they wanted to stay with a heat sink similar to the 

initial General Electric Mk II design.46 A slightly refined reentry profile reduced 

the total heat load to approximately 6,000 BTU per square foot, and this could 

be accommodated by a 1-inch-thick slab of beryllium with a unit weight of 

10 pounds per square foot. This forged heat sink was fabricated by sintering 

beryllium powder, a process generally similar to the one used to manufacture 

the early Mk II reentry vehicles.

Ultimately, NASA researchers concluded that the toxicity and the fire hazard 

posed by the beryllium heat sink in the event of a nonwater landing was unacceptable. However, by the time this decision was made, a number of heat sinks 

had been manufactured and, mostly to provide an early flight capability, were 

used on the suborbital flights of Alan Shepard and Virgil I. “Gus” Grissom. 

Once NASA had decided that beryllium posed unacceptable risks, they again 

turned to the ballistic missile researchers for an answer.47

Ablation technology was being rapidly advanced in the quest for a medium 

ballistic-coefficient reentry vehicle for the long-range Atlas and Titan ICBMs. 

In fact, an ablative-covered Jupiter reentry vehicle was successfully recovered in 

the fall of 1958, just as the Mercury spacecraft was entering preliminary design. 

The Jupiter reentry vehicle used a composite design that provided structural 

strength while allowing large volumes of hot gasses to flow between the ablative 

 45.

52
Human Space Flight and the Problem of Reentry

Structural details of the Mercury heat shield. NASA.

53
Coming Home

laminates without forcing the laminates to separate. The design appeared ideal 

for Mercury, and NASA quickly adopted it to replace the beryllium heat sink.48

Unfortunately, the Jupiter reentry vehicle used a trajectory more closely 

resembling the suborbital Mercury flights instead of the anticipated orbital missions. No directly applicable reentry experience existed, so NASA researchers 

wanted to conduct a flight qualification test to demonstrate that the ablative 

heat shield could withstand an orbital entry. The mechanics of orbital flight 

were still being developed, so researchers decided a ballistic (suborbital) mission would pose less risk and would better ensure that the heat shield could 

be recovered for postflight analysis.49 A carefully chosen flight profile would 

largely duplicate the heating rates and total heat load of an orbital entry. NASA 

tested this with the “Big Joe” flight program consisting of two flights, each 

using an Atlas launch vehicle that would loft a boilerplate Mercury capsule 

1,500 miles downrange.50

Although the outer mold line was identical, the structure of the boilerplates 

for the Big Joe tests was not typical of production Mercury capsules. For 

instance, the boilerplate only contained a half-size pressure vessel to support 

instrumentation rather than the full-size pressurized cabin contoured to the 

outer mold line. The 2,555-pound boilerplate was built in two segments, with 

the lower blunt body manufactured at the Lewis Research Center and the upper 

afterbody manufactured at Langley, both under the direction of the Space Task 

Group. The main monocoque of the afterbody was fabricated using thin sheets 

of corrugated Inconel. General Electric supplied the outer heat shield, with 

parts of the interior structure manufactured by B.F. Goodrich. The heat shield 

was instrumented with 51 thermocouples to obtain temperature and charpenetration time histories during flight. The boilerplate did not carry a retropackage or an escape tower, but it did contain a parachute recovery system.51

The Big Joe heat shields consisted of two parts: an outer ablation laminate 

that was 1.075 inches thick and an inner structural laminate that was 0.550 inch 

thick. The ablation laminate was made of concentric layers of fiberglass cloth oriented so that the layers were at a 20-degree angle to the local heat shield surface. 

The structural laminate was made of similar fiberglass cloth oriented with the 

individual layers parallel to the outer surface. Both the ablation and structural 

 48.

54
Human Space Flight and the Problem of Reentry

Engineers inspect a boilerplate Mercury space capsule with an array of bulky test equipment 

strewn about. NASA GPN-2000-003008.

layers were made from a special finish fiberglass cloth and a phenolic resin. The 

resin content of the ablation and structural laminates was, by weight, 40 percent 

and 30 percent, respectively. A 3-inch-high circular ring made from fiberglass 

and resin was attached to the back of the heat shield and was used to bolt the 

heat shield to the pressurized compartment of the boilerplate.52

The first test flight, Big Joe 1, was launched on Atlas 10-D (628) from 

the Air Force Missile Test Center Launch Complex 14 at Cape Canaveral on 

September 9, 1959. The Atlas was programmed to rise, pitch over horizontally 

as it reached its 100-mile peak altitude, and then pitch down slightly before 

releasing the boilerplate at a shallow angle downward. The Atlas had three 

engines: a sustainer engine in the middle and two booster engines strapped to 

the sides of the main missile. At a predetermined time during ascent, the two 

booster engines were designed to separate from the missile, as their thrust no 

longer compensated for their additional weight and drag. A malfunction on 

Atlas 10-D kept this from happening, and the additional weight caused the 

 52.

55
Coming Home

sustainer engine to deplete its propellants 14 seconds early and delay the separation of the boilerplate from the launch vehicle. Nevertheless, the 13-minute 

flight reached an altitude of 90 miles and traveled 1,424 miles downrange. The 

boilerplate reached a maximum velocity of 14,857 mph, a maximum dynamic 

pressure (max-q) of 675 pounds per square foot, and sustained a maximum 

of 12 g’s. The parachutes operated successfully, and the USS Strong (DD-758) 

recovered the boilerplate in excellent condition off Puerto Rico.53

Because of the launch vehicle anomaly, the peak-heating rate obtained 

during the flight was only 77 percent of that expected, and the total cold-wall 

heat load was only 42 percent of the expected 7,100 BTU per square foot.54

Despite not meeting the initial objectives, the flight was generally successful. 

The recovered boilerplate showed the heat shield withstood both the reentry 

and recovery operations with only superficial damage. The heat shield exhibited generally uniform heating, and the visibly charred region penetrated to a 

depth of approximately 0.20 inch, or some 12 percent of the total thickness. 

Small cracks and areas of delamination were found, but these did not extend 

in depth beyond the visibly charred portion and did not affect the structural 

integrity of the heat shield. Analysis showed no measurable profile change, 

and the heat shield lost only 6.1 pounds of mass due to the ablation process.55

The results from the flight test confirmed that the theories (models) being 

used by the researchers and engineers were valid and would satisfactorily predict 

heating for other trajectories. This gave NASA researchers enough confidence 

to cancel the Big Joe 2 test (Atlas 20-D), which had been scheduled for the fall 

of 1959, and the launch vehicle was transferred to another program. The Big 

Joe 1 boilerplate is currently displayed at the Smithsonian National Air and 

Space Museum’s Steven F. Udvar-Hazy Center at Chantilly, VA.

In addition to the Big Joe and Wallops Island tests, materials for the Mercury 

heat shield were evaluated in several NASA test facilities. Models ranging from 

full scale to 1 percent of full scale were used. Researchers conducted the tests 

in plasma-arc, radiant-lamp, and oxyhydrogen blowtorch facilities. In general, 

these tests yielded poor results because of contamination in the atmosphere 

(usually helium) and poor control over the heating rates. Fortunately, the 

Langley Structures Division researchers devised a way to test the heat shield 

 53.

56
Human Space Flight and the Problem of Reentry

An Atlas launch vehicle carrying the Big Joe capsule leaves its launch pad on a 2,000-mile 

ballistic flight to an altitude of 100 miles on September 1, 1959. This boilerplate model of the 

orbital capsule was recovered and studied for the effect of reentry heat and other flight stresses. 

NASA MSFC-9139360.

57
Coming Home

in their new arc-jet facility using air as a working medium, eliminating the 

contamination issues associated with the earlier tests. This facility could not 

accommodate a full-size heat shield, so samples were cut from a production 

heat shield (no. 13) that had been rejected due to defects discovered during 

ultrasonic inspection. These tests further characterized the behavior of the 

ablative heat shield and confirmed the adequacy of the theories and models 

used to develop it.56

Most of the afterbody of the Mercury capsule was protected by a series of 

René 41 shingles with fibrous batt insulation underneath. The beaded (corrugated) shingles were 0.016 inch thick and were attached by bolts through 

oversized holes that allowed the shingles to expand and contract without 

buckling. Oversized washers covered the holes to minimize heat and air penetration. René 41 consists of 53 percent nickel, 19 percent chromium, 11 

percent cobalt, 9.75 percent molybdenum, 3.15 percent titanium, 1.6 percent aluminum, 0.09 percent carbon, 0.005 percent boron, and less than 

2.75 percent iron. Other parts of the afterbody were covered with beryllium 

shingles fabricated from hot-pressed beryllium blocks. Thermoflex RF insulation blankets were located between these shingles and the primary structure. 

Both the René 41 and beryllium shingles were coated on the outer surface 

with blue-black ceramic paint to enhance their radiative properties. The inner 

surface of the beryllium shingles had a very thin gold coating to attenuate 

thermal radiation into the spacecraft.57

In October 1959, Mercury prime contractor McDonnell received the first 

production Mercury ablative heat shield from General Electric. This unit was 

used on the first Mercury capsule, which was delivered to Wallops Island on 

April 1, 1960, for a beach-abort test that took place on May 9.58 Surprisingly, 

the production heat shields were slightly thinner than the Big Joe unit, with 

an ablation laminate only 0.65 inch thick and a structural laminate 0.30 inch 

thick. Several problems were encountered during the fabrication of the production heat shields, most concerning whether the desired inclination of the 

laminates was being maintained. Ultimately, bore samples showed that it was 

not possible to keep the inclination correct near the center of the heat shield, 

raising a concern that this area would delaminate during reentry. Engineers 

from Langley, General Electric, and McDonnell concluded that the best course 

of action was to machine a 15-inch diameter section out of the center of each 

 56.

58
Human Space Flight and the Problem of Reentry

Recovery practice for the Mercury program after landing in the ocean conducted by the USS 

Strong (DD-758) in 1959. NASA.

heat shield and replace it with a molded plug that was layed-up separately 

(and correctly, since the fabric did not need to stretch the entire diameter of 

the heat shield). The plug was secured in the heat shield by 12 inclined dowel 

pins and a layer of glue.59

The background provided by the single Big Joe flight and the arc-jet tests 

gave NASA researchers confidence that the operational heat shield would perform satisfactorily. Since the suborbital Redstone launches had used capsules 

equipped with beryllium heat sinks, the first real use of an ablative heat shield 

was on the Mercury-Atlas 2 (MA-2) test flight on February 21, 1961. The flight 

was designed to subject the spacecraft (no. 6) to maximum g-loads (15.9 g’s) 

and to produce the maximum afterbody heating; consequently, the total heat 

load on the heat shield was less than the load during a nominal orbital entry. 

MA-2 flew a successful suborbital mission that lasted 17 minutes, 56 seconds 

and reached an altitude of 114 miles and a speed of 13,227 mph. The capsule 

 59.

59
Coming Home

was recovered 1,432 miles downrange, and subsequent analysis showed the 

heat shield performed satisfactorily.60

The operational experience with the Mercury heat shield was generally satisfactory, but the center plug section continued to cause problems. During the 

first orbital flight of the Mercury heat shield on the MA-4, on September 13, 

1961, the center plug section cracked free at the outer diameter and the resulting gap was about 0.1875 inch, although the inclined dowel pins still tightly 

retained the plug.61 The flight of Enos the Chimp in MA-5 on November 20, 

1961, ended with the center plug completely missing from the heat shield. In 

this case, the holes used by the dowels had been drilled too shallow, allowing 

the pins to slip free, and air bubbles were discovered in the glue. A transducer 

under the center plug showed that the plug had been in place up to the instant 

of water impact, and there was no damage to Enos or the capsule.62

McDonnell and General Electric subsequently developed improved inspection techniques, and all of the remaining heat shields were inspected and 

repaired as necessary to ensure the center plug did not separate on future missions. Despite the precautions, the center plug was again missing after the first 

U.S. piloted orbital mission of John Glenn in MA-6 on February 20, 1962, 

and after the MA-7 flight of Scott Carpenter on May 24, 1962.63

When Wally Schirra was launched in MA-8 on October 3, 1962, the 

materials and construction of the heat shield were the same as the previous 

orbital missions with the exception that the center plug was bolted to the 

structural laminate to prevent its loss after entry. During the postflight 

examination, engineers found the center plug was still firmly attached. The 

heat shield provided excellent protection during entry, and the stagnation point 

appeared to have been very close to the center of the heat shield, as expected 

of a ballistic (nonlifting) entry. The most worrisome finding was evidence of 

separation where the ablation laminate was glued to the structural laminate. 

Based on the depth of the char around the cracks, the separation appeared to 

have developed after the peak heating, perhaps as late as water impact (the 

ablator was still quite hot when it hit the cold ocean). The char depth measured 

0.33 to 0.44 inch, about the same as previous missions. The measured weight 

loss, however, was 17.43 pounds—substantially more than the 13 pounds on 

previous missions. Engineers noted that the technique used for drying the 

 60.

60
Human Space Flight and the Problem of Reentry

Inside Hangar S at Cape Canaveral, Mercury astronaut M. Scott Carpenter examines the honeycomb protective material on the main pressure bulkhead (heat shield) of his Mercury capsule, 

nicknamed “Aurora 7.” NASA S62-01420.

61
Coming Home

heat shield prior to weighing it changed for MA-8, so the results might not be 

directly comparable.64

The last Mercury flight was piloted by Gordon Cooper in MA-9 on May 14 

to 15, 1963. Concerns about the delamination on MA-8 led NASA to install 

14 bolts through the ablative laminate into the structural laminate to keep them 

from separating if the glue failed completely. This was a fortunate decision as 

postflight inspection revealed that the glue had separated, but the heat shield 

stayed together thanks to the bolts. Using the same drying procedure as on MA-8, 

engineers noted the heat shield lost 15.34 pounds.65

Project Gemini

Announced on December 7, 1961, by Robert R. Gilruth, Gemini was the 

third civilian U.S. human space flight program to be approved, after both 

Mercury and Apollo.66 Designed to fill the gap between Mercury and Apollo, 

Gemini was originally called Mark II, referring to its early status as a larger 

and slightly improved Mercury capsule. In the end, it would prove to be much 

more, although it bore a family resemblance to Mercury. The most significant 

of the proposed changes, which was never implemented, was the Rogallo wing 

paraglider recovery system, which is discussed in chapter 3.

One of the objectives for Gemini was to demonstrate a controlled reentry 

to a preselected landing site. Although the Gemini spacecraft had the same 

general shape as the ballistic Mercury capsule it was derived from, including 

the blunt heat shield and rotational symmetry about its longitudinal axis, 

it had an inherent lifting capability produced by a vertical center-of-gravity 

displacement with respect to the longitudinal axis. This center-of-gravity displacement introduced a trim angle of attack that, because of the blunt heat 

shield, produced a lift vector in the opposite direction from the trim angle. 

The vehicle could be rolled about the relative velocity vector or stability axis to 

provide a small maneuvering capability during reentry. If the vector modulation was done precisely enough, the spacecraft could be guided to a preselected 

 64.

62
Human Space Flight and the Problem of Reentry

These four shadowgraph images represent early reentry vehicle concepts. A shadowgraph is a 

process that makes visible the disturbances that occur in a fluid flow at high velocity. Light passing through a flowing fluid is refracted by the density gradients in the fluid, resulting in bright 

and dark areas on a screen placed behind the fluid. A blunt body, as pioneered by H. Julian 

Allen, produces a shock wave in front of the vehicle—visible in the photo—that actually shields 

the vehicle from excessive heating. As a result, blunt-body vehicles can stay cooler than pointy, 

low-drag vehicles can. NASA 60-108.

landing location, such as the dry lakes at Edwards Air Force Base (AFB), to 

use the Rogallo wing.67

The preselected landing location had to lie within the downrange and crossrange footprint of the vehicle—since it was unpowered, the landing site had to 

be within the range of the vehicle’s available energy—which was approximately 

300 miles downrange and 50 miles across laterally.68 An onboard digital computer was responsible for selecting a lift orientation that guided the spacecraft 

 67.

63
Coming Home

to the landing location. Alternately, the astronauts could take manual control 

using the hand controller in the cabin. Offsetting the center of gravity approximately 1.75 inches from the longitudinal centerline resulted in an L/D of 

approximately 0.22.69

After initial development problems on the early Gemini flights, the guidance and control system worked well in both manual and automatic modes. 

On one occasion, the landing was made within a few hundred yards of the 

targeted point. The first use of a blunt-body lifting reentry vehicle verified that 

the concept would work for Apollo.70

The operational Gemini heat shield consisted of a tapered layer of McDonnell 

S-3 silicon elastomeric ablator 1 inch thick at the windward edge and 0.85 

inch thick at the leeward edge (the heat shields for Gemini I and II were about 

half these thicknesses). The back structure was a 0.75-inch-thick section of 

phenolic fiberglass honeycomb consisting of two five-ply faceplates of resinimpregnated glass cloth separated by a 0.65-inch-thick fiberglass honeycomb 

core. The space between the heat shield and the back structure was filled with 

0.375 inch of RF-300 batt insulation. The maximum diameter of the heat 

shield was 90 inches, 15.5 inches greater than Mercury’s heat shield, and the 

spherical radius was 144 inches (resulting in a 1.6 ratio, instead of the 1.5 used 

on Mercury). The heat shield was attached to the cabin section by eighteen 

0.25-inch-diameter bolts. The first production heat shield was completed on 

February 3, 1964.71

Interestingly, the S-3 ablator, developed under NASA contract, proved so 

effective that Dow Corning purchased the rights to the product and sold it as 

DC-325. The ablator formed an excellent char layer during ablation, was stable 

in a vacuum, and could withstand the wide variety of temperatures in the space 

environment. Perhaps the most important change was that McDonnell had 

figured out how to lay-up the cloth across the entire heat shield, eliminating 

the center plug used on Mercury. An MX-2625 Fiberite ring was used around 

the outer edge of the heat shield where extra strength was required to absorb 

loads transferred from the launch vehicle. This was a significant improvement 

 69.

64
Human Space Flight and the Problem of Reentry

over the phenolic resin-impregnated laminated fiberglass unit used on Mercury. 

The Gemini heat shield weighed 347 pounds, compared to the Mercury heat 

shield’s 303 pounds, mostly because of a 48-percent increase in area, a 25-percent increase in the ballistic loading coefficient, and a 90-percent increase in 

the design heat load (per square foot) due to the lifting reentry.72

The afterbody heat protection used on Gemini was almost identical to that 

used on Mercury. Beaded (corrugated) 0.016-inch-thick René 41 shingles were 

used over the conical section, and the outer surfaces of the shingles were coated 

with a black oxide film to enhance the thermal balance and protect the René 

41 from oxidation. These shingles could withstand up to 1,800 °F by achieving a thermal balance through radiation to the atmosphere. The shingles were 

attached to the underlying structure using bolts through oversized holes that 

accommodated thermal expansion. Small blocks of Min-K insulation were used 

at the boltholes, and a layer of Thermoflex batt insulation was used between the 

shingles and the underlying structure. The cylindrical sections of the afterbody 

experienced temperatures too high for effective radiative cooling and used eight 

beryllium shingles as heat sinks. These shingles had a thin gold coating on the 

inside to minimize thermal radiation into the spacecraft and were 0.24 inch 

thick on the windward side and 0.09 inch thick on the leeward side.73

The first Gemini launch to use a heat shield was the Gemini 2 mission, 

which contained production units of all the equipment used on the later 

piloted missions except for the rendezvous radar, drogue parachute system, and 

auxiliary tape memory. The spacecraft was controlled through an automated 

sequencing device. Gemini 2 had been scheduled for launch on December 

9, 1964, but as the booster engines were ignited, the Malfunction Detection 

System detected a loss of hydraulic pressure and shut down the engines 1 

second later. After correcting the problems and checking the vehicle, Gemini 

2 was launched on January 19, 1965. The capsule splashed down 2,125 miles 

downrange and was recovered by the USS Lake Champlain (CVS-39). During 

reentry, temperatures near the adapter interconnect fairing on the afterbody 

were higher than expected and damaged two René 41 shingles. Subsequent 

Gemini capsules were built with slightly heavier gauge shingles, and the trim 

angle of attack was lowered to reduce local heating.74

 72.

65
Coming Home

The Gemini 3 spacecraft, with flotation collar still attached, being hoisted aboard the USS 

Intrepid (CVS-11) during recovery operations following the successful Gemini-Titan 3 flight of 

March 1965. NASA S65-18656.

66
Human Space Flight and the Problem of Reentry

As far as is known, only a single capsule has been flown twice. The Gemini 

2 capsule was refurbished and launched again on November 3, 1966, in a test 

of the Gemini-B75 configuration for the Air Force’s MOL program. This time, 

the vehicle rode atop a simulated MOL on a Titan IIIC from LC-40 at Cape 

Canaveral. The Gemini-B heat shield was identical to the standard Gemini 

heat shield with one exception: it had a 26-inch-diameter crew-transfer hatch 

cut into it to allow the astronauts to move between the capsule and laboratory. 

For the heat shield qualification (HSQ) flight, the RF-300 insulation used 

between the NASA heat shield and the capsule structure was not included 

since the reentry was expected to be more benign than the standard Gemini 

mission. In fact, the backside of the heat shield reached only 133 °F on the 

HSQ flight, compared to the Gemini’s 430 °F design limit. Researchers were 

particularly interested to learn if the new hatch would adversely affect the 

ablative performance of the heat shield. The recovered capsule showed the 

performance was satisfactory.76

The operational experience with the Gemini heat shield was decidedly better 

than Mercury, mostly because the troublesome center plug had been eliminated. The first piloted Gemini mission, the GT-3, piloted by Gus Grissom 

and John W. Young on March 23, 1965, showed that the heating effect on 

the heat shield appeared much less severe than was noted on Gemini 2. After 

the heat shield was removed from the spacecraft and dried, engineers noted 

it had lost about 20 pounds due to ablation, which was about average for the 

subsequent missions.77

Only two Gemini missions suffered any notable damage to their thermal 

protection systems. After Gemini VII, piloted by Frank Borman and James A. 

Lovell, Jr., on December 4, 1965, engineers found a small hole in one René 41 

shingle, just forward of the right-hand hatch, and a beryllium shingle around 

the bottom of the C-band antenna shifted its position enough to partially 

 75.

67
Coming Home

As a helicopter flies above, the Gemini 12 spacecraft with parachute open descends to the 

Atlantic with astronauts Jim Lovell and Edwin E. “Buzz” Aldrin aboard. NASA 66C-9516.

68
Human Space Flight and the Problem of Reentry

cover the antenna. Neither flaw was considered critical.78 On the last mission 

of the program, Gemini XII, with Lovell and Edwin E. “Buzz” Aldrin, Jr., on 

November 11, 1966, the postflight inspection revealed that the left-hand skid 

door had a shingle curled on the corner and a washer missing, and the forward 

lower centerline equipment door was deflected inward. The damage to both 

was subsequently determined to be caused by landing impact, but none of the 

anomalies was considered critical.79

Project Apollo

Although it was the third piloted U.S. program to fly, several Apollo studies 

predated even Mercury. Because of this timing, the Apollo spacecraft was not 

an evolution of Mercury or Gemini, and its development took place largely 

uncoupled from the smaller programs. Surprisingly to many people, Apollo 

was an approved program before President John F. Kennedy made his now 

famous speech challenging the Nation to land an American on the Moon 

before the end of the 1960s. At the time, a lunar landing was not planned, and 

the spacecraft was supposed to be capable of piloted circumlunar flight, along 

with several Earth-orbital tasks.80

Various reentry configurations were investigated for the Apollo spacecraft, 

including capsules and lifting reentry vehicles using lenticular shapes, lifting 

bodies, and wings. The lifting reentry shapes were considered desirable since 

they offered the capability of a horizontal landing in a relatively conventional 

manner on a prepared surface. However, these designs presented several drawbacks, chiefly that they generally weighed several thousand pounds more than 

capsules, and that they were able to cope with only prepared landing surfaces. 

It was also noted that the cross range provided by the lifting vehicles was not 

required by any of the mission models. Before the issue could be decided on 

technical merits, President Kennedy committed the Nation to a space race, and 

 78.

69
Coming Home

researchers decided a capsule was the most expedient course. The spaceplane 

would have to wait.81

During the late 1950s and into the 1960s, the three major human space 

programs—Mercury, Dyna-Soar, and Apollo—faced very different reentry 

concerns. Mercury would perform a ballistic reentry from a low-Earth orbit (or 

a ballistic trajectory for the suborbital flights). The heating rates and durations 

it faced were much the same as the ballistic missiles, and the total size of the 

capsule was not much larger or heavier. Thus, the solutions used to protect the 

missile warheads were generally adequate for the Mercury capsule. Dyna-Soar 

was a winged vehicle that used a lifting reentry, which presented a lower heating rate than did ballistic reentry but for a much longer duration. The desire 

to use the glider for multiple missions dictated a metallic heat shield almost by 

default. Apollo needed to withstand superorbital entries after its lunar flybys, 

resulting in a heating load an order of magnitude greater than Mercury, the 

ballistic missiles, or Dyna-Soar. Because Apollo would fly a semilifting entry, 

the heating duration was also quite long, although it was not as extreme as 

that of Dyna-Soar.

Engineers at NASA and prime contractor North American Aviation estimated 

that the total heat at the stagnation point might be as high as 120,000 BTU per 

square foot, with stagnation point heating rates of 520 BTU per square foot per 

second.82 It was clear that the only material that could withstand the heating at 

the stagnation point was an advanced ablator. The best material for the rest of 

the heat shield, and indeed the rest of the capsule, was less clear. For instance, 

Faget, one of the principle designers of Mercury, wanted to use René 41 and 

beryllium shingles on the afterbody, as had been done with Mercury, but the 

heating estimates were higher than could be tolerated by a shingle of any practicable thickness.83

In 1961, the possible heat shield materials that could withstand the Apollo 

environmental conditions included ablators, graphites, ceramics, and metals. 

Since the expected stagnation point heating rate was about 600 BTU per 

square foot per second, it was clear that ablators were required near the stagnation point. However, as the heating rate tapered off around the body, there 

 81.

70
Human Space Flight and the Problem of Reentry

were locations where various metal alloys, resin-impregnated ceramics, or plain 

ceramics might be used.84

Because Apollo would use a lifting reentry, followed by a precision landing 

on land, the aerodynamic qualities of the capsule were important; the outer 

mold line needed to be maintained during entry. This made the concept of 

a metallic heat shield attractive, and various NASA and contractor researchers looked into superalloy and refractory metal heat shields similar to those 

being studied for Dyna-Soar. The materials included cobalt-base Haynes 25 

and nickel-base René 41 superalloys. However, to withstand the superorbital 

entry conditions, either superalloy would have to be used as part of a Bell-type 

double-wall structure, most likely with an active water-cooling system, resulting in a very heavy installation. It was possible that the refractory metal, such 

as titanium-molybdenum alloy and F-48 columbium alloy, could withstand the 

heating conditions without melting, but the metals suffered severe oxidation 

during entry, and no suitable coatings had yet been developed. Without these 

coatings, the refractory metals could not withstand the expected environment 

without the risk of structural failure.

Efforts within the Martin Company resulted in the development of highly 

porous ceramics, including foamed zirconium oxide, foamed silicon carbide 

coated with zirconium oxide, foamed aluminum oxide, and foamed vitreous silicon dioxide. These materials exhibited temperature limits of 4,000 °F, 

3,300 °F, and 2,600 °F, respectively. The major advantage of the ceramic foams 

was that they retained their shape better than any ablator available in 1961. 

Unfortunately, they weighed considerably more than the equivalent ablators.85

The consensus of NASA researchers at Langley was that the entire capsule 

should be coated in ablator, which would also provide a small amount of radiation protection during translunar flight.86 However, the aerodynamic concerns 

still existed, and the retention of the char layer was critical to maintain the 

outer mold line of the vehicle and preserve the lifting entry characteristics. 

However, initial studies into the nylon-phenolic ablator that most nearly met 

these requirements showed that its char was of low strength and would likely 

spall due to aerodynamic pressures. This led to studies of fibrous additions to 

strengthen the char and ensure the retention of thick char layers. Eventually, 

Avco (now part of Textron) developed a superior nonreceding ablator named 

“Avcoat.” The development was largely based on research into materials for the 

high-ballistic-coefficient long-range missile warheads.

 84.

71
Coming Home

The wind tunnel experts at Langley investigated heat transfer, heating loads 

and rates, and aerodynamic and hydrodynamic stability on the Command 

Module heat shield. Avco originally proposed a layered and bonded heat shield 

similar to the one General Electric made for Mercury. In the interim, however, 

McDonnell advanced heat protection technology by developing ablator-filled 

honeycomb material for Gemini. This appeared to be a significant advancement, and Avco refined the new system to withstand the higher heating rates 

of superorbital entry. For the Gemini heat shield, a fiberglass honeycomb 

material was bonded to the back-structure and the S-3 ablator was poured into 

it and allowed to harden.87 Because of the viscosity of the Avcoat ablator, this 

technique would not work.

For Apollo, a brazed PH 14-8 stainless steel honeycomb sheet was attached 

to the structural shell, and a fiberglass-phenolic honeycomb with 400,000 

individual cells was bonded to it. Then, each cell was individually filled with a 

mixture of silica fibers and micro-balloons using a caulking gun, making the 

manufacturing process very labor intensive.88 The pot life of the Avcoat-5026 

was short, and quantities of it were frequently thrown away because technicians could not fill the cells quickly enough before it began to cure. It was a 

tedious process. The finished heat shield weighed about 32 pounds per cubic 

foot. The structural shell was attached to the aft bulkhead of the Command 

Module using 59 bolts installed in oversize holes to permit the heat shield to 

move relative to the inner structure.89

Given the limitation of the available ground test facilities, researchers 

knew they would not completely understand the conditions associated with 

superorbital entry prior to the first lunar mission. Nevertheless, they instituted 

Project FIRE (Flight Investigation of the Reentry Environment) to obtain 

data on materials, heating rates, and radio-signal attenuation on spacecraft 

reentering the atmosphere at speeds of up to 24,500 mph. NASA Langley 

announced the program, which involved both wind tunnel and flight tests, 

on February 18, 1962. The wind tunnel tests used several Langley facilities, 

including the Unitary Plan Wind Tunnel, the 8-foot High Temperature Tunnel, 

and the 9-by-6-foot Thermal Structures Tunnel.90

 87.

72
Human Space Flight and the Problem of Reentry

This 1961 General Electric concept for the Apollo spacecraft has a more X-15-like look to it than 

the blunt-body capsule that was actually built. NASA.

73
Coming Home

On March 29, 1962, NASA Langley awarded a $5 million contract to 

Republic Aviation for two experimental reentry spacecraft that were essentially 

subscale models of the Apollo Command Module. On November 27, 1962, 

NASA awarded a $2.56 million contract to Ling-Temco-Vought (LTV) to 

develop the velocity package that would simulate reentry from a lunar mission. 

The velocity package was essentially an Antares II A5 (ABL X-259d) solidpropellant rocket motor manufactured by the Allegheny Ballistic Laboratory. 

A heat shroud, manufactured for LTV by the Douglas Aircraft Company, 

protected the spacecraft from aerodynamic heating during the boost ascent. The 

velocity package also included a guidance system for maintaining stability and 

control, a telemetry system for transmitting flight data, and an ignition/destruct 

system.91 An Atlas-D booster would lift the Republic spacecraft to an altitude 

of 400,000 feet (entry interface), then the velocity package would propel the 

reentry vehicle into a –15-degree trajectory at a velocity of 37,000 ft/sec.92

The first Project FIRE flight was on April 14, 1964. The Atlas-D launch 

vehicle placed the 200-pound spacecraft into a ballistic trajectory along the 

Eastern Test Range. The Antares fired for 30 seconds, increasing the descent 

speed to 25,166 mph, and the exterior of the spacecraft reached an estimated 

20,000 °F. About 32 minutes after launch, the spacecraft splashed down into 

the Atlantic Ocean some 5,000 miles downrange, near Ascension Island. The 

second Project FIRE launch, following a similar trajectory, was on May 22, 

1965, and also used an Atlas-D from Cape Canaveral. The spacecraft entered 

the atmosphere at a speed of 25,400 mph and generated temperatures of about 

20,000 °F.93

Despite the information provided by Project FIRE, researchers believed that 

large unknowns remained concerning superorbital reentry, so NASA, North 

American, and Avco conservatively overdesigned the heat shield. The heat 

shield weighed almost 1,500 pounds, over 10 percent of the entire commandmodule weight. In fact, none of the lunar-return missions used even 20 percent 

of the available ablator, and engineers later decided they could have cut the heat 

shield thickness in half and still have had an adequate margin.

The AS-201 spacecraft/launcher test of a Saturn IB on February 26, 1966, 

had several objectives, one of which involved verifying the heat shield’s adequacy during high heat-rate reentry from low-Earth orbit. This was only partially achieved because of a fault in the electrical power system, causing a loss of 

 91.

74
Human Space Flight and the Problem of Reentry

High-angle view of Command Module 012, looking toward the –Z axis, during preparation for 

installation of the crew compartment heat shield—showing technicians working on aft bay. This 

was the ill-fated Apollo 1 spacecraft. NASA s66-41851.

steering control that resulted in a rolling reentry. This meant that data on the 

heat shield’s effectiveness was insufficient to affirm confidence in the system. 

Even so, as Apollo’s official historians concluded: “AS-201 proved that the 

spacecraft was structurally sound and, most important, that the heat shield 

could survive an atmospheric reentry.”94 A second test of the Apollo spacecraft/

Saturn IB, AS-202, was both more complex and more successful than the earlier one. Launched on August 25, 1966, this flight was intended to demonstrate 

the structural integrity and compatibility of the launch vehicle, the Command 

and Service Modules, and a host of other objectives, including the adequacy of 

 94.

75
Coming Home

the heat shield for reentry and the quality of the ablator during high reentry 

temperatures and pressures. The mission was a success, and NASA declared 

that all objectives had been achieved. In terms of the reentry sequence, the heat 

shield and all other components performed acceptably despite a steeper reentry 

than had been intended. This meant that the spacecraft landed 205 nautical 

miles uprange from the planned point and experienced greater g-forces than 

expected. Even so, heat shield engineers found that all components, in the 

words of NASA’s Apollo historians, “had come through reentry admirably.” As 

the postlaunch report noted: “During entry, spacecraft attitude was controlled 

to provide a skip trajectory resulting in a double-peak heating-rate history.”95

These tests made NASA engineers confident that they could proceed with 

the first Apollo capsule flight tests while undertaking test flights of the Saturn 

V launch vehicle, testing the capsules and their thermal protection systems 

under high-speed reentry conditions. Apollo 4 tested the worst-case scenario 

shallow reentry, which was long and drawn out with a maximum total heat 

load; and Apollo 6 tested the worst-case scenario steep reentry, which was short 

and sharp with maximum heating rate. The Apollo 4 mission was launched on 

November 9, 1967, marking the first use of a Saturn V. The payload consisted 

of a Block I Command/Service Module (spacecraft 017) and a Lunar Module 

test article (LTA-10R). Postflight inspection revealed that the Block II thermal 

protection system survived the lunar entry environment satisfactorily, and the 

aft ablative heat shield was heavily charred over its entire surface. Temperature 

data indicated surface temperatures had exceeded 5,000 °F, but the maximum 

char penetration was 0.88 inch, compared to the 1.25 inches expected.96

Apollo 6 was launched on April 6, 1968, with an unpiloted Block I 

Command/Service Module (spacecraft 020) and a Lunar Module test article 

(LTA-2R). Despite numerous problems during ascent, the mission was generally successful. Oddly, although Apollo 6 entered at 3,600 ft/sec slower than 

Apollo 4 had, the temperature on the conical section and on the leeward side 

of the heat shield was substantially higher than that of the earlier mission. 

This paradoxical situation can be attributed to three causes: Apollo 6 flew 

faster at lower altitudes, it did not skip out to as great an altitude to allow the 

ablator to cool down, and it flew approximately 80 seconds longer to reach 

the desired splashdown point. Despite the higher than expected temperatures, 

 95.

76
Human Space Flight and the Problem of Reentry

The Apollo 11 Command Module (107) is loaded aboard a Super Guppy aircraft at Ellington AFB 

for shipment to the North American Rockwell Corporation in Downey, CA. The Command Module 

was just released from its postflight quarantine at the Manned Spacecraft Center. The Apollo 11 

spacecraft was flown by astronauts Neil A. Armstrong, commander; Michael Collins, Command 

Module pilot; and Edwin E. Aldrin, Jr., Lunar Module pilot, during their lunar landing mission. 

Note damage to aft heat shield caused by extreme heat of Earth reentry. NASA s69-41985.

postflight inspection showed that the heat shield performed satisfactorily. 

During the inspection, however, engineers noticed that ablator was missing around some of the manufacturing splices in the heat shield. Since the 

edges around the missing material were not burned, the material must have 

separated after entry. Engineers had noted a similar anomaly on the AS-202 

Command Module but not on the Apollo 4 capsule. The common thread 

was that AS-202 and Apollo 6 had remained in the water for much longer 

than Apollo 4 had. Improved manufacturing processes eliminated the splices 

between the honeycomb segments on all Block II heat shields used for the 

lunar missions.97

 97.

77
Coming Home

Apollo 11 Command Module heat shield inspected on November 6, 1967. NASA S67-48970.

The operational experience with the Apollo heat shield was highly satisfactory, with no significant anomalies noted on any of the missions. Apollo 8, 

the first lunar flyby mission in December 1968, returned with a heat shield 

that was charred less than that of the Apollo 4 test. The char depth was 

0.6 inch at the stagnation point, and only 0.4 inch elsewhere. There was, 

however, a fair amount of impact damage when the capsule splashed down. 

Apollo 10, another lunar return, entered at a flightpath angle 0.02 of a degree 

steeper than planned, subjecting the crew to a maximum 6.78 g’s. Despite 

this anomaly, the heat shield performed satisfactorily. Returning from the 

first lunar landing, Apollo 11 crossed the entry interface at 400,000 feet and 

at 36,195 ft/sec at a flightpath angle of –6.488 degrees (versus a planned 

36,194 ft/sec and –6.483 degrees). Heat shield performance was listed as 

nominal. Despite Apollo 13’s faster than normal entry, Lovell reported that 

78
Human Space Flight and the Problem of Reentry

the craft’s landing decelerations were mild in comparison to Apollo 8. The 

postflight report revealed nothing unexpected about the performance of the 

heat shield. The thermal protection system worked as expected on Apollo 

15, but there was an anomaly regarding the recovery parachutes. At approximately 10,500 feet altitude, the three main parachutes deployed normally, 

but 4,000 feet later, one collapsed and deflated. Engineers believed that an 

unexpected dump of monomethyl hydrazine caused two aluminum riser 

connectors to fail, and these parts were changed to Inconel 718 for future 

flights. Although the capsule impacted the Pacific Ocean at a higher-thannormal velocity, the capsule was not damaged nor was the crew injured. The 

remaining lunar-return flights were accomplished without incident regarding 

the heat shield or parachute performance.98

 98.

79
Project Mercury recovery practice aboard USS Strong on August 17, 1959. NASA.

80
81

CHAPTER 3

Returning from space is just as difficult and risky as reaching into space. Indeed, 

coming home may well be just as risky as launch. For the United States, NASA 

has undertaken 168 flights (counting suborbital), including 135 for the Space 

Shuttle program. In addition, Russia (or the Soviet Union) has undertaken 

119 flights, mostly for the Soyuz spacecraft program (including suborbital 

flights); China has made 3 flights; and the private SpaceShipOne completed 

3 flights (suborbital flights) for a total of 293 human space flights throughout 

the more than 50-year-long space age.1

 During this time, there have been four 

accidents resulting in the loss of crew for a 1.36-percent failure rate, by flight. 

These accidents include the following:

•	 April 24, 1967: A parachute failure caused the Soyuz spacecraft to 

hit the ground at 300 mph, killing one cosmonaut.

•	 June 30, 1971: Three cosmonauts died on descent when their spacecraft decompressed prematurely.

•	 January 28, 1986: On ascent, the Space Shuttle Challenger broke 

apart, killing seven crewmembers.

•	 February 1, 2003: On descent, the Space Shuttle Columbia broke 

apart, killing seven crewmembers.

Significantly, the number of failures on descent exceeds the number of ascent 

failures, even though many treat ascent as the more risky venture. Accordingly, 

the challenge of returning to Earth is a subject of great significance. How might 

it most expeditiously be accomplished for the human spaceflight program?

Virtually all of the early concepts for human space flight involved spaceplanes that flew on wings to runway landings. Eugen Sänger’s antipodal bomber 

of the 1920s did so, and Wernher von Braun’s popular concepts of the 1950s 

did the same. However, these concepts proved impractical for launch vehicles 

1. These include 2 suborbital X-15 flights, 2 suborbital Mercury flights, 4 orbital Mercury flights, 10 

Gemini flights, 12 Apollo flights, 3 Skylab flights, and 135 Space Shuttle flights; 6 Vostok flights, 

2 Voskhod flights, 1 suborbital Soyuz flights, 110 orbital Soyuz flights, 3 Shenzhou flights, and 3 

suborbital SpaceShipOne flights.
Coming Home

available during the 1950s, and capsule concepts that returned to Earth via 

parachute proliferated, largely because they represented the art of the possible 

at the time. The United States was racing against the Soviet Union for primacy 

in space, which necessitated placing a human in space as expeditiously as possible. In such a Cold War environment, pursuing what was possible rather than 

what was desirable proved irresistible.

All of the U.S. spacecraft up to the Space Shuttle, as well as the Soviet/Russian 

and Chinese piloted capsules, used from one to three parachutes for return to 

Earth. The American capsules landed in the ocean and were recovered by ship. 

This exposed them to corrosive saltwater while waiting for expensive recovery 

efforts. The Soviet/Russian and Chinese spacecraft have always been recovered 

on land, which presented the crew with a harder landing than would be the case 

at sea. For Project Gemini, NASA toyed with the possibility of using a paraglider 

that the Langley Research Center was developing for “dry” landings, instead of 

“splashdowns” in water and recovery by the Navy. At sum, this represented an 

attempt to transform the capsule into a spaceplane. Unfortunately, the engineers 

never did get the paraglider to work properly and eventually dropped it from the 

program in favor of a parachute system like the one used for Mercury.

During the period between the creation of NASA in 1958 and the Moon 

landings of the late 1960s and early 1970s, NASA developed capsules with 

blunt-body ablative heat shields and recovery systems that relied on parachutes. 

Most people are familiar with parachutes, the umbrellalike devices usually 

made of a soft fabric that slow the motion and retard the descent of a falling 

object by creating drag as it passes through the air. The parachute’s seeming 

simplicity masked significant challenges for recovery from space as NASA 

pondered the details of how to implement it successfully for landings from 

space. Every piloted U.S. mission that has used parachutes has been successful, having safely returned crews from space back to Earth via water landings. 

However, history did not have to unfold in that way.

The Department of Defense (DOD) tested parachute-landing systems 

during Project SCORE in 1958 and employed the concept throughout the 

CORONA satellite reconnaissance program of the 1960s, when it would 

snatch in midair return capsules containing unprocessed surveillance footage dangling beneath parachutes. During the Mercury program, astronauts 

rode a blunt-body capsule with an ablative heat shield to a water landing 

and rescue at sea by the Navy. Gemini later used a similar approach, but 

NASA engineers experimented with a Rogallo wing and a proposed landing at the Flight Research Center on skids similar to those employed on the 

X-15. When the engineers working on the Rogallo wing concept failed to 

make the rapid progress required to meet project timelines, NASA returned 

to the parachute concept used in Project Mercury. Engineers incorporated 

82
Landing Under Canopies

essentially the same approach used in the Apollo program, although with 

greatly improved ablative heat shields and much larger and more complex 

parachute systems.

Pioneering Recovery: Project Mercury

In theory, the deployment of a parachute for a gentle landing upon returning 

to Earth was the most simple of engineering tasks. Such was not to be the 

case, however, because of a succession of challenges ranging from hypersonic 

reentry to alternative freezing and overheating of the system while in Earth’s 

orbit. As intended during Mercury, the spacecraft would undergo reentry 

and, after deceleration to about 270 mph, “at about 21,000 feet a six-foot 

diameter drogue parachute [would] open to stabilize the craft. At about 

10,000 feet, a 63-foot main landing parachute [would] unfurl from the 

neck of the craft. On touchdown, the main parachute [would] jettison.”2 A 

landing bag, inflated from behind the heat shield, would deploy to soften the 

impact just prior to hitting the water. Upon landing, additional bags inflated 

around the nose of the craft to keep the capsule upright in the water, and the 

parachutes were released. Finally, the astronaut would open the hatch only 

after Navy frogmen had secured the capsule and a recovery helicopter had 

connected to the vehicle.3

The full-scale testing of the Mercury recovery system began at the Langley 

Research Center, Hampton, VA, on October 9, 1958. The first task was to 

undertake tests of parachute deployment and spacecraft stability, first from 

helicopters and later from U.S. Air Force C-130s at Pope AFB, NC, where the 

317th Military Airlift Wing was located. These tests were relatively simple, at 

first using a concrete-filled 55-gallon drum attached to a deployment system. 

As the official NASA history of Project Mercury concluded, “By early January 

more than a hundred drops of drums filled with concrete and of model capsules 

had produced a sizable amount of evidence regarding spacecraft motion in free 

falls, spiraling and tumbling downward, with and without canopied brakes, to 

impacts on both sea and land.”4 These efforts demonstrated the adequacy of the 

 2. “Mercury-Atlas 6 at a Glance,” NASA release 62-8, NASA Historical Reference Collection, p. 1–6.

83
Coming Home

The sequence of events, from launch to parachute opening, of a beach abort test for the 

Mercury capsule with a launch escape motor on April 13, 1959. NASA L59-2768.

mechanical system proposed to deploy parachutes from the Mercury spacecraft. 

Using the C-130s, NASA graduated to using full-scale spacecraft and operating parachutes on these drop tests; in the process, NASA was working out the 

operational features of the parachute system and the processes necessary to 

ensure success in returning from space. The project then moved to Wallops 

Island, VA, and continued tests, “to study the stability of the spacecraft during 

free fall and with parachute support; to study the shock input to the spacecraft 

by parachute deployment; and to study and develop retrieving operations.”5

Not everything went as planned in these tests. NASA engineers found 

that the main parachute was prone to experience “squidding,” a phenomenon also referred to as “breathing” or “rebound.” Because of conditions in 

the atmosphere between an approximate range of 70,000 feet and 10,000 

feet, there were “snatch” forces, shock waves, and stability difficulties that 

resulted in the partial opening of the parachute—occurrences and conditions 

that proved unacceptable for the spacecraft returning to Earth. These problems were corrected through the acquisition of a 63-foot-diameter ribbon 

ring-sail parachute. This ribbon ring-sail parachute, developed by Theodor W. 

 5. “Mercury Parachute,” 

84
Landing Under Canopies

Knacke, had superior stability than earlier parachute designs, making it ideal 

for human space programs.6

 The parachute change resolved the problems that 

arose during testing and served effectively throughout the Mercury, Gemini, 

and Apollo programs.

For flight tests, NASA acquired 6 main-parachute and 12 drogue-parachute 

canisters from the Goodyear Aircraft Corporation—3 each to be dedicated to 

the Little Joe and Big Joe tests of the Mercury capsule. These found use in a test 

effort concerned with the problem of landing impact, specifically the question 

of whether the spacecraft could touch down at a speed of less than 30 ft/sec. 

Ensuring this capability, as well as a vertical landing, proved an exasperating 

challenge. NASA initially hoped for a land recovery, but the shock of impact 

was too great. NASA then pursued landing tests in hydrodynamics laboratories. Reducing the speed prior to impact required such efforts as the following:

McDonnell engineers fitted a series of four Yorkshire pigs into 

contour couches for impact landing tests of the crushable aluminum honeycomb energy-absorption system. These supine swine 

sustained acceleration peaks from 38 to 58 g before minor internal 

injuries were noted. The “pig drop” tests were quite impressive, 

both to McDonnell employees who left their desks and lathes to 

watch them and to STG engineers who studied the documentary 

movies. But, still more significant, seeing the pigs get up and walk 

away from their forced fall and stunning impact vastly increased 

the confidence of the newly chosen astronauts that they could 

do the same.7

As McDonnell engineers concluded, “Since neither the acceleration rates 

nor shock pulse amplitudes applied to the specimens resulted in permanent or 

disabling damage, the honeycomb energy absorption system of these experiments is considered suitable for controlling the landing shock applied to the 

Mercury capsule pilot.”8

Notwithstanding these tests, cushioning the blow to astronauts during landing and recovery operations led to a succession of project innovations. The first 

was the creation of form-fitting contour couches that helped distribute the 

 6. George M. Low, “Status Report No. 11” (April 6, 1959); Swensen, Grimwood, and Alexander, 

85
Coming Home

g-forces associated with reentry and landing throughout the body, allowing 

astronauts to withstand the loads without undue injury. During flight, “the 

astronaut, wearing the Mercury full-pressure suit, was positioned in his contour 

couch in the semisupine position and secured by shoulder and lap harnesses.” 

NASA also investigated the feasibility of including a crushable honeycombshape section of metal foil between the heat shield and the capsule to absorb the 

principle shocks of landing. In the end, NASA included a pneumatic landing 

bag in the Mercury capsule.9

In June 1959, NASA contracted with Northrop to design and fabricate the 

landing system for Project Mercury. Northrop had a long history in this arena, 

having built aircraft landing and recovery systems as far back as 1943, when it 

developed the first parachute-recovery system for pilotless aircraft. Northrop 

adopted the 63-foot ring-sail main parachute. By August 1959, Northrop had 

completed initial drop tests for the 63-foot ring-sail main parachute. It had also 

changed the drogue parachute configuration, taking it from 19.5 percent porosity with a flat circular ribbon chute to a 28-percent porosity, 30-degree conical 

canopy. The drogue parachute quickly moved through its certification and was 

qualified for deployment at speeds up to Mach 1.5 and altitudes up to 70,000 

feet. This was far beyond the standard envelope for drogue chute operations: 

deployment at 40,000 feet and below Mach 1. By September 19, 1959, after 

approximately 18 months of effort, NASA completed the qualification tests 

for the Mercury spacecraft-landing system. In all, project staff and contractors 

had undertaken “56 airdrops of full-scale engineering models of the Mercury 

spacecraft from C-130 aircraft at various altitudes up to 30,000 feet and from 

helicopters at low altitudes to simulate off-the-pad abort conditions.”10

Full-scale tests of the system in space began on December 4, 1959, when 

a mission flying Sam, an American-born Rhesus monkey, tried out the recovery system in a launch of Mercury Little Joe 2 (LJ-2) at Wallops Island, VA. 

Controllers initiated an abort sequence after 59 seconds of flight at an altitude 

of 96,000 feet and a speed of Mach 5.5. The drogue and main parachutes 

deployed properly and returned Sam safely to the ocean. Recovery of the capsule took about 2 hours, but Sam came through the flight fine. Additional tests 

thereafter found no difficulties with the system.11

 9. 

86
Landing Under Canopies

Little Joe 1 (LJ-1) was launched on October 4, 1959, at Wallops Island, VA. This was the first 

attempt to launch an instrumented capsule with a Little Joe booster. Only the LJ-1A and the 

LJ-6 used the space metal/chevron plates as heat reflector shields, as they kept shattering. 

Little Joe was used to test various components of the Mercury spacecraft, such as the emergency escape rockets. NASA L-1960-00104.

87
Coming Home

Further tests of the Mercury capsule’s extended-mission parachutes, under 

the name Project Reef, took place in 1962. Beginning on June 26, 1962, Project 

Reef began 20 airdrops to test the ability of the Mercury 63-foot ring-sail main 

parachute’s capability to support a higher projected spacecraft weight for an 

extended-range or more than 1-day mission. Tests indicated that the parachute 

could support the heavier spacecraft without undue stress.12

For the flights of Project Mercury, the configuration of the parachuterecovery system consisted of the following elements:

Above the astronaut’s cabin, the cylindrical neck section contains 

the main and reserve parachute system.

Three parachutes are installed in the spacecraft. The drogue 

chute has a six-foot diameter, conical, ribbon-type canopy with 

approximately six-foot long ribbon suspension lines, and a 

30-foot long riser made of dacron to minimize elasticity effects 

during deployment of the drogue at an altitude of 21,000 feet. 

The drogue riser is permanently attached to the spacecraft antenna 

by a three point suspension system terminating at the antenna in 

three steel cables, which are insulated in areas exposed to heat.

The drogue parachute is packed in a protective bag and stowed 

in the drogue mortar tube on top of a light-weight sabot or plug. 

The sabot functions as a free piston to eject the parachute pack when 

pressured from below by gasses generated by a pyrotechnic charge.

The function of the drogue chute is to provide a backup stabilization device for the spacecraft in the event of failure of the Reaction 

Control and Stabilization System. Additionally, the drogue chute 

will serve to slow the spacecraft to approximately 250 feet per second at the 10,000 foot altitude of main parachute deployment.

The reserve chute is identical to the main chute. It is deployed 

by a flat circular-type pilot chute.

Other components of the landing system include drogue 

mortar and cartridge, barostats, antenna fairing ejector, and sea 

marker packet.

Following escape tower separation in flight, the 21,000 and 

10,000 foot barostats are armed. No further action occurs until 

the spacecraft descent causes the 21,000 foot barostat to close, 

activating the drogue ejection system.

 12.

88
Landing Under Canopies

Assembling the Little Joe capsules. The capsules were manufactured “in-house” by Langley 

Research Center technicians. Three capsules are shown here in various stages of assembly. The 

escape tower and rocket motors shown on the completed capsule would be removed before shipping to Wallops Island. These test articles met the weight and center-of-gravity requirements of 

Mercury and withstood the same aerodynamic loads during the exit trajectory. NASA L59-4944.

89
Coming Home

Two seconds after the 10,000 foot barostat closes, power is supplied to the antenna fairing ejector -- located above the cylindrical 

neck section -- to deploy the main landing parachute and an underwater charge, which is dropped to provide an audible sound landing point indication. The ultra-high frequency SARAH radio then 

begins transmitting. A can of sea-marker dye is deployed with the 

reserve chute and remains attached to the spacecraft by a lanyard.

On landing, an impact switch jettisons the landing parachute 

and initiates the remaining location and recovery aids. This 

includes release of sea-marker dye with the reserve parachute if 

it has not previously been deployed, triggering a high-intensity 

flashing light, extension of a 16-foot whip antenna and the initiation of the operation of a high-frequency radio beacon.

If the spacecraft should spring a leak or if the life support system should become fouled after landing, the astronaut can escape 

through this upper neck section or through the side hatch.13

This sequence of operations worked effectively on all six of the Project 

Mercury piloted missions as well as during all tests.

Recovery from Long-Duration 

Earth-Orbital Flight: Project Gemini

In the fall of 1961, NASA began Project Gemini as a means of bridging the 

significant gap in the capability for human space flight between what Project 

Mercury achieved and what would be required for a lunar landing—already 

an approved program. NASA closed most of the gap by experimenting and 

training on the ground, but some issues required experience in space. This 

requirement became immediately apparent in several major areas, including 

the following major mission requirements, as defined in the Gemini crew 

familiarization manual:

•	

•	

•	

Accomplish 14-day earth orbital flights, thus validating that humans 

could survive a journey to the Moon and back to Earth.

Demonstrate rendezvous and docking in Earth orbit.

Provide for controlled land landing as the primary recovery mode.

 13.

90
Landing Under Canopies

•	

•	

Develop simplified countdown techniques to aid rendezvous missions 

(lessens criticality of launch window).

Determine man’s capabilities in space 

during extended missions.14

These major initiatives defined the Gemini 

program and its 10 piloted space flight missions conducted from 1965 to 1966.15

NASA originally conceived of Project 

Gemini as a larger Mercury Mark II capsule, 

but soon it became a totally different vehicle 

that could accommodate two astronauts for 

extended flights of more than 2 weeks. It pioneered the use of fuel cells instead of batteries 

to power the ship and incorporated a series 

of hardware modifications. Its designers also 

toyed with the possibility of using a paraglider 

that was being developed at Langley Research 

Center for ground landings instead of splashdowns in water and recovery 

by the Navy.16 The whole system was to be powered by the newly developed 

Titan II launch vehicle, another ballistic missile developed for the Air Force. 

A central reason for this program was to perfect techniques for rendezvous 

and docking, so NASA appropriated from the military some Agena rocket 

upper stages and fitted them with docking adapters to serve as the targets for 

rendezvous operations.

As it turned out, the parachute system that NASA used for Gemini was 

similar to that used on Mercury, but it had some additional features. As one 

analysis noted during the program:

Artist concept of Gemini parachute 

landing sequence from highaltitude drogue chute deployment 

to jettisoning of parachute. NASA 

S65-05398.

 14.

91
Coming Home

A single-parachute landing system is used on Gemini spacecraft, 

with the ejection seats serving as a backup. In the normal landing 

sequence, an 8-foot-diameter drogue parachute is deployed manually at approximately 50,000 feet altitude. Below 50,000 feet, 

this drogue provides a backup to the reentry control system for 

spacecraft stabilization. At 10,600 feet altitude, the crew initiates 

the main parachute deployment sequence, which immediately 

releases the drogue, allowing it to extract the 18-foot-diameter 

pilot parachute. At 2.5 seconds after sequence initiation, pyrotechnics release the recovery section, to which the pilot parachute is attached and in which the main parachute is stowed. As 

the reentry vehicle falls away, the main parachute, an 84-footdiameter ring-sail, deploys. The pilot parachute diameter is sized 

such that recontact between the recovery section and the main 

parachute will not occur during descent. After the crew observes 

that the main parachute has deployed and that the rate of descent 

is nominal, repositioning of the spacecraft is initiated. The spacecraft is rotated from a vertical position to a 35° nose-up position 

for landing. This landing attitude reduces the acceleration forces 

at touchdown on the water to values well below the maximum 

which could be tolerated by the crew or by the spacecraft.17

As NASA engineers once noted, “Main-parachute deployments take place 

in full view of the crew, and it is quite a beautiful and reassuring sight.” 18 At 

landing, the crew experienced a shock, as the “amount of wind drift, the size 

of the waves, and the part of the wave contacted also vary the load. Even the 

hardest of the landings has not affected crew performance.”19

Interestingly, Project Gemini leaders had an entirely different approach to 

landing planned for this second piloted space flight program for the United 

States. Dangling from beneath a parachute and being rescued at sea might 

have been acceptable for the Mercury program, since winning the space race 

was an overriding motivation. However, NASA found distasteful the prospect 

of continuing this parachute-recovery model indefinitely. For that reason, the 

Gemini project managers conceived and for several years relentlessly pursued 

an inflatable paraglider modeled on the airfoil invented by NASA engineer 

 17.

92
Landing Under Canopies

This is the proposed flight profile for the Gemini spacecraft paraglider concept in 1962. NASA.

Francis M. Rogallo, which eventually spawned the hang-gliding movement of 

the latter half of the 20th century. NASA’s objective was a controlled, guided, 

soft landing on a land target of its choice. NASA human space flight head 

Brainerd Holmes stated in 1962:

I think that the significant thing is that we would like to land on 

land, and not on a hostile sea. I think it is significant that we are 

planning, by having a L/D something greater than zero—as Dr. 

[Joseph] Shea said, .5—...to be able to guide this vehicle, if you 

will, just by controlling his attitude and thus having this offset 

center of gravity, to a localized landing area which might be an 

area ten miles on a side, something like that, and then much more 

localized through a parachute or paraglider.20

 20.

93
Coming Home

NASA had previously explored the possibility of recovery from space on 

the Rogallo wing. The first studies investigated the possibility of employing 

an inflated paraglider that would reenter the atmosphere at about 5,000 mph, 

after inflation and separation from an Aerobee sounding rocket. It was on this 

Inflatable Micrometeoroid Paraglider (IMP) project that flight tests achieved 

enough success to warrant continued research and development. Additionally, 

in 1963 the Air Force supported work on a reentry paraglider to rescue a 

crewmember from an orbiting space station. This program was called FIRST, 

or Fabrication of Inflatable Reentry Structures for Test.21

Pursuit of the Gemini Paraglider

For Project Gemini, efforts to develop a paraglider began in earnest on 

November 20, 1961. Through 1964, NASA engineers aggressively pursued 

the idea of such a practical landing system for the Gemini spacecraft. In theory, 

the spacecraft would carry the paraglider safely tucked away through most of its 

mission. After reentering the atmosphere from orbit the crew would deploy the 

wing and, having converted the spacecraft into a makeshift glider, they could 

fly to a runway landing. It worked in theory, but not in practice.

The basics of the story are well known.22 At the start of the Gemini program 

in 1961, NASA considered having the two-person Gemini capsule land on 

a runway after its return from space, rather than parachute into the ocean. 

This controlled descent and landing was to be accomplished by deploying an 

inflatable paraglider wing. First, NASA built and tested the Parasev, a singleseat, rigid-strut parasail, designed much like a huge hang glider, to test the 

possibility of a runway landing. The space agency then contracted with North 

American Aviation to undertake a design, development, and test program for 

a scaled-up spacecraft version of the concept. A full-scale, two-pilot Test Tow 

Vehicle (TTV) was also built to test the concept and train Gemini astronauts 

for flight. The TTV tested maneuvering, control, and landing techniques. A 

helicopter released the TTV, with its wings deployed, over the dry lakebed at 

Edwards AFB, CA, where it landed safely. Scale models of the capsules released 

at higher altitudes and faster speeds sought to duplicate reentry conditions. 

The system never worked well enough to use on the Gemini program (largely 

 21.

94
Landing Under Canopies

On June 26, 1959, then-Langley researcher Francis Rogallo examined the Rogallo wing in the 

Langley 7-by-10-foot wind tunnel. Originally conceived as a means of bringing piloted spacecraft to controlled, soft landings, Rogallo’s concept was avidly embraced by later generations of 

hang-gliding enthusiasts. NASA L59-4345.

because of control and stability issues), but the Paraglider Landing System 

Program proved useful in developing alternate landing techniques.23

But there is more to the story. In the late 1940s and early 1950s, Francis 

Rogallo, an original thinker and kite-flying enthusiast working in the Langley 

Memorial Aeronautical Laboratory’s 7-by-10-foot Tunnel Branch, experimented with the idea of a flexible wing that could enable individuals to fly. 

Working with his wife, Gertrude, they developed the familiar V-shaped flexible wing that is ubiquitous in hang gliding today, receiving patents for their 

 23.

95
Coming Home

work in 1948 and 1952. They initially focused on paragliders that resembled 

sailboat designs, but they quickly moved to concepts that had more in common 

with parachutes. When the Rogallos read the Collier’s series of articles on 

space flight beginning in 1952, they immediately recognized a use for their 

paraglider as a method for landing when returning from space. At a 1963 

American Astronautical Society conference, Rogallo said, “I thought that the 

rigid-winged gliders might better be replaced by vehicles with flexible wings 

that could be folded into small packages during the launching.”24 He pursued 

this idea for more than a decade before having much success in interesting space 

program officials in putting money into it. For Rogallo, it made more sense 

to fly home rather than splash down in the ocean, and his NASA colleagues 

in the space research-and-development (R&D) community agreed. They had 

recognized that the ballistic capsule dangling from the bottom of a parachute 

was an acceptable solution to landing, but it was also an inelegant one few 

wanted to use indefinitely. The paraglider concept was tailor-made to enable 

a step beyond the state of the art.25 It also satisfied the space agency’s desire to 

“advance spacecraft technology,” a mission viewed as critical, even as NASA 

undertook operational activities in space.26

NASA contracted with North American in 1961 to develop a paragliderrecovery system for Gemini. As the spacecraft fell through the atmosphere back 

to Earth, its ablative heat shield would protect the passengers and the machine 

from harm while slowing it to subsonic speeds. Then, a carefully designed and 

packed paraglider stowed in the spacecraft would be deployed beginning at about 

60,000 feet, and by 20,000 feet the descending spacecraft would take on the 

characteristics of a hang glider, and the astronauts would bring the craft to a 

controlled landing on either water or land. “In this application the wing is stowed 

within the spacecraft until after the high-temperature reentry and descent to low 

altitude, about 50,000 feet, where the speed is subsonic,” states a 1963 study of 

the project. “At this point the wing is deployed and the pilot directs the vehicle 

toward a predetermined landing spot by means of manual control in pitch and 

roll. The pilot executes a flare maneuver at an altitude of some 100 feet above 

 24.

96
Landing Under Canopies

The Gemini paraglider was intended to enable NASA astronauts to fly back and land on solid 

ground rather than be rescued at sea by the U.S. Navy. NASA.

the ground, and the spacecraft lands at a low sink rate.” Skids from the spacecraft 

would serve as landing legs for the crew returning from space.27

Should this system not work as intended, the Gemini spacecraft would have 

ejection seats and crew parachutes—which also would have helped crews escape 

during launch if something went wrong—that would ensure the safe return of 

the crew. “The use of this system may also be necessary when aborted missions 

result in recovery at some point where landing conditions are unsuited for use 

of the parawing landing system.”28 As conceptualized in 1962 at the start of 

 27.

97
Coming Home

the R&D program, the optimum reentry and landing phase took place in the 

following way:

1. For normal recovery the parawing would be employed for the final 

phase of letdown and landing. The parawing is initially released at 

approximately 35,000 feet and is fully deployed at 20,000 feet. A 

gliding descent is maintained below this altitude which can be pilotcontrolled to achieve landing at a preselected touchdown site.

2. Control is provided in both pitch and roll by utilizing the same 

hand control device used during reentry.

3. Glide is terminated at about 500 feet at which point the glide angle 

is increased to build up velocity for the flare maneuver. Flare is initiated at approximately 70 feet altitude, and the spacecraft is brought 

into the landing attitude with a rate of descent less than 5 ft/sec and 

a forward velocity of about 50 ft/sec.

4. The high-drag rear-landing skids and a low-drag forward skid are 

employed for touchdown to provide stability during the landing 

slide.

5. Instrumentation for control of the parawing would make use of the 

regular rate-and-attitude instruments used during other phases of 

flight. Additional instrumentation would include an airspeed meter, 

rate-of-descent indicator, and an altimeter.

6. The ejection seat/crew parachute backup system serves as an emergency escape system in the event of a malfunction in the parawing 

landing system. The use of this system may also be necessary when 

aborted missions result in recovery at some point where landing 

conditions are unsuited for use of the parawing landing device.29

NASA engineers pursued this option aggressively, and it is important to 

acknowledge how exhaustively they proceeded with the paraglider R&D program. The program engineers died very hard regarding this program once it 

proved untenable; in the various engineering memoranda, they talked openly 

about the many problems the program encountered but also about the possibilities of success. Some of those problems encountered in test operations 

include the following:

1. Leakage of the inflatable structure.

2. Failure of the structure to withstand the desired internal pressure 

and consequently blow out.

3. Malfunction of the pressurization attack release.

4. Malfunction of the attack release hook.

 29.

98
Landing Under Canopies

5. Rupture of the apex causing the pressurization valve to be torn from 

the wing.

6. Breakage of shroud lines due to snatch loads.

7. Spiral descent to impact as a result of entangled shroud lines.

8. Awkward deployment causing rotation of the spacecraft and wing.

9. Buckling of the wing due to the ratio of dynamic pressure to internal 

pressure.

10. Nonrelease of the drogue chute during deployment.

11. Mechanical failure of activating devices and electrical systems.

Other problems that had not yet been demonstrated but were certainly possible also prompted engineers to pursue solutions. “From the pilot’s point of 

view there is a big question as to whether the complicated control system will 

provide the pilot with sufficient control of the vehicle during all possible flight 

conditions such as wind gusts, precipitations, and limited visibility,” the engineers acknowledged. They were especially concerned about the performance 

of the spacecraft and the paraglider during return to Earth in harsh weather 

conditions. “If high winds are encountered,” they stated, “then it is very likely 

that the system will collapse or tumble and become entangled with the shroud 

lines.” Engineers also wanted to ensure that the spacecraft could come down in 

water as well as on land, in case that was necessary during an emergency. They 

also sought a release mechanism from the paraglider to allow use of the ejection 

system should an emergency arise at any point during recovery. Finally, NASA 

engineers worked to ensure that the paraglider would deploy appropriately 

under all conditions.30

Because of these problems with the paraglider system, the Gemini project manager initiated a “plan B” for recovery: the development of a parachute system similar to that used in Project Mercury.31 Reasonable confidence 

abounded that this system, proven effective, could also work for the Gemini 

program. Gemini design engineers also toyed with a parasail concept—a 

canopy that looked more like a parachute than a paraglider but had modest 

maneuverability. NASA contracted with Pioneer Parachute Company for the 

testing of the parasail concept in 1963. The parasail concept was something 

of a cross between the parachute and the paraglider. It had the following char-

 30.

99
Coming Home

This test used a Volkswagen Beetle as the platform to test aerodynamics of the Gemini paraglider concept. NASA.

acteristics, as noted in a 1962 proposal for the R&D program from Pioneer 

Parachute Company.

1. It is constructed of very low porosity fabric.

2. It is designed to wing and airfoil theories.

3. It uses slots and vents to relieve excess pressures and to guide air into 

and out of the canopy.

4. It makes use of central suspension lines to achieve high drag.

5. Originally this canopy was not designed for deployment at any 

appreciable speed.

6. It has been built in one size (23.2 ft equiv. nom. Dia.) only.

7. Little is known of opening forces and canopy stresses.32

With challenges before it, NASA and contractor engineers believed that this 

approach had potential for future space vehicle recovery and landing.

As a 1965 report noted: “The total Parasail land-landing system includes 

the gliding parachute for local obstacle avoidance and landing attenuation 

rockets fired just above the surface to reduce descent velocity prior to impact. 

 32.

100
Landing Under Canopies

The program we have been conducting has included component development of the parachute, the landing attenuation rockets, turn control motors 

for steering, altitude sensors, pilot display and visual reference system, and 

landing gear.” NASA successfully flew this system four times, and while it 

was developed too late for the Gemini program, many believed it could be 

incorporated into the Apollo recovery system.33 Starting April 21, 1965, a test 

program capsule, dubbed “El Kabong,” was dropped from an Air Force Reserve 

C-119 from an altitude of 11,500 feet at Fort Hood, TX. During descent tests, 

personnel steered the parasail by radio command to operate motors on the 

capsule that controlled flap angles on the sail and allowed modest control of 

drift. As the test spacecraft neared the ground, sensors ignited two 6,000-pound 

thrust motors that reduced capsule speed to less than 10 ft/sec. The capsule 

then landed on tricycle landing gear. The first two tests did not go as planned, 

but on the third, flown on July 31, 1965, the capsule landed within 40 feet of 

its target at Fort Hood’s Antelope Mound tank range. “We’ve got a winner!” 

an engineer named Lee Norman announced at the conclusion of the test. He 

announced correctly (as all earlier tests had taken place over water), “This is 

the first successful landing [of a spacecraft] in this country!”34

However, the parasail was not the recovery system of choice, rather it was 

plan B; NASA’s goal was to successfully integrate the inflatable paraglider. To 

prove the concept, the space agency first developed and tested the Parasev, the 

first one of which was tested in Langley’s full-scale wind tunnel in January 

1962. Parasev-1 used a hang-glider-type steel-tube frame onto which the fabric 

was fixed. The powerless, lightweight steel-tube vehicle was taken to what 

would become the Dryden Flight Research Center at Edwards AFB for flight 

tests. It was towed by either a ground vehicle or a light airplane to altitudes of 

up to 12,000 feet for a free-glide, and then it came in for a 100-mph landing 

on the dry lakebed. Early flights with experienced test pilots (Gus Grissom, 

Milton O. Thompson, and Neil Armstrong among them) proved extremely 

 33.

101
Coming Home

tricky. One pilot noted it flew as if “controlled by a wet noodle.” During one 

ground tow, the pilot got out of phase with a lagging control system and the 

craft developed a rocking motion that grew worse the more he tried to correct 

for it. It was a version of pilot-induced oscillations, which the Space Shuttle 

would experience a decade and a half later. Parasev-1 did a wing over into the 

lakebed, demolishing the craft and injuring the pilot. It was rebuilt several 

times and eventually made over 100 flights.35

Eight of NASA’s Flight Research Center’s engineers conducted this test 

program under the direction of Charles Richards. Its name, Parasev, an abbreviation of Paraglider Research Vehicle, served as useful shorthand for a program 

that the public saw as cute but also useful. Parasev-1 resembled an unpowered 

tricycle with an angled tripod mast, on top of which sat the Rogallo wing. 

Constructed totally in-house, Parasev-1 received Federal Aviation Agency (now 

Federal Aviation Administration) approval on February 12, 1962. The vehicle 

itself was unremarkable; it was an open frame “fuselage” made from welded 

4130 titanium tubing, with aluminum tubing comprising the keel and wing 

leading edges. The pilot sat in the open, strapped to a seat with no enclosure 

of any kind. The pilot could control the vehicle by tilting the Parasev wing fore 

and aft, and turn it by tilting the wing with a control stick. The control stick, in 

contrast to a normal aircraft’s, was mounted overhead in front of the pilot’s seat. 

The wing, made of Dacron and linen, was sown for Parasev-1 by a sailmaker.36

The Parasev proved a difficult flying vehicle to master. Tow tests behind a 

utility truck first got the vehicle airborne. Later, “souped-up” cars, motorcycles, 

and finally a helicopter towed the Parasev-1. One writer described the process 

of flying the vehicle as follows:

Floating it along the ground while getting to know the handling 

characteristics before attempting more ambitious manoeuvres, it 

was not an untroubled learning curve. Flapping violently in the 

 35.

102
Landing Under Canopies

The Parasev-1-A (Paraglider Research Vehicle) and the tow airplane (450-horse power Stearman 

sport biplane) sitting on Rogers Dry Lake, CA. The control system in the Parasev-1-A had a conventional control stick position and was cable-operated; the main landing gear used shocks and bungees 

with the 150-square-foot wing membrane made of 6-ounce unsealed Dacron. NASA E-8712.

wind, the lobes bulging and straining in a most alarming way, the 

linen wing was giving problems, including flutter at the trailing 

edges. Longitudinal and lateral stick forces were considerable. 

Flying the Parasev was, said [NASA research pilot Bruce] Peterson, 

more difficult than the later, more sophisticated, Lifting Body series 

of unpowered gliding vehicles. Several changes were made to the 

rigging arrangement and control modifications were tested but few 

responses were effective and none were predictable. On the fifth 

flight aloft, Peterson got out of phase with the control lag and a 

sinusoidal wallowing motion set in, but at the moment the towtruck began to slow down the Parasev flipped over and crashed 

to the desert floor from a height of 10ft, just 10sec after lift-off. 

Peterson survived intact but the Parasev did not, a rebuild necessitating a new designation – and a new, Dacron-covered Parawing!37

Even so, NASA believed it had a promising concept and persisted in its 

development. But NASA determined by May 1963 that the Parasev technology would not be pursued any further for Gemini. Instead, Gemini leaders 

 37.

103
Coming Home

thought the Parasev “a basic research vehicle and not directly applicable to 

Gemini paraglider landing system program.”38

As one of several design and test-development articles used to evaluate the 

paraglider concept, NASA built a half-scale test vehicle (HSTV) for dropping 

out of a helicopter and testing the paraglider. The first test took place at NASA’s 

Flight Research Center in the Mojave Desert on December 10, 1962. It was 

an inauspicious performance. As the test report stated:

The test vehicle was launched at approximately 1400 PST from 

12,300 foot pressure altitude (indicated) and a velocity of 51 

knots indicated airspeed which was in accordance with the launch 

conditions specified in the Detailed Flight Plan.... Immediately 

following release, the test vehicle tumbled into the riser lines of 

the drogue parachute. This tumbling apparently resulted from the 

break-ties used for stabilization during pre-launch flight.

Shortly after the drogue chute was untangled the HSTV went 

into a spinning condition (describing a cone).... The capsule continued to spin throughout Paraglider pressurization, aft restraint 

release and apex release. This spinning condition prompted the 

decision to jettison the Paraglider and deploy the parachute.39

The HSTV landed safely about 120 seconds after deployment, but the test 

demonstrated one of the significant problems of the paraglider throughout 

its R&D experience. The next test, on January 7, 1963, had to be aborted 

because of radio-command system failures. They tried again the next day, this 

time with disastrous results. “At an estimated 1000 feet from the ground, and 

the absence of the main chute, the emergency chute signal was transmitted 

from the Comm. Van with no effect,” the test report states. “The capsule was 

damaged severely and beyond repair.”40 Tests followed on May 22 and June 3, 

 38.

104
Landing Under Canopies

1963, but during a July 30 drop, both the main and backup parachutes failed 

and the vehicle was destroyed.41

In the end, the HSTV flights demonstrated a pattern that was replicated 

throughout the entire effort. The unfurling of the paraglider never seemed to 

proceed as intended and, sometimes, emergency systems failed to compensate. 

There was certainly no way NASA could put astronauts into a vehicle with 

such complications. NASA’s Gemini Project Manager James A. Chamberlin 

repeatedly called into question the entire effort, but he was dissuaded from 

cancelling it, at least for the time being. He told North American senior official Harrison A. “Stormy” Storms, Jr., on September 21, 1962, that there was 

“growing concern” over “repeated unsuccessful attempts...to conduct satisfactory predeployed half-scale paraglider tests.” Because of this, Chamberlin 

ordered a standdown for the test flights until North American could restructure 

its test program and give reasonable assurance that it could proceed effectively.42

Considerable changes to the system followed, and in October 1962 the 

final half-scale test flight achieved its main goal of demonstrating stability 

in the paraglider system during free flight.43 The next three flights suggested 

that the October 1962 test might have been a fluke. The first two of these 

were only partially successful; the third, on March 11, 1963, failed to deploy 

the wing, and then the emergency backup parachute failed as well. With this 

crash, two HSTV test articles were too much wreckage. The testing ended and 

Chamberlin, who had been skeptical about this for some time and wanted to 

end the paraglider program, lost his job and was replaced by a new Gemini 

program manager committed to the new landing concept.44 NASA reorganized 

the HSTV test effort once again, proposing a two-part test program in which a 

capsule would be deployed from an aircraft and then deployment of the wing 

would be tested in flight. Upon completing that successfully, North American 

and NASA personnel would tackle the second phase: landing the vehicle safely. 

 41.

105
Coming Home

NASA engineers recovering the Gemini paraglider after a test in the Mojave Desert. NASA.

106
Landing Under Canopies

That second task involved the use of two full-scale trainers towed to altitude and 

piloted down to the surface. Once this was mastered, according to the research 

proposal, NASA could move on to test the entire sequence.45

NASA never got to that stage. The half-scale tests ended much later than 

anticipated, and demonstrations of the ability to deploy the parawing in flight 

were not completed. It seemed there was virtually no way to get this system 

ready to fly for Gemini, even 2 years before the first orbital flights of the spacecraft. Some still wanted to cancel the effort, but North American pressed for its 

continuation, saying that the problems had been resolved, and the paraglider 

test program was granted a stay of execution.46

Throughout this tough cycle of R&D, few other than Chamberlin truly 

wanted to give up on the paraglider concept, and even Chamberlin would have 

preferred to make it work. Several people weighed in to preserve the R&D 

effort. For example, Major General Ben I. Funk, commander of the U.S. Air 

Force Space Systems Division, saw a need for the landing system and was willing to support it, despite it not being a military program. “I would not like 

to see the paraglider effort dropped,” he wrote to Manned Spacecraft Center 

Director Robert R. Gilruth on September 16, 1963. “[F]or all the procedural 

and technical problems which had beset its development so far, the paraglider 

may still prove the best way of land-landing a low lift/drag spacecraft, a capability in which we are interested for possible future military operations.”47

To prove the second phase of the full-scale parawing recovery and landing 

system, NASA built an “advanced paraglider trainer” for use by astronauts 

preparing for Gemini flights. It was modified in 1963 to serve as the first of 

two TTVs. The first, TTV-1, was used in perfecting maneuvering, control, and 

landing techniques. It had wheeled landing gear, but NASA also considered 

the use of landing skids. In essence, this vehicle was a boilerplate of the Gemini 

capsule that had the same dimensions as the flight article. “The cockpit floor, 

side bulkheads and substructure below the floor were simulated to permit 

installation of Gemini landing gear at a later point in time. The frames and 

 45.

107
Coming Home

This flight test of a 50-foot parawing’s ability to bring down safely a model of a human space 

capsule from a few thousand feet took place on an old army bombing range near Langley Field, 

VA, in 1961. NASA L61-8041.

bulkheads were of welded tri-ten steel with nonstructural and structural doors 

of aluminum plate for access to internal equipment.”48 The vehicle always had 

difficulties with control, especially the “flare” that allowed for a soft landing. 

For flight tests, TTV-1, and later TTV-2, flew with its paraglider wing already 

deployed; it was towed aloft and released. TTV-1 performed eight of these 

drop tests and landings at the Edwards dry lakebed.49

The first of these piloted TTV tests, albeit with the paraglider already 

unfurled, took place on July 29, 1964, when a helicopter towed TTV-1 with 

 48.

108
Landing Under Canopies

North American Aviation test pilot Charles “E.P.” Hetzel aboard. He released 

from the helicopter at approximately 10,000 feet and then undertook a descent 

to the ground that took more than 20 minutes. A second free flight, on August 

7, went well until at about 3,000 feet, when TTV-1 went into a hard left turn 

and headed for the ground; Hetzel bailed out and suffered modest injury 

on landing. TTV-1 was damaged during the test, and there were no further 

piloted flights of the vehicle until after 14 remotely controlled tests had been 

completed. A flight on December 19, 1964, also ended badly, with research 

pilot Donald F. McClusker flaring the paraglider late and suffering injuries 

in the hard landing. Persistence paid off, however, and eventually NASA was 

able to complete free flights on the TTV-1—but not before the program was 

divorced from the Gemini program.50

In addition to the piloted TTV effort, a nonpiloted, full-scale test vehicle 

(FSTV) underwent drop tests from the back of a C-130 aircraft. As the final 

report noted, this “vehicle was designed to withstand loads and stresses peculiar 

to the test trajectories with sufficient conservatism to preclude catastrophic 

failures during flight.”51 North American built two FSTVs and designed a set 

of drop tests to see how best to ensure the paraglider deployment sequence. At 

about 10,000 feet the test vehicle would then jettison the paraglider in favor of 

landing on the emergency parachute system.52 Twenty tests were scheduled, but 

the program conflicted with gravity almost from the beginning. The first test, 

on January 22, 1964, failed when the paraglider failed to deploy. Three more 

tests, February 8, March 6, and April 10, also failed to yield positive results. 

Then, after yet another failure of an April 22 test, NASA managers had no 

choice but to cancel the program. It took months to come to this decision, and 

Gemini program officials did not publicly announce the decision until August 

10, 1964, allowing North American to continue this effort as an R&D program 

unrelated to Project Gemini until the contract expired at the end of the year.53

The R&D program for the paraglider extended from May 1962 until 

December 1964. The latter date was just before the inauguration of piloted 

flights of the Gemini spacecraft, the next spring. Indeed, as the program proved 

less successful than originally envisioned, NASA engineers kept pushing back 

deployment of the paraglider, suggesting that the first few missions could use 

 50.

109
Coming Home

This technical drawing 

shows the configuration of 

the paraglider for stowage 

and flight tests. NASA.

110
Landing Under Canopies

conventional parachutes but later flights would incorporate the paraglider. At 

one point in 1964, NASA wanted to have the first seven Gemini capsules use 

a traditional parachute-recovery system and the last three missions employ 

the paraglider. This proved a pipe dream as well, and soon the program’s sole 

remaining goal was completion of a technology development effort with no 

relationship to Project Gemini. A synopsis of the test program was contained 

in the paraglider development effort’s final report:

The program incorporated a series of half-scale and full-scale 

wind tunnel tests, twenty-five deployment flight tests and four 

free-flight maneuver and landing tests. Ancillary testing included 

development of the FSTV parachute recovery system through 

boilerplate drop tests; half-scale free flight tests; one-tenth scale 

and one-twentieth scale drop tests.54

Test program officials concluded, somewhat ebulliently, and perhaps too 

generously, in the final analysis: “The flight test program was concluded with 

complete success in the FSTV and TTV phases.”55 In an irony of the first 

magnitude, after yet another failure of the program on April 22, 1964 (the fifth 

consecutive failure), NASA cut its losses and ended the paraglider program. 

On the day after NASA began phasing out the paraglider, April 30, 1964, the 

program finally had a successful test. Playing out the string must have lessened 

the pressure on the test team, as thereafter it conducted 19 more tests (out 

of 25 total), and in the summer it worked out the deployment sequence and 

began racking up successes, one after another.56

NASA’s Administrator for Manned Space Flight, George E. Mueller, tried to 

explain the strange career of the Gemini paraglider R&D program to Associate 

Administrator Robert C. Seamans in August 1965:

As you know, the Paraglider development program was instituted 

as a part of our mainstream Gemini effort in FY 1962 as a promising land-landing concept. The inflatable Rogallo wing configuration had been demonstrated in some small scale applications and 

the performance was certainly encouraging enough to warrant 

investigation for manned land-landing application. A very similar 

type of development approach to that followed for the Parasail 

 54.

111
Coming Home

was conducted for the paraglider, including a number of component developments, scale tests, and full-scale flight tests. This total 

effort in the mainstream Gemini program amounted to $27.4M 

(FY 1962 - $0.3M; FY 1963 - $12.1M; FY 1964 - $15M) with 

the bulk of the funds obligated for system development and flight 

test ($22.2M) and the remainder used for design and component 

development. Because of the inherent heavier weight encountered 

with the Paraglider, and the problems of spacecraft stowage and 

Paraglider deployment, considerable emphasis was placed throughout the development on the requirement for a light-weight configuration and a feasible and operationally suitable deployment 

sequence. This program culminated in a limited series of full-scale 

flight tests in December 1964, one of which was manned.57

Mueller concluded that the program, despite modest success toward its 

end, was only promising as a technology development program removed from 

the Gemini program.58 He let the program run its course, but as Seamans 

agreed: “the paraglider test program will phase out with the work presently 

under contract. This will be concluded by December of this year. Therefore, 

it is understood that there will be no requirement for further funding of 

paraglider activities.”59

A core question that one must ask about this program is why it failed. The 

first and most significant factor was the technological challenge. Deploying an 

inflatable structure from the capsule and gliding it to a landing on the surface 

is a task not without difficulties. North American was never able to overcome 

those difficulties. If the problem was not the deployment, it was the control 

mechanism. If it was not either of those, it was the difficulty in piloting the 

vehicle. If it was none of those, it was the size and weight of the paraglider in 

relation to the capacity of the Gemini capsule. Historian Barton C. Hacker 

concluded: “Paraglider’s failure clearly owed something to intransigent technology, something to limited human and capital resources.”60 As Dwayne Day 

 57.

112
Landing Under Canopies

concluded: “The paraglider at first seemed like a promising technology that 

could solve the problem of bringing a spacecraft down to a specific area and 

placing it gently on the ground. But the technology never matured despite 

considerable time and effort spent by NASA to perfect it.”61 Hacker also found 

fewer technical reasons for the lack of success, reasoning the program also failed 

because of “bureaucratic friction and internal politics, rigid schedules and competing demands for resources, clashing institutional and national priorities.” 

He documented in-fighting, bureaucratic politics, and a host of other issues 

that also affected the program.62

Another question that might be asked about the program is why did NASA 

project managers persist with the paraglider system as long as they did before 

admitting the technology was not yet ready for prime time. There are a number 

of possibilities that could help explain this seeming incongruity. One factor, 

without question, is that no one wants to admit defeat. A program had been 

approved, preliminary analysis suggested this was an achievable goal, and 

money was available to pursue the task to a successful conclusion. No engineer 

could have been better motivated. The paraglider represented a useful solution 

to an understandable problem in a program of national priority; the program 

was well funded, and there was a foreseeable solution. These characteristics 

helped attract the Gemini program staff, both NASA and contractor employed. 

Naysayers who thought this untested concept was too risky to the lives of the 

Gemini astronauts served to steel the supporters’ resolve.

There is one other factor that propelled the paraglider R&D effort: the desire 

to fly home like an airplane and land on a runway, or at least a skid strip, rather 

than bobbing in the ocean while waiting for the Navy to rescue the capsule 

and crew. Virtually all of the NASA engineers of the era came out of aviation 

engineering programs and cut their teeth on aerodynamics and hypersonics. 

Ultimately, people pursue design solutions that reflect their backgrounds and 

perspectives. A winged return to Earth had dominated the thinking about space 

flight from the 1920s until the ballistic era of the 1950s. The paraglider allowed 

an incremental step back toward what everyone believed should become the 

norm in space recovery. Given this historical tradition of winged space vehicles, 

the relatively small Mercury, Gemini, and Apollo spacecraft placed atop ballistic missiles seemed like anomalies, and reentry from space on parachutes 

seemed like an admission of defeat. Moreover, the primary goal of the landing 

system was pinpoint landings. A wing, or just aerodynamic lift (to include lifting bodies), was one way to achieve this goal. Rotors and lift rockets, like the 

 61.

113
Coming Home

lunar lander, were also possibilities, as were steerable gliding parachutes, Rogallo 

wings, and the parafoils advocated by this program. Each method worked in 

theory, but the weight, storage, and deployment of these options caused serious 

problems during flight tests. That it took so long to realize that the paraglider 

concept would not be ready for Gemini’s use is one of the more interesting 

aspects of the story. On some level, program participants wanted to believe they 

could be successful and move the concept to the stage of piloted space landing.

The well-respected aerodynamics researcher Bob Hoey commented on this 

problem for this study:

Engineers considered all options for achieving reusability and pinpoint return. Gene Love from Langley used to separate the concepts 

into two categories: (1) Coupled Mode – which included wings, 

lifting bodies, or other entry configurations which could be recovered, (landed) without deploying a special “landing” system. (2) 

Decoupled Mode – which included parachutes, rotors, landing 

rockets, or other special “landing” devices which were deployed after 

entry, and grossly altered the vehicle configuration and controllability as the vehicle neared the ground. Design engineers respond to 

the requirements that are laid on them, and are not responsible for 

misperceptions or errors in the requirement specifications.63

Historians of technology use the phrase “the social construction of technology” to symbolize decisions made not solely for engineering purposes but also 

for design preferences, usually resulting from psychological conceptions. At 

some level, the paraglider program seemed to have evolved from what its creators 

knew about the political, economic, social, and technological issues it addressed. 

It was, accordingly, a product of heterogeneous engineering, which recognizes 

that technological issues are simultaneously organizational, economic, social, 

and political. Such an outcome was understandable as people, institutions, and 

interests came together to launch the program. The final decision met most of 

the priorities brought to the process but left others unaddressed.64

 63.

114
Landing Under Canopies

As a plan B, NASA pursued this parasail concept, which also offered a modicum of control for 

landing. NASA.

115
Coming Home

As one observer, Valtteri Maja, asked rhetorically, “Why not use parachutes 

instead of parafoils?” The answer is, “You cannot really steer them much at 

all. There is very little lift and mostly just drag, which means that the descent 

speed is quite large compared to the chute size. And it also means that you 

can’t fly them.” Maja added:

Flying has numerous advantages compared to falling. You can do a 

flare move just before landing to make the landing very gentle. You 

might need skids or even wheels though in case the flare doesn’t 

stop you completely. And you can navigate when you fly. Precision 

skydivers routinely land their feet directly on a point target in competitions. If your ram-air parafoil has a lift to drag of 10, and you 

start gliding at 5 kilometer height, you can theoretically pick any 

landing point in a 50 km radius. You can land on a runway. You can 

have better L/D than the space shuttle, meaning an easier and more 

flexible and maneuverable landing. But I’m getting carried away, 

this post was about the last five kilometers only. The main point 

is that the parafoil is the capsule’s way out of the disadvantaged 

position of helpless passive falling and need for extensive recovery 

which has been the tradition so far. Capsules don’t have to be like 

that. The world has moved on since the sixties and offers wonderful 

opportunities. With the parafoil, capsules can fly.

65

Flying seems to have been the key here. What dominated the resolve with 

which NASA engineers pursued this concept, notwithstanding technical challenges that seemed insurmountable, ultimately was the belief that any recovery 

method that allowed astronauts to pilot their space vehicles to landings on 

Earth was superior to dangling from the bottom of parachutes and being 

rescued at sea by the U.S. Navy.66

 65.

116
Landing Under Canopies

Three Canopies into the Pacific: Project Apollo

With the failure of the paraglider program for Gemini, NASA approached 

recovery from space for the Apollo spacecraft using a parachute concept. Of 

course, the race to the Moon through Project Apollo led to sophisticated 

technological innovations in the relationship between humans and machines. 

While putting an American on the Moon in 1969 was a feat of admittedly 

astounding technological virtuosity, the codependent relationship of humans 

and the machinery of space flight, and the sophistication of mechanical systems 

under human control, accounted for the six successful Moon landings between 

1969 and 1972. Project Apollo succeeded in no small measure because it was a 

triumph of management in meeting enormously difficult systems-engineering 

and technological-integration requirements, as several historians have pointed 

out; but the machines controlled by the astronauts, more than anything else, 

went far toward making the process work. In this process, the Apollo astronauts 

proved their supremacy over all of the technological systems that were created 

to make voyages to the Moon possible. Indeed, they became the critical component in the effort, including the landing system.

For all of the effort on the Gemini’s Earth landing system—paraglider, 

parasail, or parachute—virtually nothing about the Apollo program revolved 

around its Earth landing system. Gone were the extravagant efforts to achieve 

a nonwater landing, as the Apollo astronauts would be rescued at sea after a 

parachute landing. The Apollo familiarization manual described the system 

used for this recovery as follows:

The C/M-ELS begins operation upon descending to approximately 

24,000 feet +0.4 second, or in the event of an abort, 0.4 second 

after launch escape assembly jettison.... The apex cover (forward 

heat shield) is jettisoned by four gas-pressure thrusters. This function 

is imperative, as the forward heat shield covers and protects the ELS 

parachutes up to this time. At 1.6 seconds later, the drogue mortar 

pyrotechnic cartridges are fired to deploy two drogue parachutes 

in a reefed condition. After 8 seconds, the reefing lines are severed 

by reefing line cutters and the drogue parachutes are fully opened. 

These stabilize the C/M in a blunt-end-forward attitude and provide 

deceleration. At approximately 10,000 feet, drogue parachutes are 

released, and the three pilot parachute mortars are fired. This action 

ejects the pilot parachutes which extract and deploy the three main 

parachutes. To preclude the possibility of parachute damage or failure 

due to the descent velocity, the main parachutes open to a reefed 

condition for 8 seconds to further decelerate the C/M. The three 

117
Coming Home

parachutes are then fully opened (disreefed) to lower the C/M at a 

predetermined descent rate. At 27-1/2-degree hang angle of the C/M 

is achieved by the main parachute attachment points. In the event 

one main parachute fails to open, any two parachutes will safely carry 

out the prescribed function.

The main parachutes are disconnected following impact. The 

recovery aids consists of an uprighting system, swimmers umbilical, sea (dye) marker, a flashing beacon light, a VHF recovery 

beacon transmitter, a VHF transceiver, and an H-F transceiver. 

A recovery loop is also provided on the C/M to facilitate lifting. If 

the command module enters the water and stabilizes in a stable 

II (inverted) condition, the uprighting system is activated (manually), inflating three air bags causing the command module to 

assume a stable I (upright) condition. Each bag has a separate 

switch for controlling inflation. The sea (dye) marker and swimmer’s umbilical are deployed automatically when the HF recovery 

antenna is deployed (manually initiated by crew). The marker 

is tethered to the C/M forward compartment deck and will last 

approximately 12 hours. The swimmer’s umbilical provides the 

electrical connection for communication between the crew in the 

C/M and the recovery personnel in the water.67

The spacecraft would land in the Pacific—for all of the lunar missions—

reaching the water at a velocity of “33 feet per second at 5,000 feet altitude for 

a normal or abort landing.”68

To develop the Apollo landing system, NASA contracted with North 

American Rockwell. Building on knowledge gained in the Gemini and Mercury 

parachute-landing systems, North American Rockwell undertook a rigorous 

and extensive design and testing regimen. Northrop engineer Theodor W. 

Knacke reported in 1968 the following:

Numerous interesting design details are contained in the Apollo 

parachute system. The reliability requirement of independent 

parachute deployment, coupled with large command module 

oscillations, necessitates divergent drogue parachute and main 

 67.

118
Landing Under Canopies

The Apollo 16 Command Module, with astronauts John W. Young, Thomas K. Mattingly II, and 

Charles M. Duke, Jr., aboard, nears splashdown in the central Pacific Ocean to successfully 

conclude a lunar landing mission. This overhead picture was taken from a recovery helicopter 

seconds before the spacecraft hit the water on April 27, 1972, with its three parachutes allowing 

a safe return of the capsule and crew near Christmas Island in the Pacific. NASA S72-36287.

pilot parachute deployment angles coupled with positive thruster 

type deployment. The command module oscillations create the 

possibility of contact between the parachute risers and the hot rear 

heat shield, and last but not least, the increase in CM weight without an accompanying increase in compartment volume or allowable parachute cluster loads resulted in novel design approaches 

for parachute packing, storage and shape retention.

Designed for use in both optimum and crisis situations, either during 

launch abort or return from the Moon, this system was fully redundant and 

handled forces equivalent to 3 g’s without difficulty. A study explains that as 

designed, the system should do the following:

Two ribbon drogue parachutes accomplish initial deceleration 

and stabilization, with only one parachute being required and 

119
Coming Home

the second parachute providing the back-up mode. Deploying 

both parachutes simultaneously eliminates the need for an emergency sensor, provides for faster CM stabilization and creates 

more favorable main parachute deployment conditions. After 

disconnect the two drogue parachutes are followed by three pilot 

parachute deployed Ringsail main parachutes; two of which will 

provide the rate of descent necessary for water landing.

The study concluded the following: “The basic design proved flexible 

enough to accept a substantial increase in command module weight and a 

resultant increase in recovery envelope and velocity of parachute deployment 

without changing parachute volume or load requirements.”69

Not all went well with every aspect of the Apollo parachute-recovery system. 

A number of tests failed during the runup to the missions to the Moon. For 

example, on September 6, 1963, Apollo command-module boilerplate no. 3 

was destroyed when one pilot parachute was cut by contact with the vehicle and 

one of its main parachutes did not deploy. Then, rigging problems caused the 

other two parachutes to fail. An investigation led to rigging and design changes 

on future systems. In this case, these difficulties were resolved and the program 

continued.70 The most serious failure came during the descent of Apollo 15 

from the Moon in 1971. During its reentry, all three main parachutes deployed 

without incident at an altitude of 10,000 feet, but one of the three parachutes 

deflated while the Apollo 15 capsule was obscured by clouds between 7,000 and 

6,000 feet. Regardless, the crew and spacecraft returned safely because redundancy in the system allowed one parachute to fail without compromising the 

entire landing system. However, short the one parachute, the crew descended 

to impact at a slightly higher velocity than planned. Failure analysis found that 

the parachute lines had been damaged by fuel from the reaction control system 

(RCS) during return, a normal occurrence, but in this instance, the parachute 

assembly was in the way of the RCS ejection ports. The Apollo mission summary 

reported only: “During the descent, one of the three main parachutes failed, but 

a safe landing was made.”71 As reported at the time: “The most probable cause 

of the anomaly was the burning of raw fuel (monomethyl hydrazine) being 

expelled during the latter portion of the depletion firing and this resulted in 

 69.

120
Landing Under Canopies

The Apollo 9 spacecraft, with astronauts James A. McDivitt, David R. Scott, and Russell L. 

Schweickart aboard, approaches splashdown in the Atlantic recovery area on March 13, 1969, only 

4.5 nautical miles from the prime recovery ship, USS Guadalcanal (LPH-7). NASA S69-27467.

121
Coming Home

exceeding the parachute-riser and suspension-line temperature limits.”72 Based 

on this anomaly and its occurrence frequency—only once in all of the missions 

to date—NASA investigators believed that there was only a 1-in-17,000 chance 

of failure on future missions.73

This basic approach worked well throughout Apollo, but earlier NASA engineers still wanted to have a nonwater landing system similar to that pursued 

for Gemini. Personnel at the Landing Technology Branch of NASA’s Manned 

Spacecraft Center tried to adapt the parasail landing system underway for 

Gemini to the Apollo Application Program, the follow-on to the Moon landings. The branch reported that it

expects to have a system that will be adaptable to Apollo. Their present effort is not aimed directly at incorporation of such a system, 

but rather at developing the technology and hardware necessary 

for the system itself. They are, however, basing their designs on a 

spacecraft that is of the CM size and type. The resulting system will 

most likely consist of a steerable parachute, plus some combination 

of landing rockets, deployable energy absorbers and stability aids.74

Its staff added that it had contracted with four organizations for various 

aspects of this effort:

A. Bendix Products Aerospace Division is developing a computer program to analyze land-landing dynamics.

B. North American Aviation is investigating landing gear systems for 

the Command Module.

C. Pioneer Parachute is developing a parasail type of steerable parachute.

D. Northrop Ventura is developing a cloverleaf type of steerable parachute.

The primary concern was whether this landing system could handle a 

14,000-pound capsule and be containable within a 1.5-cubic-meter space. 

While this system was considered even less heavy and bulky than a water 

system, the addition of landing rockets to cushion a landing might push the 

total weight above what was already envisioned for the Apollo Command 

Module. “Probable weight increase and cost of incorporation must be weighed 

against the added capability and decrease in cost of recovery operations,” the 

 72.

122
Landing Under Canopies

The Apollo 13 Command Module splashes down, and its three main parachutes collapse, as the 

week-long, trouble-plagued Apollo 13 mission comes to a premature but safe end. The spacecraft, 

with astronauts James A. Lovell, Jr., commander; John L. Swigert, Jr., Command Module pilot; and 

Fred W. Haise, Jr., Lunar Module pilot aboard, splashed down on April 17, 1970, in the South Pacific 

Ocean, only about 4 miles from the USS Iwo Jima (LPH-2) prime recovery ship. NASA S70-35644.

study concluded.75 This was the first reference in this recovery literature from 

the 1960s concerning the significant tradeoff engineers had to make between 

added weight and reduced-recovery operational cost. The Navy was generally 

quite agreeable during the space race era to deploying its ships for recovery, and 

NASA was not required to pay for that total operation. That made water recovery, at least from NASA’s perspective, not only the most expedient option, but 

also the least expensive method of recovery. Even so, this program concluded 

without adopting anything more sophisticated than the parachute system used 

for Mercury, Gemini, and Apollo. NASA would not return to the parasail/

paraglider concept for landing until the 1990s.

 75.

123
Wen Painter of NASA’s Flight Research Center drew this cartoon in 1966. It captured what some 

saw as a key difference between space capsule splashdowns at sea and spaceplane landings 

on a runway. NASA.

124
CHAPTER 4

While NASA’s space capsules presented a near-term capability for human 

space flight, few in the industry believed the vehicles were ideal. The future, 

as envisioned by almost everybody, included a lifting-reentry spaceplane that 

would land on a conventional runway. Even the capsule designers recognized 

the advantages of a lifting reentry vehicle, and Gemini and Apollo each had a 

small L/D ratio to allow more precise targeting of a landing site.1

During a lifting reentry, a vehicle generates lift perpendicular to its flightpath, which can then be adjusted to change both vertical motion and flight 

direction. Practical entry angles have an upper and lower limit. The lower 

limit, also called the overshoot boundary, is the angle at which the vehicle 

will skip back out of the atmosphere (essentially the normal angle proposed 

for skip-gliders). The upper limit, or undershoot boundary, is the load-factor 

limit established by vehicle-structural, human-tolerance, or operational considerations. The primary measure for lifting reentry is the hypersonic L/D coefficient.2

 Increasing the L/D has significant effects on decreasing the maximum 

entry load and increasing the allowable entry corridor depth. Low L/D values 

(0.5–1.0) produce durations essentially the same as a semiballistic reentry, 

with survivable g-loads, moderate heating levels, and low maneuverability. 

 1. Some purists might take exception to our use of the term “spaceplane” to characterize winged/

125
Coming Home

Increasing the hypersonic L/D above 1.0 brings only marginal improvements 

in entry loads, but extends the cross range as a nonlinear function of the L/D 

(an L/D of 1.5 might generate a 1,000-mile cross range, but an L/D of 3 will 

generate a 4,000-mile cross range). High L/D values produce low g-loads and 

allow significant variations in the flightpath, but they result in long-duration 

reentries with continuous heating. Although the peak temperatures of a lifting 

reentry are lower than those of a ballistic reentry, the total heat load (temperature multiplied by duration) is generally higher.3

Researchers expected that a spaceplane would land on a preselected runway, 

like an airplane, and would not require a large recovery force, like capsules. 

Theoretical and wind tunnel studies, however, showed the aerodynamic configurations that produced the highest L/D during entry did not necessarily produce 

a high L/D at landing speeds. Long slender cones or wedges without wings were 

ideal for entry, but the best landing configurations used long, gliderlike wings. 

Engineers soon discovered that by providing two flight modes—one for hypersonic velocities through the transition to subsonic flight, and another for landing—many of the flight-control problems of a spaceplane could be decoupled.

Since the majority entry occurs at hypersonic speeds, the engineers decided 

to concentrate on developing shapes that exhibited adequate stability and control at high speeds, and then they would add the equipment for the decoupled 

landing mode. During the late 1950s and early 1960s, the more prominent 

landing techniques (see chapter 2) included gliding parachutes, paragliders 

(Rogallo wings), rotors, sustained propulsive lift, and deployable wings (usually, incorrectly, called variable-geometry wings). However, it quickly became 

evident that this type of equipment greatly expanded the complexity of the 

vehicle and added a significant amount of weight. Researchers eventually found 

that the delta wing provided a compromise that allowed a single shape to perform well at hypersonic velocities and adequately during landing.

Unfortunately, thermal protection systems for a spaceplane were a great deal 

more challenging than for capsules due to the higher total heat load generated 

by the longer entry time. At the same time, the large surface of the delta wing 

further exacerbated matters. Needing to maintain a smooth aerodynamic surface added yet another complication.

In spite of any perceived drawbacks, lifting reentry spaceplanes capable 

of landing on a normal runway were considered necessary to increase human 

access to space at a reasonable cost, and became the Holy Grail of the aerospace 

 3. 

126
Spaceplane Fantasies

industry. At the time, NASA officials compared the space program’s traditional 

use of expendable launch vehicles with capsules atop to throwing away a railroad locomotive after every train trip, whereas a reusable spaceplane would 

offer cost-effective, routine access to space.4

Choices

In the immediate postwar era, three basic vehicle types were being considered 

for high-speed global flight: (1) ballistic, (2) boost-glide, and (3) skip-glide. 

Early in 1954, H. Julian Allen, Alfred J. Eggers, Jr., and Stanford E. Neice at 

the NACA Ames Aeronautical Laboratory put together a theoretical discussion of these different concepts in A Comparative Analysis of the Performance 

of Long-Range Hypervelocity Vehicles. In that report, the researchers examined 

the relative advantages and disadvantages of each vehicle type. The conclusions proved significant. They found that ballistic and boost-glide vehicles 

could probably be developed, with proper attention to design. However, even 

at this early date, the researchers warned that skip-gliders appear substantially 

less promising.5

The flight profile of each vehicle type is subtly different. The ballistic vehicle, 

best exemplified by the long-range missile, leaves the atmosphere and returns 

as a single maneuver, generally in the form of a parabola. A boost-glider is 

accelerated to a speed and altitude where the dynamic pressure allows the 

vehicle to glide at some given L/D coefficient. At this altitude, termed by Eugen 

Sänger the “equilibrium altitude,” the aerodynamic lift required for flight is 

less than the centrifugal force resulting from the curved flight around Earth.6

 4. Roger D. Launius, “NASA and the Decision to Build the Space Shuttle, 1969–72,” 

127
Coming Home

The required aerodynamic lift lessens as speed increases, and it becomes zero at 

orbital velocities. Surprisingly, the equilibrium altitude is fairly low—usually 

below 250,000 feet—until the speed approaches orbital velocity. Once at the 

equilibrium altitude, the vehicle maintains its maximum L/D until aerodynamic drag causes it to slow and lose altitude until it lands.7

On the other hand, the trajectory of a skip-glider (such as the Sänger 

Silverbird) is composed of a succession of ballistic paths, each connected to 

the next by a skipping phase during which the vehicle enters the atmosphere, 

negotiates a turn at some given L/D coefficient, and is ejected from the atmosphere by aerodynamic lift. Each upward skip results in a lower peak altitude 

than the last since the vehicle is unpowered, and this sequence continues until 

the vehicle no longer has sufficient energy to leave the sensible atmosphere, at 

which time it glides to a landing.8

While he was investigating the relative merits of the various vehicle concepts, Allen compared the likely structural weights of each vehicle. He noted 

that the apparent advantage of the ballistic missile—a lack of wings—was 

largely negated by the need for increased propellant tankage since the vehicle 

could not rely on aerodynamic lift at any point in its trajectory. The ballistic 

vehicle also experienced the highest aerodynamic loads while entering the 

atmosphere, adversely affecting structural weight (although the skip-glider 

suffered a similar flaw). In this regard, a boost-glider, or spaceplane, was the 

most efficient configuration in terms of structural weight, assuming a suitable, 

lightweight thermal protection system could be found.9

The researchers noted that the three vehicle types reacted to entry heating 

in vastly different ways. The blunt-body theory made it possible to design a 

ballistic reentry vehicle because the majority of the heat was carried away from 

the vehicle by shock waves and other aerodynamic phenomena; heat sinks or 

ablators could handle the residual heat. Obviously, this was impractical for both 

of the glide vehicles since, by definition, a glider needs a high L/D to increase 

its range and avoid plunging through the atmosphere.10

However, a boost-glider could gradually convert its kinetic energy over a 

longer time than a capsule, radiating much of the heat back into the atmosphere 

 7. Allen, “Hypersonic Flight and the Reentry Problem,” p. 6.

128
Spaceplane Fantasies

This early 1960s comparison of the relative merits of a parachute landing system versus a 

winged approach clearly favored flying back to a runway. NASA.

and maintaining comparatively low structural temperatures. Testing by Allen 

and Eggers revealed that with sufficiently low wing loading, it might be possible to build a glider that could radiate enough heat to maintain a structural 

temperature under 1,600 °F, which was within the capability of some available 

materials. The wing loading had to be kept low since increasing it adversely 

affected the radiation-equilibrium temperature, resulting in large wings and 

more thermal protection. Active cooling systems could be employed at the 

hot spots (nose, wing leading edge, etc.), where the equilibrium temperature 

exceeded the tolerance of available materials.11

The skip-glider, on the other hand, seemed to be unworkable. A large fraction of the kinetic energy was converted to heat in a short time during each skip, 

but the interval between skips was not sufficient to radiate the heat into space.12

 11.

129
Coming Home

Eggers determined that any skip-glider would need to use an extensive active 

cooling system, and the weight penalty would probably be so excessive as to rule 

this vehicle out as impractical, or even impossible, except for very short flights.

Despite having discovered the blunt-body theory that made ballistic reentry 

possible, Eggers became convinced of the overall desirability of a lifting reentry 

vehicle instead of a ballistic capsule. He revealed this preference at the June 

1957 annual meeting of the American Rocket Society in San Francisco, CA. 

Eggers was skeptical about the relatively high heating loads and deceleration 

forces characteristic of ballistic reentry, warning that “the g’s are sufficiently 

high to require that extreme care be given to the support of an occupant of a 

ballistic vehicle during atmospheric reentry.” He also pointed out that such an 

object, entering the atmosphere along a shallow trajectory to hold deceleration down to 7.5 g’s, would generate a surface temperature of at least 2,500 °F. 

Thus, in Eggers’s judgment, “the glide vehicle is generally better suited than the 

ballistic vehicle for manned flight.”13 He also saw the difficulty of recovering 

a ballistic capsule since it was not controllable in the atmosphere and might 

need a target area of several thousand square miles. Of course, several of these 

challenges were subsequently overcome by providing a small amount of lift, 

and developing better ablators, for the capsules.

Many concepts for accomplishing a lifting reentry followed by a horizontal 

landing were proposed during the late 1950s. There were inflatable vehicles 

with very low wing loading; delta-winged flattop shapes; delta-winged flatbottom shapes; semiballistic shapes with extendable wings; and several lifting 

bodies. A team at NACA Langley that included Charles H. McLellan, who 

had solved many of the aero-thermal issues associated with the X-15, determined in 1959, “in many ways body shape is of secondary importance to other 

design parameters such as size, weight, planform loading, and lift-drag ratio. 

Furthermore, for a given lift-drag ratio and weight-to-base-area ratio, no specific body shape was found to possess sufficiently superior qualities to exclude 

consideration of other body shapes. However, the configurations with more 

nearly flat bottoms appeared to offer some advantages from the standpoint of 

performance and heat-transfer considerations.”14

 13.

130
Spaceplane Fantasies

Lifting Bodies

Other researchers realized that the blunt-body shape could be modified into 

a lifting body that would generate a modest 1.5 hypersonic L/D and enable 

potential landing sites to be several hundred miles cross range. Even this relatively low L/D would be a great improvement over the capsules: Mercury 

had essentially zero L/D, but the slight refinements to Gemini raised this to 

about 0.25 at hypersonic velocities. Apollo is generally credited with an L/D 

between 0.6 and 0.8, but the original specifications only required 0.4 to provide 

adequate margins for guidance errors during super-orbital reentries (i.e., direct 

return from the Moon). However, the guidance techniques actually used during 

Apollo were sufficient. Therefore, most of the L/D was not required, and the 

flown missions used an L/D between 0.29 and 0.31, subjecting the crew to a 

deceleration of about 6 g’s.15

Apparently, Sänger had first proposed the concept of a lifting body as early as 

1933.Sänger argued that a flat-bottom lifting body would have a higher L/D at 

high supersonic speeds than a normal-shape aircraft of similar size.16 However, 

he did not pursue this concept to the extent of determining the optimum shape 

for a lifting body, nor did he prove that such a vehicle should necessarily have 

a flat bottom—although all of the Silverbird designs did feature flat bottoms.

In 1954, Meyer M. Resnikoff at NACA Ames began looking at the earlier work by Sänger and determined that a lifting body would enjoy a 40- 

to 100-percent advantage in L/D over a conventional aircraft at hypersonic 

velocities. Lacking access to a hypersonic wind tunnel, the study was based 

instead on formulas derived from Sir Isaac Newton’s impact theory.17 Resnikoff 

concluded that “the lower surface of such a body must be flat [thus verifying Sänger’s speculation] and rectangular, and that if the maximum available 

 15.

131
Coming Home

volume is utilized, the minimum drag body...is a wedge.”18 It should be noted 

that the lifting bodies investigated by Resnikoff bore little resemblance to most 

of the vehicles ultimately flown during the 1960s. Other researchers, such 

as Dr. Wilbur L. Hankey, Jr., at the Air Force Flight Dynamics Laboratory, 

reached similar conclusions around the same time.19

A different lifting body concept was presented in a paper presented by 

Ames’s Thomas J. Wong at the same March 1958 conference at which Maxime 

A. Faget detailed his blunt-body capsule design. Wong showed a 30-degree 

half-cone with an L/D high enough to limit deceleration during reentry to 

approximately 2 g’s and to allow a lateral cross range of ±230 miles, with a 

longitudinal variation of 700 miles. Unfortunately, much like the winged-craft 

proposal presented by NACA Langley Research Center’s John V. Becker at the 

same conference, this vehicle weighed substantially more than a pure bluntbody shape, and the throw-weight of the Redstone and Atlas boosters being 

used for Project Mercury did not allow such luxuries.20

Nevertheless, the lifting body concept interested the Air Force, which subsequently funded several small research programs, including one with The 

Aerospace Corporation that would enjoy sustained support over many years. 

The Air Force was interested in a maneuverable intercontinental ballistic missile 

(ICBM) warhead that could reduce the effectiveness of any antiballistic missile 

system, and the ability to select a landing site away from the ground track of a 

photoreconnaissance satellite could allow more opportunities for a film-return 

vehicle to be recovered. This was particularly important to polar operations, 

where the natural rotation of Earth moves the launch site considerably cross 

range during even a short orbital flight. High cross range allows a vehicle to 

land at its launch site after a single orbit, a desired capability that would later 

drive much of the Space Shuttle configuration. Ultimately, the warhead community moved in an entirely different direction with the high-ballistic coefficient reentry vehicles.

There were other perceived benefits of the lifting body design. It appeared 

that for any given planform size (wing area), the lifting body offered a large 

internal volume compared to a capsule or wing-body shape. However, this 

would prove misleading because much of the volume was composed of areas 

 18.

132
Spaceplane Fantasies

Three lifting body vehicles at NASA’s Flight Research Center in California in 1969. Left to right: 

the X-24A, M2-F3, and HL-10. These aircraft helped pave the way for later spaceplane concepts. NASA EC69-2358.

with complex curves that did not lend themselves to being efficiently filled with 

normally square (cube) equipment. Even propellant tanks presented a challenge 

since they needed to be odd shapes—often a difficult proposition with cryogenic 

or high-pressure storage vessels. Ultimately, packaging equipment effectively 

into a lifting body proved challenging both conceptually and practically.

By the end of 1958, four very different lifting body shapes were being seriously studied, including the Ames M1, the Langley HL-10, the Langley lenticular vehicle, and The Aerospace Corporation’s A3. Several other shapes were 

also under evaluation at the Air Force Flight Dynamics Laboratory (AFFDL) 

at Wright-Patterson AFB in Dayton, OH.21 Each represented a distinctive 

design approach.

Conceived largely by Eggers at NACA Ames, the M1 began as a modified 

13-degree half-cone, mostly flat on the top, with a rounded nose to reduce 

heating. The shape had a hypersonic L/D of 0.5 (only marginally better than 

a ballistic blunt-body capsule) and demonstrated a pronounced tendency to 

tumble end over end at subsonic speeds. In fact, the shape had virtually no 

subsonic L/D and could not land horizontally. Eventually, the Ames researchers 

Eggers, Clarence A. Syvertson, George G. Edwards, and George C. Kenyon 

 21.

133
Coming Home

discovered that most of the stability problems could be cured by modifying 

the aft end with body flaps that looked much like a badminton shuttlecock, 

creating the M1-L. The slightly improved L/D would give the vehicle about 

200 miles of lateral maneuverability during reentry and about 800 miles of 

longitudinal discretion over its landing point.22

This design gradually morphed into a blunted 26-degree cone that offered 

improved stability, and it became the M2 shape. Further flattening of the cone 

top provided additional lift and a hypersonic L/D of 1.4. The base area was 

reduced by adding a boat tail (first proposed in 1960 by Edwards and David H. 

Dennis at Ames) to the upper and lower surfaces of the half-cone, simultaneously 

improving the reentry trim capability and reducing drag. Soon, a protruding 

canopy, twin vertical stabilizers, and various control surfaces were added, leading 

to it being called the M2b “Cadillac.” Preliminary wind tunnel tests showed that 

this configuration would have a subsonic L/D of about 3.5 and could probably 

be landed safely if an adequate control system could be developed.23

At NACA Langley, a different shape was conceived under the general guidance of Eugene S. Love, John W. Paulson, Jr., and Robert W. Rainey. It was 

based on studies initiated in 1957 that showed a flat-bottom reentry configuration with negative camber (cross section like an inverted-wing airfoil) would 

have a 14-percent higher L/D than a blunt half-cone. Sharply upswept tips 

provided control with a single centrally mounted vertical stabilizer. Langley 

 22.

134
Spaceplane Fantasies

continued developing this concept and, in 1962, the configuration was designated HL-10 (HL stands for horizontal lander).24

Researchers at Langley also proposed a lenticular shape. Donna Reed, the 

wife of Robert Dale Reed, who will play a major role in this story a bit later, 

called this a powder puff; others called it a flying saucer. The design was a circular shape that had no noticeable protuberances during entry. It transitioned to 

horizontal flight by extending control surfaces after following a reentry profile 

similar to a symmetrical capsule. Although various artists’ concepts show the 

shape, and it is briefly mentioned in several technical reports, the only serious 

investigation of it appears to have been during the development of the Convair 

Pye Wacket defensive missile for the B-70 (WS-110A) bomber.25

A team led by Frederick Raymes at The Aerospace Corporation developed 

the original A3 shape. The design was subsequently transferred to the Glenn 

L. Martin Company, where engineers, under the direction of Hans Multhopp, 

found they could improve its L/D at both hypersonic and subsonic speeds by 

using a more slender body and a flatter lower surface. After consulting with 

engineers at the AFFDL, the Martin engineers introduced positive camber in 

the longitudinal axis (the opposite of that used on the HL-10). John Rickey, 

an aerodynamicist at Martin, was assigned to refine the configuration into a 

practical design as part of Air Force Project M-103. The resulting A3-4 or SV-5 

(space vehicle-5) shape had a severe delta planform with pronounced rounding 

and twin vertical stabilizers. It appears that the shape was specifically tailored 

to meet anticipated Air Force cross-range requirements for the CORONA 

film-return vehicles.26

Wind tunnel tests conducted on the SV-5 shape led to several changes to 

enhance the transonic characteristics, including slightly altering the angle of 

 24.

135
Coming Home

NASA engineer Dale Reed holds a model of the M2-F1 lifting body with the full-scale version 

directly behind him in 1967. In support of the M2 lifting body program in the early 1960s, Reed 

had built a number of small lifting body shapes and drop-tested them from a radio-controlled 

mothership. NASA E-16475.

136
Spaceplane Fantasies

the nose ramp to provide better high- and low-speed trim capability. As finalized in late 1963, the configuration incorporated both a center-fin and tip-fin 

airfoil that included inward camber.27

The concept of lifting reentry was taken seriously enough that a Joint Air 

Force–NASA Conference on Lifting Manned Hypervelocity and Reentry 

Vehicles was held at NASA Langley on April 11 to 14, 1960. The goal was 

to disseminate research results relating to piloted hypervelocity and reentry 

vehicles, which have lifting capability and the ability to maneuver in the atmosphere. Papers were presented that covered many aspects of vehicle design, wind 

tunnel and theoretical testing, flutter, guidance and control, heat transfer, and 

other issues of interest to vehicle designers. The results from hypersonic wind 

tunnel and hypervelocity-facility tests allowed researchers to predict that each 

of the lifting bodies had acceptable hypersonic characteristics.28

Air Force–Sponsored Research

In 1959, the Air Force evaluated the high-speed characteristics of a lifting 

body during three test flights that were part of a larger research effort called 

WS-199. With significant involvement from the AFFDL, McDonnell developed the Model 122B Alpha Draco boost-glider as part of WS-199D. Each 

lifting body–shaped glider was launched to 92,000 feet by a two-stage launch 

vehicle derived from the Sergeant Battlefield missile. After being accelerated 

to Mach 5, the Alpha Draco vehicle glided toward a preprogrammed location 

about 240 miles downrange, where it entered a terminal dive into the ocean. 

The first two Alpha Draco flights on February 16 and March 16, 1959, were 

successful, but the third flight, on April 27, 1959, did not fare as well and was 

destroyed by range safety after it deviated from its planned flightpath. Despite 

the failure, the program verified the basic principles of boost-glide vehicles, 

and provided limited data on hypersonic aerodynamics and thermodynamics.29

By August 1959, engineers at the AFFDL believed the advent of new hightemperature materials and sophisticated guidance packages would permit 

 27.

137
Coming Home

testing hypersonic gliders during actual reentry flights. The laboratory originally envisioned the gliders as simple wing-body vehicles with a pronounced 

keel on the ventral surface. However, Alfred C. Draper, Jr., recognized that 

the keel design would introduce a dihedral effect and result in serious heating 

problems during reentry. Draper successfully argued that the gliders should 

instead use the WLB-1 shape that had recently been developed by the laboratory. Ultimately, however, the final configuration was essentially the forward 4 

feet of the Dyna-Soar glider, with a flat-bottom delta planform, rounded nose 

cap, rounded wing leading edges, and a tilted nose for hypersonic trim. The 

leading edges, lower surfaces, and upper surfaces were of the same design and 

materials as Dyna-Soar, allowing designers to take advantage of the large body 

of wind and shock tunnel data that was being accumulated by Dyna-Soar.30

On January 31, 1961, the effort was designated Project 1466 and named 

Aerothermodynamic/Elastic Structural Systems Environmental Tests (ASSET). 

The flight program was designed to assess the applicability and accuracy of 

analytical methods and experimental techniques in the areas of structures, 

aerodynamics, aerothermodynamics, and aerothermoelasticity for lifting reentry vehicles. In April 1961, the Air Force awarded a contract to McDonnell 

for two slightly different glider configurations: aerothermodynamic structural 

vehicle (ASV) and aerothermoelastic vehicle (AEV). McDonnell built four 

of the former and two of the latter. The gliders were 5.7 feet long, weighed 

1,130 pounds (ASV) and 1,225 pounds (AEV), and had a sharply swept (72.5 

degrees) low-aspect-ratio delta wing. The anticipated Blue Scout boosters were 

soon replaced by Thor launch vehicles, which were more readily available.31

Although superficially similar, and sharing common subsystems, the two 

types of gliders differed completely in mission and research capabilities, and 

 30.

138
Spaceplane Fantasies

Dale Reed, who inaugurated the lifting body flight research at NASA’s Flight Research Center 

(later, Dryden Flight Research Center, Edwards, CA), originally proposed that three wooden outer 

shells be built. These would then be attached to the single internal steel structure. The three 

shapes were (from viewer’s left to right) the M2-F1, the M1-L, and a lenticular shape. Milton O. 

Thompson, who supported Reed’s advocacy for a lifting body research project, recommended 

that only the M2-F1 shell be built, believing that the M1-L shape was “too radical” while the 

lenticular one was “too exotic.” Although the lenticular shape was often likened to that of a flying 

saucer, Reed’s wife, Donna, called it the “powder puff.” NASA EC62-175.

this was reflected in their flight profiles. The ASVs were to evaluate materials 

and structural concepts and to measure temperature, heat flux, and pressure 

distribution during hypersonic gliding reentry. The ASVs were boosted to 

altitudes of 190,000 to 225,000 feet and velocities from 16,000 to 19,500 ft/

sec with ranges varying from 1,000 to 2,300 miles. On the other hand, the 

AEVs were boosted to 168,000 and 187,000 feet and a velocity of 13,000 ft/

sec, obtaining ranges from 620 to 830 miles. The primary AEV experiments 

were a 1-by-2-foot body flap on the rear body surface to evaluate unsteady 

aerodynamic effects on control surfaces and a corrugated columbium panel 

located ahead of the flap to investigate aeroelastic panel flutter.32

 32.

139
Coming Home

The first launch (ASV-1) took place on September 18, 1963, just a few 

months prior to the cancellation of Dyna-Soar. ASSET provided the first realworld reentry exposure for a new generation of refractory materials (like carbongraphite, thoria-tungsten, and zirconium) and superalloys such as columbium, 

the cobalt-based alloy L-605, and TZM molybdenum). Advanced thermalinsulating materials were also tested. Bell, Boeing, Martin, McDonnell, Solar, 

and Vought all provided panels or nose caps that were used on the gliders. 

Through these tests, ASSET provided a wealth of data that contributed to the 

development of more advanced materials (primarily composites and carboncarbon) that would be available when Space Shuttle development began later 

in the decade.33

The ASSET program furnished a great deal of information on the aerothermodynamics and aeroelastic characteristics of a flat-bottom wing-body shape 

reentering from near-orbital velocities. From 1966 to 1967, the Air Force tested a 

lifting reentry vehicle under similar conditions. Unlike the structures and heating 

research-oriented ASSET vehicles, PRIME explored the problems of maneuvering reentry, including pronounced cross-range maneuvers. ASSET, PRIME, 

and PILOT (a low-speed piloted demonstrator) were all part of Program 680A, 

START (Spacecraft Technology and Advanced Reentry Tests).34

During late 1964, the Air Force selected the SV-5D for further development, creating a subtle variation of the basic SV-5 lifting body for the PRIME 

reentry vehicle. The research and development on this system consumed over 

2,000,000 engineering work hours, including configuration and material studies, wind and shock tunnel work, and 50 low-speed glide flights using a recoverable model launched from a ballute.35

The PRIME vehicle (often reported as the X-23A, although this appears to 

be unofficial) was an 890-pound lifting body constructed primarily of 2014-

T6 titanium alloy, with some beryllium, stainless steel, and aluminum used 

where appropriate. The vehicle was completely covered with a nonreceding, 

charring ablative heat shield consisting of ESA-3560HF for the larger flat 

surface areas and a more robust ESA-5500M3 for the leading edges. These 

 33.

140
Spaceplane Fantasies

This 1960 concept from General Electric for an Apollo spacecraft bore a striking resemblance to 

the ASV-1 spaceplane developed for the ASSET program. NASA.

Martin-developed ablators used silica-carbon fibers and siloxane resin interlaced through a silicone-based honeycomb to hold the char. The thickness of 

the ablators varied between 0.8 and 2.75 inches, depending on local heating 

conditions. The nose cap was constructed of carbon-phenolic composite.36

The PRIME flights would terminate at approximately Mach 2 with the 

deployment of a drogue ballute. As the drogue ballute deployed, its cable would 

slice through the upper structure of the main equipment compartment, where 

a 47-foot-diameter recovery parachute was stored. Once the recovery chute was 

deployed, the SV-5D hung in a tail-down attitude awaiting aerial-retrieval by 

a Lockheed JC-130B Hercules.37

On December 21, 1966, the first PRIME vehicle (FV-1) was launched from 

Vandenberg AFB, CA, on a trajectory simulating a reentry from low-Earth orbit 

 36.

141
Coming Home

with a zero cross-range maneuver. The performance of the Atlas launch vehicle 

and the PRIME spacecraft was nominal, except for a failure of the recovery parachute. The vehicle fell into the Pacific Ocean and was lost. However, all flight 

objectives other than recovery had been accomplished, and over 90 percent of 

all possible telemetry had been received.38 On March 5, 1967, the second vehicle 

successfully completed a 654-mile cross-range maneuver during reentry, the first 

for a returning spacecraft. Several stringers failed to be cut during the parachute 

separation process, and the vehicle was suspended in a manner that the JC-130B 

could not recover; the second vehicle was lost in the Pacific Ocean, as well.39

On April 19, the third SV-5D completed a 710-mile cross-range maneuver 

and achieved a hypersonic L/D of 1.0 at velocities in excess of Mach 25, only 

slightly less than the 1.3 predicted. The performance of the PRIME vehicle 

and all of its subsystems was perfect, and this time everything worked well 

for the recovery. The waiting JC-130B successfully snagged the SV-5D at 

12,000 feet, less than 5 miles from its preselected recovery site. A complete 

postflight inspection revealed that the vehicle was in satisfactory shape and 

could be launched again if needed. This was an important, if somewhat 

unheralded, milestone in demonstrating the potential reusability of lifting 

reentry spacecraft.40 Satisfied with the results of the first three flights, the Air 

Force cancelled the last PRIME flight.

It would be difficult to overestimate the importance of the ASSET and 

PRIME tests to the development of the Space Shuttle. These nine suborbital 

tests, along with the hypersonic database generated by the X-15, provided a 

disproportionate amount of the actual aerothermodynamic data on which the 

Shuttle orbiter designs were based.

Piloted Lifting Bodies

In the early 1960s Robert Dale Reed, an aerospace engineer at the NASA 

Flight Research Center, near Edwards AFB, had been following the development of the lifting bodies with considerable interest. Reed noted that while 

the hypersonic flying qualities of the designs were no longer in question, there 

was still considerable doubt concerning their low-speed stability. In February 

 38.

142
Spaceplane Fantasies

1962, Reed built a 24-inch model of the M2 shape and launched it from a 

60-inch-wingspan, radio-controlled carrier aircraft. During several of the drop 

tests, his wife, Donna, used an 8-millimeter home movie camera to record the 

flights, which were later shown to Eggers and the Flight Research Center director, Paul F. Bikle. The results were encouraging enough for Bikle to authorize 

a 6-month feasibility study of a lightweight piloted M2 glider, construction 

of which was subsequently authorized in September 1962.41 The M2 shape 

was the first shape to be tested, but Reed anticipated testing the M1-L and 

lenticular shapes as well. This plan eventually was dropped, and the program 

went straight into the mission weight phase having only tested the M2-F1.42

The size of the M2-F1 was dictated by the desire to have a wing loading of 9 

pounds per square foot, resulting in a vehicle that was 20 feet long, 10 feet high, 

and 14 feet wide. 43 Complete with its pilot, the M2-F1 weighed 1,138 pounds.44

Following ground checkout of the control system, the M2-F1 was trucked to 

NASA Ames, where it was tested in the 40-by-80-foot, full-scale wind tunnel. 

On many of the wind tunnel runs, which ultimately totaled about 80 hours, 

NASA test pilot Milton O. Thompson was seated in the M2-F1, operating the 

controls while data was collected. The wind tunnel tests at Ames were completed 

in March 1963, and the M2-F1 was trucked back to the Flight Research Center. 

The first of many ground tows behind various trucks took place on April 5, 

1963. These were intended to lead to captive flights using a Parasev-like canopy, 

but none of the vehicles proved fast enough to get the M2-F1 airborne.45

In response, the Flight Research Center purchased a stripped-down Pontiac 

convertible with a 455-cubic-inch engine, four-barrel carburetor, and a fourspeed stick shift. Ninety-three car tows were performed before the first air tow 

 41.

143
Coming Home

was attempted.46 Air tows were accomplished behind a Douglas R4D (Navy 

version of the C-47/DC-3), and the first piloted free flight was on August 16, 

1963, with Thompson at the controls. On September 3, 1963, the glider was 

unveiled to the news media, where it immediately became a hot item in the 

popular press. The M2-F1 completed over 100 flights and 395 ground tows 

before being retired on August 18, 1966.47

By early 1963, preliminary studies were underway on an air-launched, 

rocket-powered lifting body built largely with systems and equipment left over 

from the X-1 and X-15 programs. Engineers intended to expand the envelope 

of lifting body research into the low-supersonic and transonic speed regions, 

and to evaluate the landing behavior of a mission-weight lifting body.48

In early 1964, Paul Bikle, Dale Reed, and Milt Thompson proposed building two mission-weight versions of the M2 that would be launched by the 

NB-52 carrier aircraft used by the X-15 program at the Flight Research Center. 

Officials at NASA Headquarters suggested, however, that the Langley HL-10 

shape also be tested. Accordingly, on June 2, 1964, the Flight Research Center 

awarded a fixed-price contract to the NorAir Division of Northrop.49

Interestingly, designers made a conscious decision to size the M2-F2 for a 

possible future orbital flight-test series. The diameter at the base of the original 

cone was intentionally made the same diameter as the upper stage of a Titan 

launch vehicle. It was hoped that by constraining the size of the glider to the 

mold lines of the booster, the destabilizing influence of a lifting shape could 

be minimized. This was, at least partly, in response to problems experienced 

with Dyna-Soar in which wind tunnel tests showed that the glider had an 

adverse effect on the launch vehicle. Of course, the M2-F2 was not designed 

to perform an actual reentry, but it was hoped that an identically sized orbital 

vehicle might be built in the future.50

M2-F2

The M2-F2 was rolled out of the Northrop plant on June 15, 1965, and was 

trucked to Edwards the next day. Of conventional aluminum construction, the 

 46.

144
Spaceplane Fantasies

The M2-F1 lifting body under tow at the NASA Flight Research Center. The wingless, lifting body 

design was initially conceived as a means of landing a spaceplane horizontally after atmospheric 

reentry. The absence of wings meant that less surface area would need to be protected from reentry 

heating. In 1962, NASA undertook a program to build a lightweight, unpowered lifting body as a 

prototype to flight-test the wingless concept. It would look like a “flying bathtub” and was designated 

the M2-F1, the “M” referring to “manned” and “F” referring to “flight” version. NASA ECN-408.

M2-F2 was 22 feet long, spanned 9.6 feet, and weighed 4,630 pounds without 

its single Reaction Motors XLR11 rocket engine. A full-span ventral flap controlled pitch, while split dorsal flaps controlled roll (lateral) motion through 

differential operation and pitch trim through symmetrical operation. Twin 

ventral flaps provided directional (yaw) control and acted as speed brakes.51

On March 23, 1966, the M2-F2 completed its first captive-flight under the 

wing of the NB-52. However, there was still concern that the lifting body might 

fly upward and impact the carrier aircraft after it was released, so two wind 

tunnel test series preceded the first glide flight. A subscale model was tested in 

the Langley 7-by-10-foot High Speed Tunnel to define the launch transients 

produced by the flow field of the NB-52 and to establish the proper carry angle 

for the pylon adapter. These tests predicted that an abrupt, controllable right 

roll would occur at launch but that the lifting body would fall downward. The 

second flight was a full-scale test using the M2-F2 in the Ames 40-by-80-foot 

wind tunnel test section. Data showed that the performance and stability were 

adequate to begin flight-testing.52

 51.

145
Coming Home

Jay L. King, Joseph D. Huxman, and Orion D. Billeter assist NASA research pilot Milton O. Thompson 

(on the ladder) into the cockpit of the M2-F2 lifting body at the NASA Flight Research Center. The 

M2-F2 is attached to the same NB-52 wing pylon used by the X-15 program. NASA EC66-1154.

NASA research pilot Thompson made the first unpowered free flight on 

July 12, 1966, and 12 additional glide flights were made in the following 

4 months. On the morning of May 10, 1967, during the 16th glide flight, 

NASA test pilot Bruce Peterson exited the second of two planned S-turns only 

to find the M2-F2 rolling and banking wildly. Peterson recovered control but 

was no longer lined up with the marked runway on the lakebed, and he was 

too low to make a correction. Peterson successfully completed the flare, and 

pulled the landing gear deployment handle just as the vehicle touched the 

lakebed. The M2-F2 rolled and tumbled end over end several times at more 

than 250 mph before coming to rest upside down on the lakebed.53 Peterson 

 53.

146
Spaceplane Fantasies

was seriously injured and the M2-F2 was essentially destroyed. Fortunately, 

Peterson made an excellent recovery, although he lost the use of one eye.54

Subsequent investigation revealed that settings within the control system 

allowed Peterson to trigger a severe pilot-induced oscillation during the final 

approach. Later theoretical analysis would show that this instability was the result 

of a coupled roll-spiral mode of the pilot-airplane combination.55 This was the 

unfortunate end of the M2-F2 flight program, and the wreckage of the M2-F2 

was transported to the Northrop facility in Hawthorne, CA, for inspection.

M2-F3

On January 28, 1969, NASA announced the crashed vehicle would be rebuilt 

as the M2-F3, with a large center vertical stabilizer, in addition to the upswept 

wingtips. The new stabilizer was not used for additional directional stability, 

but rather to prevent the strong yawing moment produced when the upper 

flaps were deflected asymmetrically for roll control.56

NASA test pilot William H. Dana took the rebuilt M2-F3 for its first glide 

flight on June 2, 1970, and found the handling qualities much improved. 

The M2-F3 made its fastest flight, Mach 1.6, on December 13, 1972, and its 

highest flight on December 20, when it reached 71,493 feet. The vehicle was 

retired after completing 43 flights—16 as the M2-F2 and 27 as the M2-F3.57

At one point, Northrop submitted an unsolicited $200 million proposal 

to build an orbital version of the M2 shape launchable by a Titan booster. 

Interestingly, neither NASA nor the Air Force took action, mainly because there 

were no funds, and the proposal faded from sight. However, while evaluating this 

concept, researchers noted that the degradation in performance and stability due 

to the roughened surface of an ablative heat shield after reentry was not factored 

into the lifting body flight-test program. Subsequent analysis showed that this 

 54.

147
Coming Home

NASA research pilot John A. Manke is seen here in front of the M2-F3 lifting body in 1972. 

NASA ECN-3448.

degradation was sufficient to raise serious questions about the true ability to land 

any of these vehicles following an actual reentry, and limited testing eventually 

took place using the HL-10 in the Ames full-scale wind tunnel.58

HL-10

Concurrently with the M2 flight tests, the Flight Research Center was also 

flying the HL-10. This vehicle was 22 feet long and spanned almost 15 feet 

 58.

148
Spaceplane Fantasies

across its aft fuselage and also was conceived with the intent of developing an 

orbital spaceplane concept. Internally, the M2-F3 and HL-10 were similar, 

with nearly identical subsystems and structural details. Unlike the other lifting 

bodies, the HL-10 did not incorporate a canopy, and the pilot was completely 

dependent on the nose and side windows for visibility during landing. The 

control system consisted of upper-body and outer stabilizer flaps for transonic 

and supersonic trim, blunt trailing-edge elevons, and a split rudder on the 

central vertical stabilizer. Since this was a test vehicle not intended for space 

flight, the thermal protection system was minimal.59

As with the M2-F2, subscale tests were conducted in the Langley 7-by10-foot High Speed Tunnel to define the launch transients produced by the 

flow field of the NB-52 and the proper carry angle for the pylon adapter. 

The HL-10 was rolled out in Hawthorne, CA, on January 18, 1966, and was 

subsequently trucked to the Ames 40-by-80-foot wind tunnel test section for 

final testing. The results were generally satisfactory, although some momentary 

flow separation was noted on the tip stabilizers under certain flight conditions. 

Engineers at Ames and Northrop did not consider this particularly serious and 

cleared the HL-10 for flight.60

Peterson took the HL-10 on its first glide flight on December 22, 1966, 

revealing that the flow separation was much more serious than initially thought, 

and NASA Langley embarked on yet another round of wind tunnel tests. The 

solution was to add an inward-cambered glove to the leading edge of the tip 

stabilizers that allowed the airflow to stay attached at high angles of attack and 

low speeds. The HL-10 was grounded for 15 months while the problem was 

identified and corrected.61

During its next flight, on March 15, 1968, research pilot Jerauld R. Gentry 

found the modifications worked well and the HL-10 handled nicely. On 

November 13, research pilot John A. Manke made the first successful powered 

flight, and on May 9, 1969, the HL-10 made the first supersonic flight of any 

of the piloted lifting bodies. Air Force test pilot Peter C. Hoag reached Mach 

1.86 on February 18, 1970, and 9 days later Dana topped out at 90,303 feet, 

 59.

149
Coming Home

HL-10 lifting body model in NASA Langley Research Center’s Full-Scale Tunnel in 1963. NASA 

L-1963-06125.

making the HL-10 the fastest and highest flying of the lifting bodies. It was on 

this flight that the HL-10 for the only time approached a situation in which 

the thermal protection system came into play in the slightest, but the reality 

was that this concern could wait for the development of future systems. The 

basic research program consisted of 35 flights, and the HL-10 was deemed the 

best performer of the lifting bodies. The shape later became the basis for several 

early Space Shuttle concepts.62

Two additional HL-10 flights were undertaken to determine whether turbojet engines were needed to land a spaceplane returning from space, or if it could 

be accomplished with a glider approach. Three Bell Aerosystems 500-pound 

force hydrogen peroxide engines replaced the XLR11 rocket. These engines 

fired as the vehicle passed through 6,500 feet on the way to landing, reducing 

 62.

150
Spaceplane Fantasies

Comparison of refractory and ablative shingle structures for a proposed orbital HL-10 thermal 

protection system. NASA.

the approach angle from the normal 18 degrees to 6 degrees and boosting the 

airspeed to over 350 mph. At 200 feet above the lakebed, the pilot would shut 

down the engines, extend the landing gear, and make a routine landing. Hoag 

made flights on June 11 and July 17, 1970, marking the last flight of the HL-10 

in testing this approach to landing.63

Hoag’s comments about the landing engines were quite negative. He 

reported that the shallow approach made it difficult for him to judge where 

the vehicle would touchdown. He also reported that the nose-high attitude 

during the entire maneuver forced the pilot to rely on the nose window, which 

provided poor depth perception, for a long time before touchdown. The consensus was that the weight and complexity of the engine installation and the 

increased pilot workload outweighed the benefits of landing engines—namely, 

the ability to make a missed approach, due to shallower descent angles and 

 63.

151
Coming Home

higher approach speeds.64 It was an important lesson for the future of spaceplanes, and accordingly NASA proceeded with plans for the Space Shuttle to 

perform a glider landing on a runway.

Air Force Efforts

Although heavily involved in supporting the M2-F2 and HL-10 programs, the 

Air Force also conducted independent research into lifting body concepts for 

reentry and recovery from space. In fact, as early as the mid-1950s, the AFFDL 

had begun investigating lifting body designs (SAMOS, SAINT, etc.) for a variety of purposes, including for use as a film-return vehicle for reconnaissance 

satellites. The laboratory conducted a detailed evaluation of a variety of shapes 

beginning in 1959 and continuing until the creation of X-24B. Most of these 

shapes tended to reflect the relatively high L/D preferred by researcher Draper.65

As early as 1962, Draper proposed flight-testing piloted prototypes of several of the designs, but the Air Force, much to the dismay of both Draper and 

NASA, declined. Nevertheless, Draper and his associates proceeded with their 

laboratory studies, generating a large series of tailored shapes. Some of these 

reentered the atmosphere as lifting bodies but deployed variable-geometry 

wings to increase their transonic L/D. The laboratory also explored interference configurations that used complex undersurface designs to position shock 

flows for favorable increases in lift.66 Similarly, the laboratory investigated 

Nonweiler caret-wing wave rider configurations that made use of favorable flow 

interference to increase their hypersonic L/D.67 This approach continued to be 

pursued during 1991 in support of the X-30 NASP, prior to its cancellation.68

 64.

152
Spaceplane Fantasies

X-24A

While NASA was testing the M2 and HL-10, the Air Force was building a 

piloted lifting body demonstrator. Project PILOT (piloted low-speed tests) 

was part of the ongoing START program, which also included the ASSET 

and PRIME vehicles. This would provide the last part of a knowledge base 

that would capture data for the A3-4/SV-5 shape from the Mach 25 achieved 

by the SV-5D (X-23A) all the way to approach and landing with the SV-5P 

(X-24A). The objectives of PILOT were to investigate static and dynamic 

stability characteristics and to verify control techniques for low-supersonic, 

transonic, and landing-speed regimes.69

Martin Marietta was awarded a contract on March 2, 1966, for a singlerocket-powered SV-5P, designated the X-24A. Construction of the vehicle 

consumed a little over 12 months under the direction of Bastian “Buzz” Hello 

and Lyman Josephs.70 When the X-24A was rolled out on August 3, 1967, it 

was 24 feet long, spanned 13 feet, and was essentially four times larger than the 

PRIME SV-5D shape, with a revised canopy area. Power was supplied by the 

seemingly irreplaceable XLR11. The vehicle had an empty weight of less than 

6,000 pounds, increasing to 11,000 pounds with propellants and a pilot. Since 

testing was to be limited to less than Mach 2, the X-24A was of conventional 

aluminum construction, with no special attention paid to heat protection.

The basic aerodynamic control system consisted of eight movable surfaces. 

Pitch control was derived from the symmetrical deflection of the lower and/or 

upper flaps, depending upon flight conditions. Differential deflection of the 

flaps provided the primary roll control. Pitch or roll commands, which caused 

either lower flap to close fully, resulted in the transfer of control to the corresponding upper flap through a simple clapper mechanism. The two pairs of 

rudders were deflected symmetrically as a bias feature with directional control 

provided by the deflection of the surfaces in unison.71

The X-24A was added to the existing cadre of lifting bodies (the M2-F1, 

M2-F2, HL-10) in the NASA Dryden Flight Research Center hangar, where 

they were being flight-tested by a joint Air Force/NASA test team. The X-24A 

arrived at Edwards on August 24, 1967, for initial ground testing, but the 

vehicle was soon airlifted to NASA Ames for testing in the 40-by-80-foot wind 

tunnel test section. While at Ames, additional tests were performed with a 

 69.

153
Coming Home

This 1968 photo shows the bulbous X-24A lifting body on the lakebed adjacent to the NASA 

Flight Research Center. The X-24 was one of a group of lifting bodies flown by the NASA Flight 

Research Center in a joint program with the U.S. Air Force at Edwards AFB from 1963 to 1975. 

NASA ECN-2006.

simulated rough ablator applied to the vehicle. Water-soluble glue was sprayed 

on the vehicle, and a wire-mesh screen was laid over the surface to simulate 

the pattern of the ablator honeycomb. Coarse sand was then sprayed over the 

surface. When the glue was dry, the wire mesh was removed, leaving a rough 

protruding pattern on the surface. It was hoped that these results could be 

correlated with the effects that were observed on the PRIME vehicle following 

an actual reentry. The results from the full-scale tunnel tests showed a 20-percent reduction in maximum L/D and some degradation in stability caused by 

surface roughness.72 As a result, any thoughts of flying and landing the vehicle 

with the simulated ablator roughness were quickly dismissed. Although this 

was an interesting data point, and used during the development of the Space 

 72.

154
Spaceplane Fantasies

Shuttle, no further work on it was done using the X-24A.73 However, it did 

raise doubts about the usefulness of an ablative heat shield on any operational 

lifting reentry vehicle that would be landed horizontally, and it eventually 

pointed R&D efforts toward the thermal protection system used on the Space 

Shuttle with its ceramic tiles and reinforced carbon-carbon.74

The X-24A arrived back at Edwards on March 15, 1968, but it was not 

until a year later that the vehicle was cleared for flight. Finally, on April 17, 

1969, Air Force test pilot Gentry made the first of nine glide flights. Gentry 

flew the first powered flight on March 19, 1970, and during the course of its 

test program, the X-24A made 28 flights, reaching a maximum speed of Mach 

1.60 and an altitude of 71,400 feet. The last portion of the X-24A program 

was dedicated to simulating Space Shuttle approaches and landings, and the 

vehicle completed its last flight on June 4, 1971.75

X-24B

While the X-24A was still being built, Martin also completed two jet-powered 

low-speed lifting bodies based on the SV-5 shape as a company-funded venture 

for use as astronaut trainers at the Aerospace Test Pilot School. However, the 

Air Force considered the vehicles significantly underpowered and they never 

flew. During 1968, the Air Force Systems Command solicited suggestions for 

possible uses for the two unused SV-5J airframes, and the AFFDL responded 

on January 23, 1969, with a proposal by Draper and Bill Zima to modify 

one of them into the FDL-7 configuration. To minimize modification costs, 

it was proposed that the basic SV-5J body structure and vertical stabilizers be 

retained, with the FDL-7 shape being gloved over the existing forebody, creating the FDL-8. A flat-bottom, 72-degree sweep, double-delta configuration 

contributed to a high hypersonic L/D, while a 3-degree nose ramp provided the 

proper hypersonic trim condition. Despite tailoring the airframe for hypersonic 

velocities, the X-24B was never intended to fly that fast.76

The laboratory saw this as an opportunity to perform a low-cost flight 

demonstration on one of their configurations with a hypersonic L/D of 2.5, 

in contrast to the other lifting bodies that had hypersonic L/Ds of between 

 73.

155
Coming Home

1.2 and 1.4. The proposed jet-powered FDL-8 was to be air-launched from 

one of the NB-52 carrier aircraft, but as the studies matured, the advantages 

of rocket propulsion became apparent, and Draper soon proposed using the 

X-24A instead of the SV-5J to reduce the costs of the program.77 When the 

X-24B (also known as the SV-5P-2) was rolled out on October 11, 1972, the 

vehicle had grown over 14 feet in length and 10 feet in span, its weight was up 

to 13,700 pounds, and the center of gravity was considerably farther forward 

relative to the main landing gear.78

The X-24B was delivered to Edwards on October 24, 1972, for several 

months of ground testing. NASA test pilot Manke made the first glide flight 

on August 1, 1973, and the first powered flight on November 15. On October 

25, 1974, Air Force test pilot Michael V. Love made the fastest X-24B flight 

at Mach 1.76, and Manke took the vehicle to 74,130 feet on May 22, 1975, 

marking its highest flight.79

By the summer of 1975, the Space Shuttle designers were again debating 

whether to provide air-breathing landing engines. The primary concern was 

whether low-L/D reentry shapes could successfully complete unpowered landings on conventional runways. From the beginning, engineers believed that the 

X-24B could accomplish this, and nine accuracy-landing flights showed that 

pilots could successfully touch down within 500 feet of a designated marker 

on the lakebed runway. Since the vehicle was traveling at about 220 mph 

at landing, the pilot was within the touchdown zone for only 3 seconds—a 

remarkable demonstration of energy management.80

Based on these results, research pilots Love and Manke proposed landing on the 15,000-foot concrete runway at Edwards, something that had not 

been done with the X-15 or earlier lifting bodies due primarily to their lack of 

ground-steering capability. Johnny Armstrong, the X-24B program manager, 

presented the plan to the commander of the Air Force Flight Test Center, 

Major General Robert A. Rushworth. The general was well acquainted with 

 77.

156
Spaceplane Fantasies

The X-24B flying over the lakebed at NASA Dryden Flight Research Center in 1975. The X-24B 

was the last of the lifting bodies to fly, after being modified from the original X-24A. The X-24B’s 

design evolved from a family of potential reentry shapes, each with higher lift-to-drag ratios, 

proposed by the Air Force Flight Dynamics Laboratory. NASA EC75-4642.

unpowered landings, having flown the X-15 more than any other pilot (34 

times). Armstrong remembers Rushworth was supportive but cautioned, “I will 

forgive you if you land long and end up rolling onto the lakebed, but landing 

short of the concrete runway will be unacceptable.”81

The desired touchdown point was a stripe painted on Runway 04, 

approximately 5,000 feet from the approach end. Manke performed the first 

runway landing on August 5, 1975, and landed in a slight bank with one main 

wheel touching before the stripe, and the other wheel touching after the stripe. 

Love made the second runway landing on August 20, 1975, touching down 

400 feet beyond the stripe. Both pilots reported that the additional visual cues 

provided by roads and Joshua trees made the final phase of the landing easier 

than on the smooth, flat lakebed. The two runway-landing rollouts were about 

34 percent longer than equivalent lakebed landings, and a similar difference 

 81.

157
Coming Home

was later noted between lakebed and runway landings of the Space Shuttle 

during flight tests.82

On September 23, 1975, Dana, who had been the last pilot to fly the X-15, 

also completed the last powered X-24B flight, marking the end of rocketpowered research flights at Edwards AFB.83 Following this flight, the Air Force 

decided to provide six checkout flights for three test pilots who had not had the 

opportunity to fly any of the lifting bodies. These flights included an unrelated 

test where access panels on the lower surface were used to evaluate Space Shuttle 

thermal protection system tiles. One array of tiles was bonded to the panel to 

assess the ability of the tiles to withstand aerodynamic shear forces in flight and 

to investigate the susceptibility to damage during landings on the dry lakebed. 

The last flight was on November 25, 1975.84 In all, the X-24B flight program 

had consisted of 24 powered and 12 glide flights.

Given the state of the art in 1970, the lifting bodies were not necessarily 

suitable as spacecraft. While the lifting body flight-test data was being gathered, 

the effects of ablation surface roughness on low-speed drag were also being 

evaluated, and full-scale wind tunnel tests of the X-24A with a simulated rough 

ablator surface showed a 20-percent reduction in L/D. Wind tunnel tests at 

Wright-Patterson AFB on an 8-percent model of the X-24A showed similar 

results. Flight tests of the X-15A-2, which used a thin ablative coating, showed 

a reduction in L/D of about 15 percent after a relatively mild exposure to the 

aerodynamic heating environment. Comparison tests of two PRIME vehicles, 

one before flight and one after flight, showed a 30-percent reduction in L/D. 

These effects were also accompanied by reductions in stability that could obviously be quite detrimental to the handling qualities.85

 82.

158
Spaceplane Fantasies

It must be concluded that the first three lifting body vehicles, as originally 

conceived, would probably not have been landable following a reentry with 

a normally ablated thermal protection system.86 Nevertheless, the lightweight 

ceramic-tile technology developed for the Space Shuttle opened the door to 

all of these lifting reentry concepts, and highly maneuverable entries with over 

2,400 miles cross range are theoretically possible with X-24B-like configurations. The lifting bodies, however, presented other challenges for designers, 

such as how to package propellant tanks and other subsystems into the oddly 

shaped fuselages.

Orbital Lifting Bodies

During the early and mid-1960s, the lifting body seemed like an ideal platform 

for an orbital vehicle, and, not surprisingly, several proposals were floated to 

develop such a vehicle. One of these was a study conducted by Martin Marietta 

for the Langley Research Center using the basic HL-10 shape. This study, led by 

R.H. Lea, was completed in February 1966 and did not so much concentrate 

on the changes to a lifting body required for orbital flight but, rather, was a 

sales tool for the Titan III family of launch vehicles then under development.87

Nevertheless, the results of this orbital lifting body study looked promising 

enough that NASA Langley issued a contract to Martin in April 1966 to study 

the concept further. Robert L. Lohman was the study manager, and the final 

report was delivered in May 1967. Parametric design and performance data 

was developed for five different lifting body sizes (1, 2, 4, 6, and 8 men) and 

four different launch vehicles (the Titan GLV, a 2- and 5-segment Titan IIIC, 

and a Saturn IB).88

Ultimately, Martin believed that a 25-foot-long, 3-person lifting body 

weighing 12,342 pounds would be the most cost-effective research tool. The 

proposed vehicle included a conventional 2219T-6 aluminum airframe and 

 86.

159
Coming Home

a refurbishable ablative heat shield designed for all-turbulent heating over its 

entire surface. The vehicle could carry a crew of six by removing some of the 

research instrumentation. As part of the study, Martin completed a detailed 

preliminary design of the vehicle and its subsystems, and a minor amount of 

effort was expended on possible modifications to support rendezvous, docking, 

and superorbital (34,000 ft/sec) entry. Although the study was based on the 

HL-10 shape, it was felt that many of the conclusions were equally applicable 

to the M2 and SV-5 vehicles.89

Two nonpiloted flights would be followed by a series of nine piloted flights 

at a cost of $1 billion. The flights would be launched into a 92-by-230-mile 

orbit from Cape Canaveral and recovered at Edwards AFB, with Eglin AFB, 

FL, used as a contingency landing site. Martin believed the first nonpiloted 

suborbital flight could occur 35 months after the authority to proceed. Four 

piloted vehicles would be manufactured, and each could be refurbished and 

reflown every 129 days. It was an aggressive program, and as happened so often 

in high-technology development programs, the engineers were too optimistic 

in their enthusiasm for the system and its capabilities.90

In the end, there would be no serious attempt to launch a lifting body into 

orbit, although researchers at both Langley and the Flight Research Center 

thought such an exercise would be useful. However, neither the Air Force nor 

NASA had sufficient funds to do so given the large expenditures on Apollo 

and later on the Space Shuttle program.

BoMi and Brass Bell

Major General (Dr.) Walter R. Dornberger had been the military director of the 

German Peenemünde test site during World War II, and Dr. Krafft A. Ehricke 

was a close associate. Both men came to the United States in 1947 and worked 

for the Government on various missile-related programs. In 1950, Dornberger 

joined the staff of Bell Aircraft, and Ehricke followed in 1952. While at Bell, 

they advocated the development of a Silverbird-like skip-glide vehicle. Early 

Bell skip-glide designs were simple variations of the Sänger-Bredt Silverbird, 

but the triangular-section straight wings soon began to morph into the delta 

planform more typical of later concepts. During these studies, the company 

devoted a great deal of attention to thermal protection systems and investigated 

both passive and active techniques in considerable detail.

 89.

160
Spaceplane Fantasies

On April 17, 1952, Bell proposed a 1-year study of a piloted BomberMissile (BoMi) consisting of a five-engine rocket booster with a three-engine, 

double-delta skip-glider. The two-person booster was nearly 120 feet long with 

a 60-foot wingspan and was to be manufactured of aluminum alloys with a 

titanium hot-structure wing leading edge. An active cooling system kept the 

airframe within its thermal limits by circulating water between a thin outer 

alloy skin and a heavier inner aluminum skin. A heat exchanger cooled the 

water and jettisoned superheated vapor overboard. The single-seat skip-glider 

was 60 feet long with a 35-foot wingspan and carried a 4,000-pound nuclear 

weapon in a rear-ejecting bomb bay. There appeared to be no military usefulness for an orbital vehicle, so this suborbital design was capable of 2,650 mph 

at 100,000 feet with a 3,300-mile range. The pair took off horizontally, and 

the booster provided power for 2 minutes before the skip-glider separated 

and ignited its engines. The booster flew back to the launch site for a runway 

landing and subsequent reuse. The glider then skipped across the atmosphere 

following the path described by Sänger and Irene Bredt for the Silverbird.91

An initial Air Force review of the suborbital BoMi on April 10, 1953, 

uncovered several deficiencies. It was difficult to envision how the vehicle 

could be adequately cooled, despite a proposed water-wall concept, and there 

was insufficient information concerning stability, control, and aeroelasticity 

at the proposed speeds. The projected range of 3,300 miles was too short for 

intercontinental operations, minimizing its usefulness as either a bomber or a 

reconnaissance vehicle.92

Notwithstanding these misgivings, on April 1, 1954, the Air Force awarded 

Bell a 1-year study contract as project MX-2276. Despite the lack of intercontinental range and the difficult technological questions, the skip-glide concept 

seemed to offer significant strategic potential and, at a minimum, it would provide a useful test vehicle for future hypersonic programs. However, the Air Force 

requirements for MX-2276 were a significant increment beyond what Bell had 

proposed. The vehicle was to be capable of bombardment and reconnaissance 

missions with a maximum velocity of 15,000 mph at 259,000 feet and a range of 

12,000 miles. The MX-2276 contract was not, per se, a hardware-development 

 91.

161
Coming Home

The BoMi (Bomber Missile) was a “boost-glide” concept not unlike the Eugen Sänger’s earlier 

efforts, officially designated the MX-2276. USAF.

162
Spaceplane Fantasies

effort. Instead, Bell was to define requirements for future programs, with a 

focus on such issues as the necessity for a piloted space vehicle, possible mission profiles, structural performance at high temperatures, and the feasibility 

of various guidance systems.93

During early 1955, the Air Force asked the NACA to evaluate the BoMi 

concept. Largely in response to the Air Force request, Allen and Alvin Seiff at 

NACA Ames conducted a classified comparison of the skip-glide and boostglide techniques, building on Allen’s 1954 evaluation. The conclusion was that 

the skip-glider had a slightly longer range for any given initial velocity than an 

equivalent boost-glide concept. However, the researchers again determined that 

the skip-glider would encounter significantly higher heating rates and experience greater magnitude and longer duration g-loads during reentry. This was 

because the transfer of kinetic energy to heat occurred in abrupt pulses during 

the skipping phases of flight—the first skip, of course, was the most severe, 

but all of the early skips had to endure extreme heating. The high heating rate 

necessitated more thermal protection, creating a heavier vehicle, and probably negated the small advantage in predicted range, most likely presenting 

unsolvable technical challenges. The boost-glider, on the other hand, gradually converted kinetic energy to heat over the entire entry trajectory, resulting 

in a relatively low, constant heating rate, but it still created a substantial total 

heat load. The August 1955 report concluded that the boost-glide concept 

“appeared to hold the most promise for an operational military system but 

noted that it would be of questionable usefulness since it could not...be a very 

maneuverable vehicle in the usual sense and thus is, perhaps, limited to use 

for bombing and reconnaissance.” Those, fortunately, were the missions that 

interested the Air Force.94

Based largely on the Ames report, on September 30, 1955, Dr. Ira H. 

Abbott, the NACA Assistant Director for Research, reported to the Air Force 

that the NACA was generally supportive of the boost-glide concept but believed 

that more research was required before a development program could be initiated. Abbott reported that the NACA was less enthralled with the skip-glide 

concept, believing it to be essentially unworkable due to concerns about reentry 

and recovery from space, especially regarding atmospheric heating, structural 

stability at high temperatures, and guidance and control of a vehicle at extreme 

speeds. With encouragement (and, likely, funding) from the Air Force, the 

 93.

163
Coming Home

NACA undertook further investigation into the boost-glide concept, with 

Allen and Neice issuing another report in March 1956.95

The Ames reports marked the end for the dynamic-soaring technique first 

proposed by Sänger, and almost all work on skip-gliding ceased in favor of 

boost-glide concepts. Bell soon replaced the BoMi skip-glider with a generally 

similar boost-glide concept.

It should be noted that although most early lifting reentry vehicles were 

called gliders, this description was only appropriate because they were not 

continually powered. With some exceptions, all of these vehicles carried turbojet engines that would extend the terminal portion (landing) of their flight 

and frequently included rocket engines to act as upper stages to achieve higher 

orbits or to help the vehicle maneuver in space.

Rocket-Bomber

Near the same time, at the request of Trevor Gardner, the Assistant Secretary of 

the Air Force for Research and Development, several boost-glide presentations 

were given to Air Force Headquarters in November 1955. The concept met with 

a warm reception since it offered a chance to leapfrog the technology bottleneck 

that was hampering development of high-speed, long-range bombers.96

In response, on December 19, 1955, the Air Force asked the aerospace 

industry to investigate the concept of a hypersonic bomber. Boeing, Convair, 

Douglas, North American, and Republic ultimately responded, and three companies—Convair, Douglas, and North American—were awarded 1-year study 

contracts. All five companies—later joined by Bell, Lockheed, and Martin—

also used company funds for the study.97 However, midway through the study, 

on June 12, 1956, the Air Force released SR-126 to determine the feasibility 

of an intercontinental bombardment and reconnaissance system that could 

circumnavigate the globe at 100,000 feet altitude, with a maximum altitude 

of 260,000 feet. This piloted hypersonic Rocket Bomber (RoBo) needed to be 

operational by 1965 and superseded the earlier study.98

 95.

164
Spaceplane Fantasies

In late July, 1957, Dornberger presented “An Approach to Manned Orbital 

Flight” to the Air Force Scientific Advisory Board; the information that was 

presented relied heavily on the boost-glide designs Bell had been investigating for the previous 5 years. During the presentation, Dornberger stated that 

engineering solutions have been found for all major problem areas. In most 

cases, these engineering solutions have already been reduced to practical design. 

Frequently, this was backed up by solid evidence.99

For instance, Bell had investigated three solutions to reentry heating. Two 

ideas were quickly discounted: a hot structure similar to the one developed for 

the X-15, and a sweat cooling transpiration system that required too much 

fluid to be practical. The third idea, the same as proposed for BoMi, was an 

active cooling system that consisted of an aluminum inner wall that carried the 

structural load, a layer of Dyna-Flex insulation, and a heat-resistant alloy outer 

skin that could withstand high temperatures but carried no load. Pumps would 

circulate liquid lithium, liquid sodium, or water between the inner and outer 

walls and expend vapor overboard. Surprisingly, Bell estimated that the weight 

of this structure, including the necessary coolant, was comparable to a heatsink structure constructed of aluminum or titanium alloy. The active cooling 

system had been tested extensively in the laboratory, and Bell believed it was 

technically feasible for large-scale applications. During these tests, a liquidsodium cooling loop for the wing leading edge operated for over 600 hours 

(total, not continuously) at 1,200 °F, and silicon carbide leading-edge samples 

had been successfully tested to 4,000 °F. In addition, a 2-foot square sample of 

the water-wall skin structure had been successfully tested for 90 minutes under 

full-intensity heating.100

Researchers at the Arnold Engineering Development Center had evaluated 

the proposed nose shape at Mach 16 and at temperatures as high as 10,000 °F. 

The glider configuration had been tested in a NACA Ames hypersonic wind 

tunnel up to Mach 10, and various pieces had been tested above Mach 12 using 

sounding rockets. Panel flutter had been investigated in a NACA Langley wind 

tunnel at Mach 4 and at temperatures up to 850 °F. Significant test-facility 

time was being expended on the boost-glide concept.

Bell was still not content with the suborbital approach and believed a spaceworthy BoMi could become the first recoverable piloted satellite. The S-BoMi, 

as it was called, was expected to achieve an altitude of 480,000 feet (90 miles) 

at a velocity of 25,640 ft/sec. Bell engineers believed that their knowledge 

 99.

165
Coming Home

was sufficient to allow them to develop the suborbital BoMi within 2 years 

but noted several areas required further study for the orbital S-BoMi. These 

included the continued development of materials, structural insulation, and 

cooling techniques, followed by the development of orbital-navigation equipment. Bell also expected that additional aerodynamic problems would result 

from the interaction of the vehicle skin and the atmosphere in its free molecular 

state—the composition of the atmosphere at the 50-mile altitude proposed for 

BoMi was relatively well understood but at 100 miles was essentially unknown. 

This might have been the most important finding from the various BoMi and 

RoBo studies; nobody truly understood the environment that the vehicles were 

expected to operate in nor the technologies needed to survive in it.101

To correct this, on November 14, 1956, the Air Force issued SR-131 for 

the Hypersonic Weapons and Research and Development Supporting Systems 

(HYWARDS) program, also known as System 610A. The intent of the larger 

program was to provide research data on aerodynamic, structural, human factors, and component problems associated with high-speed atmospheric flight 

and reentry in support of BoMi and RoBo. The HYWARDS vehicle was designated System 455L and was primarily intended to serve as a test vehicle for 

subsystems to be employed in future boost-glide weapons systems.102

In support of the Air Force, researchers at NACA Ames and Langley investigated possible configurations for the HYWARDS vehicle. The Langley effort 

was led by Becker, and the January 17, 1957, report contained the surprising 

recommendation that the design speed of HYWARDS should be increased to 

18,000 ft/sec and at a somewhat lower altitude. Becker pointed out that this 

was where boost-glide vehicles approached their maximum heating environment, mainly because the rapidly increasing altitudes necessary for speeds 

above 18,000 ft/sec caused a reduction in heating rates, and the heating rate 

in space is negligible. An analysis conducted by Langley’s Peter F. Korycinski 

suggested major advantages for a configuration having a flat-bottom surface 

for the delta wing, with the fuselage located in the relative cool of the shielded 

area on the leeside (top) of the wing. This flat-bottom design had the smallest 

critical-heating area for a given wing loading, thus reducing the amount of thermal protection needed. In this respect, the configuration differed considerably 

from the earlier Bell designs that had used a mid-mounted wing. This was one 

101. Ibid.

102. System Requirement 131, Hypersonic Weapons Research and Development Supporting 

Systems, November 14, 1956; Research and Development Project Card, System 455L, 

December 18, 1956. The “L” suffix was generally applied to reconnaissance vehicles, implying 

the Air Force may have expected to build some type of operational follow-on vehicle.

166
Spaceplane Fantasies

of the first clear indications that aerodynamic-design choices could significantly 

alleviate some of the heating and structural concerns associated with entry. 

Interestingly, researchers at Ames reached exactly the opposite conclusions.103

The Ames HYWARDS configuration was devised by a team that included 

Allen and Eggers. The configuration presented in a January 1957 report was a 

flat-top wing body with drooped wingtips that used a combination of surface 

radiation and an internal-cooling system to control aerodynamic heating. Ames 

suggested that the “vehicle appears to pose difficult but solvable structural 

problems.” The vehicle had a maximum speed of approximately Mach 10 but 

a range of only 2,000 miles, compared to 3,300 miles for the Langley design. 

To produce the highest possible L/D, the Ames design made use of the favorable interference lift that occurred when the pressure field of the underslung 

fuselage impinged on a high-mounted wing. Unfortunately, the entire fuselage 

was located in the hottest region of the flow field and the additional weight 

required to keep the airframe cool outweighed the higher L/D. This would be 

addressed further in a September 1957 report.104

In the September report, Ames discussed two configurations, known as A and 

B; each could carry a single pilot, his support equipment, and 1,200 pounds of 

research instrumentation. Both configurations had relatively low wing loadings 

(≈20 pounds per square foot), used a single XLR99 rocket engine, and required 

large external boosters to attain the 18,000-ft/sec design velocity. The airframe 

was similar for both configurations, with a structure that was thermally insulated 

from the outer skin and actively cooled where necessary.105

Research at Ames suggested that the region of highest heating was generally 

on the high-pressure side of the vehicle. In the A configuration, the fuselage 

was removed from the high-pressure region by locating it above the wing, 

which acted as a heat shield, much like with the Langley design. The lower 

surface of the wing was essentially flat in order to facilitate prediction of pressure distributions and heating rates. The shielding effect of the wing kept the 

fuselage relatively cool, greatly simplifying the problem of insulating the pilot 

and instrumentation, and an active cooling system was not required.

Configuration B was essentially the same vehicle proposed in the January 

1957 report. This design combined the flow field of a slender expanding fuselage 

103. Telephone conversations between the author and John V. Becker, various dates in 2001 

and 2002.

104. “Preliminary Investigation of a New Research Airplane for Exploring the Problems of 

Efficient Hypersonic Flight”; “Study of the Feasibility of a Hypersonic Research Airplane,” draft 

NACA Ames Aeronautical Laboratory report (no number) (September 3, 1957).

105. “Study of the Feasibility of a Hypersonic Research Airplane,” pp. 6–11.

167
Coming Home

with the lower surface flow field of the wing to achieve a higher hypersonic 

L/D at a lower angle of attack, which designers expected would minimize the 

effects of heating due to generally slower speeds during reentry. Locating the 

fuselage below the wing also cambered the configuration, making it essentially 

self-trimming at the maximum L/D, thereby reducing drag due to trimming 

effects. The drooped, toe-in wingtips were used to achieve static stability and 

control at hypersonic speeds, supplemented by a retractable ventral stabilizer 

to improve stability below Mach 6.

The Ames study was notable in that it delved deeply into various thermal protection schemes. Three approaches were investigated in detail: (1) an 

unprotected hot structure, (2) an internally cooled structure, and (3) an airframe protected from the heat by external insulation. Although the last two 

approaches were expected to be more complex, the potential weight savings 

intrigued researchers.

Since Inconel X was being used in the X-15 hot structure, Ames concentrated on the more-advanced alloys that had recently become available. A 

cobalt-based L-605 alloy demonstrated adequate mechanical properties (it did 

not become brittle) and oxidation resistance (it did not ablate) up to about 

1,800 °F. Some molybdenum alloys retained good strength characteristics even 

above this temperature but oxidized rapidly without protective coatings, which 

did not exist at the time. Ceramic materials were not seriously considered for 

structural elements since they tended to be brittle and had minimal loadcarrying abilities, but they were considered essential for small, hot areas such 

as the wing leading edge.106

An investigation of active cooling systems demonstrated the ability to keep 

temperatures within existing hot-structure technology but left little hope that 

a workable vehicle could be designed around them. The problem was that 

coolant is heavy, and even for the small research aircraft being contemplated, 

over 2,000 pounds of coolant—water or liquid helium—would be needed. The 

study eventually concludes that “unless high heat-capacity coolants can also 

be used efficiently for propulsion, the weight of a design in which surfaces are 

directly cooled appears prohibitive.”107 Since the boost-glide designs were, as 

the name implies, gliders for most of their flights, even this approach did not 

hold much promise.

Preliminary investigations into insulating the airframe from heat effects 

looked promising, but researchers realized that any protective layer must also 

be capable of withstanding high-speed aerodynamic effects and anticipated 

106. “Study of the Feasibility of a Hypersonic Research Airplane,” pp. 11–17.

107. Ibid., p. 18.

168
Spaceplane Fantasies

acoustic levels of over 170 decibels. The most promising design for temperatures in excess of 1,800 °F was a metallic superalloy outer skin with closely 

spaced stand-offs that allowed a thick layer of Thermoflex insulation between 

the outer skin and the titanium structure underneath. Researchers estimated 

that this would add about 2 pounds per square foot to the weight of the 

vehicle. Because of differences in thermal expansion between the insulating and 

primary structures, expansion joints would be required at appropriate intervals. These, in turn, could lead to surface irregularities that could significantly 

increase local-heating rates at high speed (due to boundary layer interaction), 

and they would have to be designed and maintained carefully. The study notes 

that the “successful development of a smooth and lightweight insulating structure with adequate life expectancy is believed to represent a major development 

effort.” Truer words were never written, but the general approach evolved into 

the thermal protection system used on the Space Shuttle orbiters.108

The September Ames study had added the low-wing A configuration 

largely in response to criticisms from Langley regarding the initial highwing configuration proposed in January. Nevertheless, researchers at Ames 

continued to champion the high-wing design, arguing that, “it is apparent 

that the flat-top arrangement is the more efficient, claiming tests showed the 

design offered a 35 percent increase in performance.”109 This design brought 

to a head the debate within the NACA on the relative merits of flat-top 

versus flat-bottom designs, a battle that the flat-bottom advocates eventually 

won. However, while the Air Force and the NACA tried to sort out exactly 

what HYWARDS should look like, the entire effort was soon overcome by 

other events.

The Stillborn Round Three X-Plane

In mid-1955, the Soviet Union and the United States had separately announced 

intentions to launch Earth-orbital satellites as part of the 1957 International 

Geophysical Year. Nevertheless, when the Soviet Union launched Sputnik on 

October 4, 1957, the event created a stir among the popular press. The perceived lack of response by President Dwight D. Eisenhower further antagonized the media and, shortly after, the American people, as well. However, it 

108. “Study of the Feasibility of a Hypersonic Research Airplane,” pp. 18–19.

109. Ames Aeronautical Laboratory, “Preliminary Investigation of a New Research Airplane 

for Exploring the Problems of Efficient Hypersonic Flight,” NACA, Washington, DC, January 18, 

1957, p. 32.

169
Coming Home

was the 1,100-pound Sputnik II that ultimately caused the administration 

to take action since it graphically portrayed the heavy-lift capability of Soviet 

launch vehicles, and indirectly, of their ICBM program.

By the end of August 1957, the Air Force had already consolidated the 

BoMi, RoBo, and HYWARDS efforts into a single three-step development 

program designated System 464L and named Dyna-Soar (for dynamic soaring, as Sänger had referred to skipping reentry). The name was ironic since the 

skip-gliding concept had been abandoned as unworkable 2 years earlier. The 

classified title of the new program was “Hypersonic Glide Rocket Weapons 

System,” and by the end of the year, government and industry had spent a little 

over $8 million ($93 million in 2010 inflation-adjusted dollars) studying the 

boost-glide concept.110

Independently, the NACA Hypersonic Research Steering Committee met 

on October 15, 1957, to determine the direction of the Round Three research 

airplane. All agreed that whatever came after X-15 should have a large increment in performance, but no specific requirements were discussed. Perhaps 

understandably given the launch of Sputnik I, the meeting concentrated more 

on the future of human space flight, with three different approaches being 

proposed. A minority led by Faget from NACA Langley argued for a ballistic 

blunt-body shape—essentially what became the Mercury capsule. Another 

minority favored the lifting body approach of tailoring a blunt reentry shape to 

provide a modest L/D that permitted limited maneuvering.111 The remainder of 

the committee preferred the flat-bottom hypersonic glider described by Becker. 

The proposal from Becker appeared to mesh nicely with the new Dyna-Soar 

effort, so the committee recommended the Round Three requirements be 

included in the Air Force program.112

Step I of the three-step Dyna-Soar program would include a small single-seat hypersonic boost-glide conceptual test vehicle that would provide 

110. R&D Project Card, System 464L, August 23, 1957, p. 1. Dyna-Soar also stood, apparently, 

for Dynamic Soarer, as the Air Force called the vehicle itself.

111. Interestingly, later NASA capsule designs had substantial increases in their L/D. In fact, 

Apollo, with a hypersonic L/D between 0.6 and 0.8, rivaled some of the lifting bodies in crossrange capability, although it was never used operationally.

112. John W. Crowley to Milton Ames, “Meeting of Round III Steering Committee to be held 

at NACA Headquarters, July 2, 1957” (June 18, 1957); memorandum to the Director from 

Clotaire Wood, “Presentation to Air Force Headquarters on Round III” (July 11, 1957); and 

John W. Crowley, NACA Associate Director for Research, to NACA field centers, “Round Three 

Conclusions” (November 15, 1957), all in possession of authors; “This is Dyna-Soar,” Boeing 

News Release S-6826, 1963.

170
Spaceplane Fantasies

information on aerodynamic, structural, human factors, and component 

problems and serve as a test bed for subsystems development. After preliminary air-drop flights, a single-stage launch vehicle would boost the glider to 

12,000 ft/sec and 360,000 feet in altitude sometime in 1966. This phase of 

Dyna-Soar largely satisfied the anticipated Round Three research airplane 

goals.113 Step II included a vehicle that could provide high-quality photographic and radar intelligence as well as perform limited bombardment missions. A two-stage booster would propel the vehicle to 22,000 ft/sec at an 

altitude of 170,000 feet, enabling it to glide 5,750 miles. It was expected 

that the engines and guidance systems being developed for the ballistic missile programs would be used beginning in 1969.114 Step III, expected to be 

operational in 1974, incorporated most of the strategic reconnaissance and 

bombardment capabilities previously envisioned for RoBo.115

By January 25, 1958, the Air Force had screened a list of 111 potential bidders 

for the conceptual test vehicle, and 10 companies were selected to receive requests 

for proposals, including Bell, Boeing, Chance Vought, Convair, Douglas, General 

Electric, Lockheed, Martin, North American, and Western Electric. Three additional contractors—McDonnell, Northrop, and Republic—were subsequently 

added to the list, although several teaming agreements resulted in only nine actual 

contenders for the contract.116 After reviewing the proposals, on June 16, 1958, 

the Air Force awarded two-phase predevelopment contracts to Boeing-Vought 

and Martin-Bell to refine their concepts.117

X-20 Dyna-Soar

Perhaps the most interesting of all discussions about reentry and recovery from 

space concerning a winged, reusable vehicle in the 1960s involves the storied 

career of the X-20 Dyna-Soar. It is one of those flying vehicles about which 

113. R&D Project Card, System 464L, August 23, 1957, pp. 2–3.

114. Intercontinental missile programs does not necessarily mean ICBMs since Navaho was 

also considered a prime source.

115. R&D Project Card, System 464L, August 23, 1957, pp. 6–7.

116. Colonel H.M. Harlow to Source Selection Board, “Report of Working Group for System 

464L” (January 25, 1958).

117. Major General Ralph P. Swofford, Jr., Acting DCS/Development, Headquarters USAF to 

Commander ARDC, “Selection of Contractor for WS-464L (Dyna-Soar) Development” (June 25, 1958).

171
Coming Home

The U.S. Air Force transformed the 1930s spaceplane concepts of Eugen Sänger and others into 

an experimental vehicle: the Dyna-Soar (Dynamic Soaring). Had it flown, it might have been useful for reconnaissance and long-range bombing from the edge of space. Research on the piloted 

spaceplane lasted from 1957 to 1963, when the program was canceled before the first vehicle 

ever flew. The X-20 program was the first serious effort to build and test a winged vehicle that 

could fly to and from space. USAF.

172
Spaceplane Fantasies

legend and myth abound.118 Dyna-Soar was officially designated System 620A 

on November 9, 1959, and the Boeing-Vought team was declared the winner 

on December 11, 1959.119 At the same time, the Glenn L. Martin Company 

was selected to develop a human-rated version of the Titan launch vehicle.120

After several reviews and much political infighting, on April 27, 1960, the Air 

Force awarded a contract to Boeing as the overall integration contractor and 

ordered 10 Dyna-Soar gliders.121 The procurement schedule called for two 

vehicles to be delivered during 1965, four in 1966, and two during 1967. 

The other two airframes were to be used for static tests beginning in 1965.122

The development of Dyna-Soar ranks as perhaps the most convoluted 

process in modern defense history. Constant fighting between the Air Force 

and the Department of Defense, particularly Secretary Robert S. McNamara, 

caused the program to change direction countless times and hindered any real 

sense of progress. On the other hand, it is a tribute to the technical sides of the 

Air Force and Boeing that the development effort made real progress. DynaSoar was the first serious attempt to build a piloted, reusable, lifting reentry 

spacecraft, and it began before Project Mercury had taken flight. Most in the 

industry considered the capsule concept crude, while Dyna-Soar was perceived 

as elegant since it was maneuverable, reusable, and landed on a runway like an 

airplane. It was also vastly more complicated and required substantially more 

research, particularly in the area of thermal protection systems. An entire book 

would be needed to tell the story, so we will skip most of it since it is beyond 

118. The most thorough of all studies on this subject is Roy A. Houchin, U.S. Hypersonic 

Research and Development: The Rise and Fall of Dyna-Soar, 1944–1963 (New York: Routledge, 

2006).

119. This was contract AF33(600)-39831. Vought’s involvement with the Dyna-Soar program 

eventually dwindled to the design and fabrication of the high-temperature nose-cap. This work 

would later serve Vought (later LTV) well when asked to develop the carbon-carbon nose cap and 

wing leading edges for the Space Shuttle.

120. This was well before the Titan II was selected for the NASA Gemini program. The letter 

contract for the Titan II GLV was issued by the Air Force Space Systems Division on January 19, 

1962. The Glenn L. Martin Company merged with the American-Marietta Corporation in 1961 to 

form the Martin Marietta Corporation.

121. The Boeing contract was AF33(657)-7132, and the gliders were assigned serial numbers 

61-2374 through 2383.

122. Geiger, “Strangled Infant,” pp. 227–226; TWX, RDZSXB-31253-E, Headquarters ARDC to 

Headquarters Ballistic Missile Division, November 10, 1959; TWX, AFDAT-90938, Headquarters 

USAF to Commander, ARDC, November 17, 1959. 

173
Coming Home

The launch vehicle for the X-20 evolved considerably over the course of the program, beginning 

with a Titan II (shown) and ending with a human-rated version of the Titan III. USAF.

174
Spaceplane Fantasies

the scope of this monograph and concentrate on the aspects of Dyna-Soar 

related to reentry and recovery from space.

Despite the funding issues, political infighting, and confusion over roles, 

by the summer of 1961, Boeing had made significant progress on the design 

of the Model 844-2050 Dyna-Soar I glider. The development effort was 

directed by Boeing Vice President George M. Snyder in Seattle, WA. The final 

Dyna-Soar glider was the result of over 13,000 hours of wind tunnel tests that 

included approximately 1,800 hours of subsonic, 2,700 hours of supersonic, 

and 8,500 hours of hypersonic time. Several innovative test techniques were 

used, including a temperature-sensitive paint that was applied to wind tunnel 

models to determine the heating levels experienced during the test. DynaSoar was among the first users of this new paint, which allowed researchers 

to use relatively inexpensive wood models, instead of metal models, equipped 

with a myriad of temperature transducers. In addition, a great deal was 

learned about material science, including new techniques for the manufacture, welding, and extruding of various high-strength superalloys.123

The Air Force and NASA reviewed a full-scale mockup on September 

11, 1961. The Model 844-2050 was 35.3 feet long, spanned 20.4 feet, and 

had approximately 345 square feet of wing area. The glider weighed 10,830 

pounds empty, and it could carry a maximum payload of 1,000 pounds in 

a 75-cubic-foot compartment that was pressurized at 0.7 atmosphere with 

100-percent nitrogen.124

The wing on early Dyna-Soar configurations used a double-wedge upper 

surface with a flat bottom that provided good hypersonic stability and was easy 

to manufacture. Unfortunately, this design required the addition of retractable stabilizers for low-speed flight, so the upper wing surface was modified 

to improve low-speed handling without compromising hypersonic stability. 

This modification, however, resulted in some transonic instabilities that were 

eliminated by adding an aft fuselage ramp on the Model 844-2050E, giving 

the final Dyna-Soar its distinctive appearance.125

123. “X-20 Dyna-Soar Information Fact Sheet,” January 1963, p. 5; Terry Smith, “Dyna-Soar 

X-20: A Look at the Hardware and Technology,” Quest: The History of Spaceflight Magazine, 

Winter 1994, pp. 23–28; X-20 Program Report, Boeing report D2-80852 (March 1963).

124. Dyna-Soar I Structure Description, Boeing report D2-6909-2 (November 1961); “X-20 

Dyna-Soar Information Fact Sheet,” p. 5; Smith, “Dyna-Soar X-24,” pp. 23–28; X-20 Program 

Report, Boeing report D2-80852 (March 1963).

125. Memorandum, to the Chief of the Research Division, ARDC, “Development of the DynaSoar Configuration” (August 23, 1962).

175
Coming Home

To handle the thermal stresses during entry, the internal structure of the 

fuselage differed significantly from conventional aircraft of the period. It consisted of a René 41 truss-space frame with fixed and pinned joints in square 

and triangular elements that looked similar to bridge construction. There were 

forward and aft beams on each side of the fuselage, with six cross frames connecting each side. In the aft bay, a superstructure of truss members spanned 

between the main-beam upper chords to react with the skin-support beam 

loads and provide a load path for torsional flight loads. Wherever possible, the 

axially loaded truss members were round, swaged-end tubes with sheet-metal 

tabs or fusion-welded machined fittings. Swaging the tube ends allowed for 

compact joints at which up to 11 members intersected at a common point. 

Members subjected to other than pure axial loads (bending or combined bending and axial) were built-up rectangular tube sections fusion-welded together. 

The space frame was capable of deforming in response to thermal gradients 

without introducing major stresses or losing structural rigidity. The radiationcooled structure was designed to survive four reentries.126

The thermal protection system for Dyna-Soar was a major challenge, presaging the problems on the Space Shuttle. The structure was designed to achieve 

an equilibrium-heating condition during entry by radiating heat back to space. 

Engineers were concerned about the emissivity of the external skins; they were 

hoping to achieve 0.80, but even a small reduction had a major impact on the 

expected internal and external temperatures. Unlike ablative or heat-sink concepts, this design placed no time constraints on entry but could only withstand 

somewhat milder heat pulses. The wide range of L/Ds (and therefore angle of 

attack) planned for the glider meant that the thermal protection system for 

each location had to be designed for the worst-case heating environments. The 

nose cap and wing leading edges received the highest heating at low angle of 

attack (high L/D), while the bottom of the wing and control surfaces received 

their highest heating at high angle of attack (low L/D).127

Where temperatures were expected to be less than 1,000 °F, the internal 

truss structure was covered by skin using a series of René 41 panels that were 

corrugated to add stiffness and allow expansion during entry heating. On the 

lower surface of the wing, a series of D-36 columbium panels were attached to 

the underlying René 41 panels using standoff clips with a layer of Dyna-Flex or 

126. “X-20 Dyna-Soar Information Fact Sheet,” p. 5; Smith, “Dyna-Soar X-20,” pp. 23–28; X-20 

Program Report, Boeing report D2-80852 (March 1963).

 127. Testing Lifting Bodies at Edwards. Note that hypersonic gliders utilized the backside of the 

L/D curve rather than the frontside, where a lower angle of attack would normally yield a lower L/D.

176
Spaceplane Fantasies

Micro-QuartzTM insulation sandwiched between them.128 Each external surface 

panel overlapped the adjoining panel, allowing adjacent panels to expand and 

move relative to each other without buckling. Special attention was given to 

heat leakage at the expansion joints, access panels, landing gear doors, and 

hinge area at the elevon-control surfaces. The wing leading edge, expected to 

encounter temperatures up to 3,000 °F, was constructed from segmented TZM, 

an alloy of titanium, molybdenum, and zirconium.129

Oxidation of the refractory metals used in the heat shield could lead to 

structural failure, so a silicide coating was used on the D-36 columbium and 

TZM panels. A final layer of synar-silicon carbide applied over the silicide coating gave Dyna-Soar its distinctive black color. Although this solved the oxidation issue, these coatings would have had to be refurbished after each flight.130

There were two different nose-cap designs, with the early Boeing concept 

becoming a contingency design after Vought proved to have a better idea. 

Vought developed a siliconized graphite structure overlaid with zirconia tiles 

restrained by zirconia pins. In case of cracks in the structure, the tiles and 

pins were held in place by platinum-rhodium wire. The nose cap was flexibly 

mounted to a support ring, which in turn was supported at three locations by 

a linkage arrangement attached to the glider truss structure. The contingency 

Boeing design used a single-piece zirconia structure reinforced with platinumrhodium wire. During the molding process, shaped tiles were cast in the outside 

surface to allow thermal expansion and to prevent possible cracks from spreading. Each design was capable of withstanding 4,300 °F and was later tested 

during the ASSET series.131

The cockpit windows were the largest designed for a piloted spacecraft at 

the time, and the forward three windows were covered by a D-36 columbium 

heat shield during the 2,000 °F entry. The cover was jettisoned as the glider 

slowed to supersonic speeds to provide better visibility during landing. The 

single window on each side remained uncovered since they were not subjected 

to high heating loads. NASA test pilot Neil Armstrong used a Douglas F5D 

128. Dyna-Flex is also known as Cerachrome, while Micro-Quartz goes by the name Q-Fiber 

Felt. Both are fibrous batt insulating materials that continue to be used in a variety of hightemperature applications.

129. System Development Plan, Dyna-Soar (Step I) Program, 620A, April 1, 1960, pp. III-8 to 

III-12, III-49 to III-52, III-65 to III-72, III-195 to III-207; System Package Program, System 620A, 

Dyna-Soar, April 26, 1961, pp. 84–92, 132–139, 182–187.

130. System Package Program, System 620A, Dyna-Soar, April 26, 1961, pp. 84–92, 

132–139, 182–187.

131. Testing Lifting Bodies at Edwards.

177
Coming Home

Dyna-Soar was to employ a refractory metal heat shield that engineers expected to use multiple 

times with only limited refurbishment. This differed significantly from the ablative heat shields 

used by the capsules and the ceramic heat shield used by the later Space Shuttle. USAF.

Skylancer at Edwards AFB to demonstrate a pilot could land with side vision 

only in the event that the heat shield did not jettison as planned.132

Unlike the capsules, Dyna-Soar was designed to land on a concrete runway 

8,000 feet long and 150 feet wide using a three-point landing-skid arrangement 

based partially on X-15 experience. The Goodyear main landing skids used René 

41 bristles wound around a series of longitudinal rods, and the Bendix nose skid 

(called a dish or a plate) was a single-piece René 41 forging. As it turned out, 

the landing gear was one of the larger unknowns as the program progressed.133

Dyna-Soar had a cross range of 2,000 miles, and the glider was designed to 

fly at any hypersonic L/D between 0.6 and 1.8, corresponding to angles of attack 

between 55 and 18 degrees.By combining different bank angles with different 

L/Ds, the vehicle could fly a wide variety of entry trajectories. Selecting a low 

L/D with zero bank angle would result in a short, straight-ahead trajectory, while 

selecting a high L/D with a large bank angle would produce a long, turning entry 

132. Six pilots had already been selected to fly Dyna-Soar—Captain Albert H. Crews, Major 

Henry C. Gordon, Captain William J. Pete Knight, Major Russell L. Rogers, Major James Wood 

(chief pilot), and NASA test pilot Milton O. Thompson. NASA research pilots Neil A. Armstrong 

and William H. Dana also supported development of the Dyna-Soar.

 133. Testing Lifting Bodies at Edwards; X-20 Program Report, Boeing report D2-80852 (March 1963).

178
Spaceplane Fantasies

and high cross range. The footprint for Dyna-Soar at the entry interface (400,000 

feet) was approximately 3,000 miles wide and 8,000 miles long, with the size 

of the footprint getting smaller as the vehicle decelerated. The glider could land 

anywhere within the footprint. The subsonic L/D was 4.25, and the glider used 

approach and landing techniques similar to the X-15.134

Unfortunately, the end of Dyna-Soar was as convoluted as the beginning. 

By the summer of 1963, those on the program knew the end was near, but the 

machinations are well outside the scope of this monograph. When it became 

obvious that McNamara wanted to cancel Dyna-Soar in favor of a new program suggested by Dr. Harold Brown, Director of Defense for Research and 

Engineering, NASA made it clear that the Agency did not agree. Dr. Raymond 

L. Bisplinghoff, The NASA’s Associate Administrator for Advanced Research 

and Technology, pointed out that advanced flight-system studies had repeatedly 

shown the importance of developing a maneuverable lifting reentry vehicle with a 

high-temperature, radiation-cooled metal structure. Test facilities were unable to 

adequately simulate this environment, and, consequently, a research vehicle such 

as Dyna-Soar was necessary. Bisplinghoff reminded the Department of Defense 

that NASA had always supported the Dyna-Soar program and would probably 

need to initiate a substitute research program if Dyna-Soar were cancelled, effectively negating any cost savings to the U.S. Treasury.135

On December 4, 1963, Dr. Alexander H. Flax, the Assistant Secretary 

of the Air Force for Research and Development, wrote to McNamara firmly 

supporting the continued development of Dyna-Soar and rebutting Brown’s 

argument against the program. Since $410 million had already been expended 

on Dyna-Soar, Flax questioned the proposal to cancel Dyna-Soar and initiate a 

new program with similar objectives.136 The Secretary of the Air Force, Eugene 

M. Zuckert, and Under Secretary of the Air Force, Dr. Brockway McMillan, 

agreed with Flax.137 Zuckert further stated that he did not wish to see the Air 

 134.

179
Coming Home

Each of the Dyna-Soar glider’s external surface panels overlapped the adjoining panel, allowing 

adjacent panels to expand and move relative to each other without buckling. USAF.

Force abandon Dyna-Soar and start a new program that had been projected 

based on undoubtedly optimistic costs and schedules.138

It was for naught. On December 10, 1963, McNamara cancelled the DynaSoar program. In its place, the Air Force would begin developing the Manned 

Orbiting Laboratory (MOL) serviced by Gemini-B capsules.139 The Secretary 

of Defense explained that Dyna-Soar was not designed to perform space138. ASAF/R&D to the Secretary of the Air Force, “Manned Military Space Program,” December 

4, 1963; Secretary of the Air Force to the Secretary of Defense, “Manned Military Space 

Program,” December 4, 1963.

139. The MOL program itself would also be canceled before its first flight. The launch complex 

(SLC-6) constructed at Vandenberg AFB to support the MOL program was later modified to serve 

as the west coast Space Shuttle launch site. The STS-33/51-L Challenger accident caused the 

Air Force to mothball and later cancel plans to use the launch complex. During 1990, it was 

proposed to rebuild SLC-6 to support Titan IV, a variation of the booster originally intended for 

Dyna-Soar and MOL. This was also cancelled prior to completion. Over $7 billion was spent on 

the facility to support various projects, yet the only launches to take place from it were several 

small Lockheed Launch Vehicle boosters. During 2003–2005, the complex was modified by 

Boeing to support the Delta IV EELV program. The first successful launch of a large booster was 

a Delta IV Medium on June 27, 2006.

180
Spaceplane Fantasies

logistics operations, place substantial payloads into space, or fulfill extended 

orbital missions. Somehow, McNamara completely ignored what the program 

had intended to accomplish and criticized it for not being things it had never 

aspired to. This ended the first serious attempt to build a reusable piloted lifting reentry vehicle.140

On December 18, Zuckert approved the continuation of 10 Dyna-Soar 

activities, including the fabrication of the thermal protection system, the construction of the full pressure suit, the development of the nose cap, and the 

flight-testing of the coated molybdenum panels and nose caps on ASSET.141

At the time of cancellation, nearly all of the engineering drawings for the X-20 

glider had been released and all of the basic material had been delivered. The 

construction of the crew compartment was essentially complete and the wing 

spars, vertical stabilizer spars, and fuselage primary structure were in the final 

assembly jig. Also at that time, Boeing had 6,475 people involved in the X-20 

program, while Minneapolis-Honeywell had 630 and RCA had 565.142

Despite this apparent progress, there were several known technical problems looming, although none were considered major hurdles at the time. For 

instance, Boeing was developing production methods for forming, fastening, 

drilling, and otherwise working with coated columbium, coated molybdenum, 

and René 41, and this learning process was causing some delays in the assembly of the glider. Tests had shown that the landing-gear brushes and nose dish 

were not suited for landings on concrete runways due to excessive wear and 

140. TWX, AFCVC-1918/63, Headquarters USAF to All Commands, December 10, 1963; News 

Briefing, Secretary of Defense, “Cancellation of the X-20 Program,” December 10, 1963.

141. TWX, ASZR-10-12-1011, Headquarters Aeronautical Systems Division (ASD) to 

Headquarters Space Systems Division, December 10, 1963; TWX, ASZRK-12-10-249, 

Headquarters ASD to Air Force Plant Representative (AFPR), Boeing, December 10, 1963; 

TWX, SAF to Headquarters AFSC, December 11, 1963; Secretary of the Air Force to the Air 

Force Chief of Staff, “Dyna-Soar Termination,” December 12, 1963; X-20 Phase-out Plan, X-20 

SPO, December 13, 1963, pp. IV-I to IV-II; TWX, AFRDD-79094, CS/AF to Headquarters AFSC, 

December 18, 1963.

142. P.J. DiSalvo, Dep. Dir/Procurement, X-20 SPO to AFPR, Boeing Co., “AF33(657)-

7132, The Boeing Company Percentage of Completion and SPO Recommendations for 

Final Settlement,” March 12, 1964; DiSalvo to AFPR, Minneapolis-Honeywell Regulator Co., 

“AF33(657)-7133, Minneapolis-Honeywell Regulator Company Percentage of Completion 

and SPO Recommendation for Final Settlement,” February 27, 1964; DiSalvo to AFPR, RCA, 

“Contract AF33(657)-7134, Radio Corporation of America, Percentage of Completion and SPO 

Recommendation for Final Settlement,” March 30, 1964; X-20 Detailed Termination Plan, X-20 

SPO, January 23, 1964, p. III-2.

181
Coming Home

inadequate tracking stability. Alternate concepts were under development, but 

in the meantime, the glider would land on the dry lakebeds at Edwards, just 

like the X-15.143

Several relatively minor structural problems existed at the interface between 

the glider and the booster, but there was a much more serious problem related 

to the aerodynamics of the mated vehicle. The winged surfaces of the glider, 

mounted far forward on the launch vehicle, created instabilities during the early 

boost phase.144 The addition of fins at the bottom of the Titan III booster (as 

proposed on an earlier Titan II version) regained the static stability, but wind 

tunnel tests uncovered some unexpected aerodynamic interference between the 

forward wing and the aft fins, largely negating their effectiveness. The added 

weight of the fins was also undesirable. Increasing the control authority of the 

rocket nozzles allowed the instability to be properly controlled without the fins, 

but this introduced excessive structural loads in the booster and at the glider/

booster attachment point while transiting wind shears early in the ascent. At 

the program’s termination, a solution was still being sought.145

When Dyna-Soar was canceled, the first X-20 air launch from the B-52C 

was about 18 months out, and the first flight on a Titan IIIC was 30 months 

in the future.146 Approximately $410 million ($3.7 billion in 2009 inflationadjusted dollars) had already been spent, and another $373 million was needed 

prior to the first orbital flight. Even if Dyna-Soar had never flown an operational mission, it would have provided valuable information on entry flight 

control and heating, something that was seriously lacking during the development of the Space Shuttle 10 years later.

The 1970 Hypersonic Vehicles Report: 

Setting the Stage for the Future

Test models of many of the concepts were flown in wind tunnels, and a few in 

free flights. For instance, in 1969 North American built a 450-pound model 

of an elongated, conical piloted reentry vehicle that had a set of rotors stowed 

in the upper fuselage. As the vehicle slowed to subsonic speeds, the rotors 

deployed and the blades began to unfold. As speed reduced further, the rotors 

143. Testing Lifting Bodies at Edwards.

144. The lifting body concepts eliminated, or at least minimized, this destabilizing effect since 

the overall vehicle width would have been about the same as the diameter of the booster.

145. Testing Lifting Bodies at Edwards.

146. Ibid.

182
Spaceplane Fantasies

opened fully and the vehicle autorotated until it reached the landing area. The 

tests demonstrated that the concept worked, and the weight penalty was not 

as extreme as first expected. However, design studies indicated that available 

technology would limit the application of the technique to vehicles weighing 

less than 50,000 pounds. This was unacceptable for projected programs such 

as the Space Shuttle.147

147. E.P. Smith, “Space Shuttle in Perspective: History in the Making,” AIAA paper 75-336 

presented at the AIAA 11th Annual Meeting and Technical Display, Washington, DC, February 

24–26, 1975, p. 6.

183
When the concept of a space shuttle emerged in the latter 1960s, NASA envisioned it as a 

two-stage, fully reusable vehicle with both parts landing on a runway. This wind tunnel model 

represented the “DC-3” concept developed by Max Faget at the Manned Spacecraft Center 

(renamed the Johnson Space Center in 1973). NASA.

184
CHAPTER 5

Although we tend to think of the Space Shuttle as more modern than the capsules that preceded it, such was not truly the case. Dyna-Soar, Mercury, and 

Apollo were all first-generation spacecraft, and their development took place 

mostly concurrently, with little chance for one program to benefit greatly from 

the others. Gemini came later and benefited from Mercury to become the first 

second-generation spacecraft (although one could argue that the Apollo Block 

II Command Module also qualifies). Since the development of Dyna-Soar was 

cancelled before its first flight, the Space Shuttle represents the first-generation 

lifting reentry spaceplane. When the initial studies that led to the Space Shuttle 

began, the United States had completed fewer than two-dozen piloted space 

flights using the three capsule types; indeed, when Columbia made her maiden 

voyage, the United States had only made 30 orbital flights. It was an extremely 

small base of experience upon which to begin such an ambitious program. 

The audacity of the space community in successfully bringing this revolution 

in space flight to fruition should not be minimized in assessing the history of 

the Space Shuttle.

The engineers who created the Space Shuttle did so with grand expectations. 

However, the exercise must be looked at in the context of its time. The United 

States was riding high on the successes of the early human space flight programs, 

particularly the Apollo Moon landing program, and visions of space stations and 

the space shuttles that serviced them were firmly implanted in the minds of the 

engineers and the public. Arthur C. Clarke and Stanley Kubrick furthered these 

ideas to the music of “The Blue Danube” in the movie 2001: A Space Odyssey.

Since the beginning, most visions of human space flight revolved around 

the concept of a spaceplane, a logical extension of the airplane that had become 

commonplace during the 20th century. It came as a disappointment to many 

engineers and researchers when the first humans ventured into space aboard 

capsules. Test pilots were known to remark that astronauts were “Spam in a 

can”—many engineers were not much kinder in their critiques of the capsules 

used by Mercury, Gemini, and Apollo. However, the simple fact was that there 

was no other way to get humans into space at the time. Redstone and Atlas 

simply were not powerful enough to launch a spaceplane, and the political 

realities dictated that the United States could not wait.

185
Coming Home

Titan probably was powerful enough to launch a small winged vehicle, 

perhaps like the one proposed by John Becker at the NACA Conference on 

High-Speed Aerodynamics in March 1958. However, by the time Titan flew, 

President John F. Kennedy had committed to a Moon landing before the end of 

the decade, so the momentum for capsules continued. Nevertheless, visions of 

spaceplanes were abundant, and the Air Force committed to building the first 

reusable, lifting reentry spacecraft when it began the development of DynaSoar. The X-20 effort took little from the NASA capsules since their development was concurrent, and the entire concept was much more advanced. Instead 

of throwing away the vehicle after every flight, as with the capsules, Dyna-Soar 

was designed to fly four times with only minor refurbishment of its superalloy metallic heat shield. It would be maneuverable during entry and able to 

land on a conventional runway. It was also incredibly controversial within the 

Department of Defense and ultimately was cancelled by Secretary of Defense 

Robert S. McNamara before the first vehicle was completed. Regardless of 

whether there was a valid military mission for the spacecraft, its development 

and flight-test would have provided a much-needed body of knowledge when 

it came time to build the next winged space vehicle—the Space Shuttle.

NASA and the Air Force began studying vehicles that resembled space shuttles 

in the late 1950s and continued throughout the 1960s. The development of the 

actual vehicle called the Space Shuttle began when the NASA Manned Spacecraft 

Center (MSC) and the Marshall Space Flight Center (MSFC) released a joint 

Request for Proposals (RFP) for an 8-month study of an Integral Launch and 

Reentry Vehicle (ILRV). Five companies—the Convair Division of General 

Dynamics, Lockheed, McDonnell Douglas, Martin Marietta, and North 

American Rockwell—expressed an interest in participating in the study, which 

was Phase A of the Space Shuttle design and development cycle.1 In NASA’s 

four-phase process, Phase A was labeled advanced studies, Phase B was project 

definition, Phase C was vehicle design, and Phase D was production and operations. The 21st century Constellation program is following a similar process.

Throughout the Phase A studies, there was considerable technical controversy 

within NASA over the size and configuration of the space shuttle. These studies—supported by more than 200 man-years of engineering effort and backed 

up by countless hours of wind tunnel testing, material evaluation, and structural design—resulted in four basic vehicle types. These included straight-wing, 

delta-wing, and stowed-wing vehicles as well as lifting bodies. Combined with 

independent Air Force ILRV studies, Phase A confirmed that cross range was the 

major controversy when trying to consolidate Air Force requirements with the 

 1. Request for Proposals (RFP), MSC-BG721-28-9-96C and MSFC-1-7-21-00020, October 30, 1968.

186
Spaceplane Reality

needs of NASA. The major issue was the Air Force wanted the space shuttle to 

be able to land at its launch site after only a single polar orbit. During this time, 

Earth had rotated approximately 1,265 miles, meaning the returning spacecraft 

had to be able to fly at least that distance during entry. On the other hand, 

NASA simply wanted an opportunity to land back at the launch site from lowinclination orbits once every 24 hours, requiring a relatively modest 265 miles 

in cross-range capability (although various abort scenarios raised this to almost 

500 miles in most studies). Still later, abort considerations would raise the NASA 

requirement to over 1,000 miles, almost coinciding with the Air Force’s requirement. In addition, national-security considerations also drove the decision on 

the size of the payload bay. The reconnaissance program had a particular satellite 

being planned for deployment from the Shuttle, and this drove the length of 

the payload bay (the diameter was driven by the size of the expected modules 

for a space station).

During Phase A, engineers found the lifting body concept to be the most 

poorly suited to space shuttle applications, primarily because the shape did not 

lend itself to efficient packaging and installation of a large payload bay, propellant 

tanks, and major subsystems. The complex double curvature of the body resulted 

in a vehicle that would be difficult to fabricate, and further, the body could not 

easily be divided into subassemblies to simplify manufacturing. In addition, the 

lifting body’s large base area yielded a relatively poor subsonic lift-to-drag ratio 

(L/D), resulting in a less attractive cruise capability. Although lifting bodies continued to be studied throughout Phase B, the concept was a dark horse, at best.

At the same time, stowed-wing designs were found to have many attractive features, including low burnout weight and the high hypersonic L/D 

needed to meet the maximum cross-range requirements. In addition, the 

stowed approach permitted the wing to be optimized for the low-speed flight 

regime, simplifying landing. Its drawbacks included a high vehicle-weightto-projected-planform-area ratio, which would result in a higher average base 

temperature relative to either straight- or delta-wing concepts. In addition, the 

mechanisms needed to operate the wing and transmit the flight loads to the 

primary structure resulted in significantly increased design and manufacturing complexity. The maintenance required between flights was expected to be 

high, and insufficient data existed to reliably determine potential failure modes, 

which were thought to be numerous.

Maxime A. Faget, having moved from Langley to MSC in Houston, TX, was 

not a proponent of lifting reentry. Instead, Faget held to the idea of a high-drag 

blunt body. He designed a straight-wing space shuttle design popularly called 

the DC-3 (more officially, the MSC-001), which operated as a blunt body. Faget 

proposed to enter the atmosphere at an extremely high angle of attack with the 

broad lower surface facing the direction of flight, creating a large shock wave that 

187
Coming Home

would carry most of the heat around the vehicle instead of into it. The vehicle 

would maintain this attitude until it got below 40,000 feet and about 200 mph, 

when the nose would come down, and it would begin diving to pick up sufficient 

speed to fly toward the landing site, touching down at about 140 mph. Since the 

only flying the vehicle would perform was at very low speeds during the landing 

phase, the wing design could be selected solely on the basis of optimizing it for 

subsonic cruise and landing; hence the simple straight wing proposed by Faget. 

The design did have one major failing (at least in the eyes of the Air Force): Since 

it did not truly fly during reentry, it had almost no cross range.2

Faget had convinced many in NASA that his simple straight-wing concept would be more than adequate. Nevertheless, not everybody agreed. In 

particular, Charles Cosenza and Alfred C. Draper, Jr., at the Air Force Flight 

Dynamics Laboratory did not accept the idea of building a space shuttle that 

would come in nose high, then dive to pick up flying speed. For most of the 

entry, the vehicle would be in a classic stall, and the Air Force disliked both 

dives and stalls, regarding them as preludes to out-of-control crashes. Draper 

preferred to have the vehicle enter a glide while still at hypersonic speeds, thus 

maintaining much better control while still avoiding much of the severe aerodynamic heating.3

 It should be noted that even Draper agreed with the falling 

concept for the first part of entry to take advantage of the blunt-body theory; 

the difference lay with exactly when to begin flying.

However, if Draper’s vehicle was going to glide across a broad Mach range, 

from hypersonic to subsonic, it would encounter another aerodynamic problem: a shift in the wing’s center of lift. At supersonic speeds, the center of lift 

is located about midway along the wing’s chord (the distance from the leading 

to the trailing edge); at subsonic speeds, it moves much closer to the leading 

edge. Keeping an aircraft in balance requires the application of aerodynamic 

forces that can compensate for this shift. Another MSC design, the Blue Goose, 

accomplished this by translating the entire wing fore to aft as the center of lift 

changed—a heavy and complex solution.

Air Force engineers had found that a delta planform readily mitigated most 

of the problem, so Draper proposed that any space shuttle should use delta 

wings instead of straight ones. Faget disagreed, pointing out that his design did 

not fly at any speed other than low subsonic; at other speeds, it fell and was not 

subject to center-of-lift changes since it was not using lift at all. Faget argued 

 2. Maxime Faget and Milton Silveira, “Fundamental Design Considerations for an Earth-surface-to188
Spaceplane Reality

that a straight wing with a narrow chord would be light and have relatively 

little area that needed a thermal protection system. To achieve the same landing speed, a delta wing would be large, add considerable weight, and greatly 

increase the area that required thermal protection.4

Draper argued that delta wings had other advantages. Since it was relatively 

thick where it joins the fuselage, a delta wing offered more room for landing 

gear and other systems that could be moved out of the fuselage, freeing space 

for a larger payload bay. Its sharply swept leading edge produced less drag at 

supersonic speeds, and its center of lift changed slowly compared to a straight 

wing. In addition, the delta wing offered one other advantage—one that became 

increasingly important as the military became more interested in using the Space 

Shuttle. Compared to a straight wing, a delta wing produces considerably more 

lift at hypersonic speeds. This allowed a returning vehicle to achieve a substantial 

cross range during entry, a capability highly prized by the Air Force. It was a 

complicated problem, and one that would take at least another year to solve.

Phase B Studies

To help resolve the controversy, two concepts were baselined for the Phase 

B study, including a Faget-style straight-wing, low cross-range orbiter, and a 

Draper-supported delta-wing, high cross-range orbiter. The straight-wing orbiter 

would be configured to provide design simplicity, low weight, decent handling, 

and good landing characteristics. The vehicle would be configured to enter at a 

high angle of attack to minimize heating and facilitate the use of a heat shield 

constructed from materials available in the early 1970s. The delta-wing orbiter 

would be designed to provide the capability to trim over a wide angle-of-attack 

range, allowing initial entry at a high alpha to minimize the severity of the heating environment, then transitioning to a lower angle of attack to achieve a high 

cross range. Of course, the entry concept for the Space Shuttle was significantly 

different from that intended for Dyna-Soar. Dyna-Soar was designed to complete 

an entire entry at any angle of attack from about 18 degrees to 55 degrees. This 

resulted in over-design of the thermal protection system for those regions that 

were only critical during the extreme low or high angles of attack. The Space 

Shuttle was designed for a predefined angle-of-attack schedule with a Mach 

number that allowed optimization of the thermal protection system but less 

flexibility in achieving the desired cross range or downrange.

 4. Ibid., pp. 211–212.

189
Coming Home

Four design concepts for the Space Shuttle. NASA/Dennis R. Jenkins.

NASA issued two Phase B contracts on July 6, 1970: one to a team consisting 

of McDonnell Douglas and Martin Marietta, and the other to North American 

Rockwell, which was joined by Convair as a risk-sharing subcontractor. This phase 

involved a detailed analysis of mission profiles and a preliminary look at the R&D 

needed to complete them. The effort would also include a great deal of study on 

thermal protection systems, a critical component in any spaceplane concept.

The Phase B studies were the first comprehensive review of the thermal 

protection system for a large lifting reentry vehicle. These studies included 

extensive investigations into the expected heating environments, and the various materials that might be used on a space shuttle. The earlier Dyna-Soar work 

was largely overcome by the expected, and unrealistic, high flight rate and the 

size and weight of the new vehicle.

Any thermal protection system is defined by the natural and induced environments to which it is exposed during various phases of the mission. Natural 

environments include rain, dust, wind, and micrometeoroids, which may require, 

in certain instances, that additional protective devices (e.g., coatings) be applied to 

the basic heat shield. Induced environments include heating, pressures, and acoustics that result from vehicle aerodynamic design and the selected flight profile.5

 5. D.W. Hass, “Final Report: Refurbishment Cost Study of the Thermal Protection System of a Space Shuttle 

190
Spaceplane Reality

One of the most promising ideas to come out of the Phase A studies was 

to mount the space shuttle thermal protection system so that it could be easily 

removed from the vehicle. This allowed the systems and structure under the 

thermal protection system to be serviced and the thermal protection system itself 

to be refurbished without taking the entire vehicle out of service for a prolonged 

period. In this concept, whatever material the thermal protection system was 

made out of would be mounted on panels, and these panels would be installed 

over the basic vehicle structure with an appropriate insulation between. When 

the vehicle returned from orbit, the panels would be replaced; the old panels 

would be sent to a shop for refurbishment, and then used on a future mission. 

This made sense at the time since it appeared that the most likely materials were 

ablators and refractory metals, both of which would need significant maintenance 

between flights.

Ablators were the departure point for most of the thermal protection system 

studies because they were well understood and proven. Researchers studied 

various materials, including silica/silica-fiber composites; mixed inorganic or 

organic composites with silica, nylon, or carbon fiber-reinforced resins (phenolics, epoxies, and silicones); and carbon- or graphite-based materials. These 

ablators had densities ranging from 10 pounds per cubic foot (using microballoons to keep the material lightweight) to 150 pounds per cubic foot (for 

high-heat protection). Researchers evaluated both charring and noncharring 

materials, but the need to maintain a precise outer mold line for aerodynamic 

efficiency usually mandated charring ablators.

The primary application technique for elastomeric materials was spraying, 

although the experience of applying MA-25S ablator on the X-15A-2 had been 

less than ideal. The ablator could be sprayed directly on primary structures and 

cured (as with the X-15A-2) or sprayed on panels and attached by fasteners. 

A major problem with spraying an ablator directly onto the vehicle was that 

most ablators required curing at elevated temperatures, meaning a very large 

oven would need to be constructed.6

One of the bigger unknowns with ablator heat shields was how to refurbish 

them. Researchers had intended to investigate this aspect of ablators using 

the X-15A-2 at Edwards AFB, but the first high-speed flight with the ablative 

coating resulted in substantial damage to the airplane, and further flight tests 

were cancelled. In truth, the damage was not a failure of the ablator but an 

engineering flaw that failed to recognize the potential for shock impingement 

on an experimental ramjet that was hung below the ventral stabilizer. In addition, the X-15 used a hot structure to tolerate the heat on its flights, and the 

 6. Ibid.

191
Coming Home

ablator interfered with the normal radiance and absorption of heat that the 

structure depended on. It was not a happy experience.7

Refurbishing ablators requires the removal of the used ablator and replacing it with new material. For a panel design, removing the used panels and 

substituting new ones minimized vehicle turnaround time. The used panels 

could then be refurbished or discarded. Direct application to the vehicle skin 

required that the used ablator be removed from the primary structure and 

new ablator applied, likely increasing vehicle turnaround time. In either case, 

elastomeric ablator facilitated refurbishment, as it was soft and easily cut with 

a knife or other sharp edge, layer by layer, down to the substrate.8

However, elastomeric ablator refurbishment did not necessarily require the 

complete removal of the used ablator. It would be possible to apply a thick 

enough coat of ablator to last for several flights. After the first flight, only the 

charred or spent ablator would be removed—a relatively simple task. However, 

this imposed an initial weight penalty. Another approach was to refurbish the 

ablator by removing only the charred or spent ablator down to virgin material. 

New ablator could then be applied over the virgin material. This was how the 

X-15 program intended to refurbish the X-15A-2. The thickness was measured 

by pins embedded in the ablator at specified intervals, followed by manual 

sanding until the desired height was attained—a tedious, manual process. 

Obviously, a better technique would need to be developed.

Dyna-Soar had eschewed ablators in favor of a metallic heat shield, and this 

seemed an elegant solution to most researchers. The use of radiative metallic 

panels supported by a cool structural shell with a layer of insulation between 

would work for all but the hottest regions of the vehicle. Metals are intrinsically 

durable materials and therefore capable of extensive reuse, but they have oxidation or strength limitations that needed resolution prior to using them in an 

operational thermal protection system. However, temperature limitations are 

largely determined by the material’s ability to resist oxidation. Unfortunately, 

most of the exotic and superalloys tested for these programs tended to oxidize quickly when exposed to high temperatures. To minimize this oxidation, 

researchers developed coatings that could be applied to the metal to shield it 

from the effects of the atmosphere. At the time, coatings had been developed 

that appeared to permit an upper-limit temperature of about 2,500 °F for 100 

cycles, but actual real-world data was missing.9

 7. Dennis R. Jenkins, 

192
Spaceplane Reality

A fully reusable Space Shuttle design concept from 1970. NASA.

The most likely candidates for metallic radiative heat shields included titanium, nickel- and cobalt-based superalloys, thoria dispersed nickel-chromium 

(TD NiCr), and coated columbium alloys. Typically, nickel alloys and cobalt 

alloys were used in environments in which temperatures exceeded 1,000 °F. 

Due to their chemical compositions, these materials tended to be quite expensive, with long lead times from the mills. Some of the more well-known trademarked alloys included Inconel®, Hastelloy®, Monel®, Nitronic®, CBS600, 

Haynes 188®, L605, and René 41®.

Selecting the upper-use temperature limit for a metallic material involved 

the evaluation of creep strength, metallurgical stability, rate of oxidation, life 

of the oxidation-protective coating, and effect of oxidation-protective coatings 

193
Coming Home

on the mechanical properties of the base materials. In general, titanium alloys 

were shown to be effective below 900 °F, and if the panels were exposed to only 

minor pressure loads, short excursions to environments up to 1,200 °F could 

be tolerated. For the large surface areas subjected to temperatures between 

1,200 °F and 1,800 °F, conventional nickel- and cobalt-based superalloys 

proved efficient and economical.10

Researchers had more difficultly finding suitable material for temperature 

ranges between 1,800 to 2,200 °F. Conventional superalloys proved structurally 

inefficient above 1,800 °F, while coated columbium was not weight efficient 

below 2,200 °F. Researchers believed that TD NiCr alloy was the most promising candidate to fill this temperature range. In addition to being lighter, a 

major advantage of TD NiCr over columbium alloys was that it did not require 

protective oxidation coatings. However, at the time, TD NiCr was a relatively 

new material and very little information was available regarding its use in a 

thin-gauge reusable heat shield application.11

New materials, including hardened compacted fibers (HCF) and oxidationinhibited carbon-carbon, which were neither ablative nor metallic, presented 

engineers with an entirely different set of options. The most common HCF 

investigated during the early studies was called mullite, though it was not truly 

mullite (a rare chemical that is found only on the Isle of Mull off the west 

coast of Scotland). Mullite had long been used as a refractory material in the 

glass and steel industries since it exhibited high temperature strength, good 

thermal-shock resistance, and excellent thermal stability. Since mullite was 

rare, industry researched and discovered methods of using various materials 

to fabricate synthetic mullite ceramics.12

HCF was characterized by a layer of rigidized inorganic fibers that combined 

functions of a high-temperature radiative surface and an efficient insulation. 

Researchers believed that HCF was a promising candidate for space shuttle 

applications because of its potential weight savings (it weighs only 12–15 pounds 

per cubic foot), availability, and temperature-overshoot capability. These HCF 

materials were relatively soft, extremely porous, and had inherently low emittance values. Adhesive bonding had to be used since HCF material would not 

support mechanical fasteners due to structural weakness. The materials that were 

candidates to produce HCF included aluminosilicate, silica, aluminosilicatechromia, artificial mullite, and zirconia. Researchers believed these materials 

 10.

194
Spaceplane Reality

were potentially useful for 100 flights in the range of 2,000 to 3,000 °F. They 

also believed that the development of suitable reusable coatings was the most 

critical problem that had to be solved before the material could be successfully 

used on any space shuttle. In many ways, the mullite shingles were an indication 

of things to come as the ceramic Lockheed insulation was perfected.13

A relatively new class of materials, all of which benefited significantly from 

the investment in Dyna-Soar, was also extensively investigated. Researchers 

believed that oxidation-inhibited, fiber-reinforced, carbonaceous materials 

offered the potential for a reusable, reasonably cost efficient, high-temperatureresistant structural material. The strength-to-weight ratio of these materials 

at high temperatures was unsurpassed, leading to its use on the wing leading 

edge and nose cap, where aerodynamic forces were highest. The carbonaceous 

material was reinforced by carbon or graphite fibers, usually in the form of 

cloth. It should be noted that these materials, like the reinforced carbon-carbon 

actually used on the Space Shuttle, were not insulators, and the backside of the 

material was essentially as hot as the frontside. What the materials provided was 

a means of creating the appropriate outer mold line on the hottest parts of the 

vehicle (the nose and wing leading edge), but some sort of insulating material 

was still needed to protect the primary structure. The insulating material was 

usually a fibrous material characterized by low strength, low density, and low 

thermal conductivity. Insulators typical of this class include Q-Felt®, MicroFiber®, and Dynaflex® batt insulation.

Reusable Surface Insulation

However, a new material was just being developed. Although it was similar in 

many respects to the artificial-mullite shingle investigated during Phase B, this 

new material represented a quantum leap in technology. During the original 

studies of lifting reentry vehicles in the late 1950s, there had been a great debate 

over the relative merits of active cooling systems versus passive systems for thermal protection. The active systems were attractive on paper but nobody could 

quite figure out how to make them work at a reasonable weight. Therefore, the 

choices were largely narrowed to either a hot structure, like the X-15, or a more 

conventional structure protected by some sort of insulation. The hot-structure 

approach required the use of rare and expensive superalloys, and there was 

always a great deal of doubt if it would have worked on a vehicle as large as a 

space shuttle. Generally, most preferred a fairly conventional structure made 

 13.

195
Coming Home

of titanium or aluminum protected by a series of metallic shingles with a thick 

layer of insulation in between the two. There was some investigation of ablative 

coatings, but the unhappy X-15A-2 experience made just about everybody shy 

away from this technology except as a last resort.

Things began to change as the Lockheed Missiles and Space Company 

made quick progress in the development of a ceramic reusable surface insulation (RSI) concept, and by December 1960 Lockheed had applied for a patent 

for a reusable insulation material made of ceramic fibers. The first use for the 

material came in 1962 when Lockheed developed a 32-inch-diameter radome 

for the Apollo spacecraft made from a filament-wound shell and a lightweight 

layer of internal insulation cast from short silica fibers. However, the Apollo 

design changed, and the radome never flew.

Nevertheless, this experience led to the development of a fibrous mat, 

called Lockheat, that had a controlled porosity and microstructure. The mat 

was impregnated with organic fillers such as methylmethacylate (plexiglass) 

to achieve a structural quality. These composites were not ablative, since they 

did not char to provide protection. Instead, Lockheat evaporated, producing an 

outward flow of cool gas. By 1965, this had led to the development of LI-1500, 

the first of what became the Space Shuttle tiles. This material was 89-percent 

porous, had a density of 15 pounds per cubic foot (hence the 15 in the designation), was capable of surviving repeated cycles to 2,500 °F, and appeared to 

be truly reusable. A test sample was flown on an Air Force reentry test vehicle 

during 1968, reaching 2,300 °F with no apparent problems.

Lockheed engineers continued developing the silica RSI, and Lockheed 

decided to produce the material in two densities to protect different heating 

regimes: 9 pounds per cubic foot (designated LI-900) and 22 pounds per 

cubic foot (LI-2200). The ceramic consisted of silica fibers bound together 

and sintered with other silica fibers, then glaze-coated by a reaction-cured 

glass consisting of silica, boron oxide, and silicon tetraboride. Since this mixture was not waterproof, a silicon polymer was coated over the undersurface 

(i.e., unglazed) side. The material was very brittle, with a low coefficient of 

linear thermal expansion, and therefore Lockheed could not cover an entire 

vehicle with a single piece of it. Rather, the material needed to be installed 

in the form of small tiles, generally 6-by-6-inch squares. The tiles would 

have small gaps between them to permit relative motion and to allow for the 

deformation of the metal structure under them due to thermal effects. Since 

a tile would crack as the metal skin under it deformed, engineers decided to 

bond the tile to a felt pad and then bond the felt pad to the skin. Both of 

these bonds were accomplished with a room temperature vulcanizing (RTV) 

adhesive. It was a breakthrough, although everybody involved approached 

the concept with caution.

196
Spaceplane Reality

This schematic of the 

Space Shuttle shows the 

major components of the 

vehicle. NASA 0102619.

197
Coming Home

Phase C/D Developments

It appears to be the fate of winged spacecraft that their development is shrouded 

in political maneuvering and controversy. So it had been with Dyna-Soar, and 

so it was with the Space Shuttle. After long and heated battles within NASA, 

the Department of Defense, the White House, the Office of Management 

and Budget, and Congress, President Richard M. Nixon finally approved the 

development of the Space Shuttle on January 5, 1972.14

The RFP was issued on March 17, 1972, to Grumman/Boeing, Lockheed, 

McDonnell Douglas/Martin Marietta, and North American Rockwell for the 

production and initial operations of the Space Shuttle system. The technical 

proposals were due on May 12, 1972, with the associated cost and contractual 

data due a week later. By this time, the Faget low cross-range concept had been 

abandoned, and the Government baseline orbiter, called MSC-040C, had a 

large delta wing that met all of the Air Force’s requirements.15

The RFP stated that each orbiter should have a useful life of 10 years and 

be capable of up to 500 missions. However, the RFP asked each contractor 

to provide information on the impact of lowering this to only 100 missions, 

a figure that was subsequently adopted. Three reference missions were specified: (1) lifting 65,000 pounds of payload into a 310-mile, due-east orbit 

from the Kennedy Space Center (KSC); (2) 25,000 pounds into a 310-mile, 

55-degree orbit from KSC while carrying air-breathing landing engines; and 

(3) 40,000 pounds into a 115-mile, circular polar orbit from Vandenberg 

AFB, CA. The orbiter was to have the capability to return to its launch site 

after a single orbit, although an actual cross range was not specified. An 

unspecified RSI thermal protection system was to be used, although ablative 

material or other special forms of thermal protection systems were allowed 

where beneficial to the program. The orbiter thermal protection system was 

to be capable of surviving an abort scenario from a 500-mile circular orbit.16

Unsurprisingly, the four contractor-proposed vehicles looked remarkably 

similar, since all were based on the MSC-040C concept. Grumman and North 

American were the only competitors that designed the orbiter for the baseline 

500-mission service life; the others opted for the 100-mission alternate.17

 14.

198
Spaceplane Reality

The thermal protection system of the Space Shuttle relied on ceramic tiles that were especially 

heat resistance but were fragile and required considerable care to ensure their successful use. It 

also used reinforced carbon-carbon on the wing leading edge and nose cap. This drawing of the 

orbiter wing shows the structure of this critical component of the vehicle. NASA.

Grumman did not completely trust the Lockheed-developed RSI and proposed adding a layer of ablative material around the crew compartment and the 

orbital maneuvering system (OMS)/reaction control system (RCS) modules 

during the early development flights to guard against a possible RSI burnthrough. After landing at KSC, the Grumman orbiter would be purged and 

secured, towed to a new refurbishment facility where the thermal protection 

system would be repaired as necessary, then sprayed to maintain the thermal 

protection properties and prevent water absorption.18

The Lockheed Phase C/D proposal took some liberties with the basic 

MSC-040C design. Specifically, the wing planform showed a decided bias 

toward a double-delta design long before the program officially sanctioned this 

change. Surprisingly, the proposal did not specifically address why Lockheed 

had chosen this design, other than it slightly simplified the thermal protection 

 18.

199
Coming Home

system. For instance, Lockheed proposed LI-1500 tiles for the inboard part of 

the leading edge, using the more expensive carbon-carbon system only on the 

outboard sections. The overall orbiter thermal protection system was designed 

to provide 100 normal operational entries at 2,500 °F, or a single-contingency 

reentry at 3,000 °F.19

McDonnell Douglas remained fairly true to the original MSC-040C design, 

but its thermal protection system was not the expected Lockheed-developed 

RSI silica tiles. Instead, McDonnell Douglas proposed using mullite HCF 

shingles20 that were glued to the bottom and side of the orbiter over a thin layer 

of foam for strain relief. The top of the orbiter would use ESA-3560 or SLA561 ablator, as would the leading edges of the wing and vertical stabilizer.21

Like Grumman, the engineers at North American Rockwell projected a 

500-mission life for the orbiter. To accomplish this, the company selected a 

thermal protection system designed to subject the primary structure to no 

more than 350 °F. North American was also not completely convinced that 

the Lockheed-developed silica tiles would mature quickly enough for the 

early orbital flights, so the baseline used mullite HCF shingles instead. North 

American continued to monitor the progress being made on the silica tiles since 

they were lighter weight, and the contractor expected to switch to them at an 

appropriate time. The leading edges and nose cap were carbon-carbon, and no 

substantial use of ablators was made anywhere on the orbiter.22

Orbiter Contract Award

James Fletcher, George Low, and Richard McCurdy met the morning of July 

26, 1972, for the final review of the Space Shuttle orbiter proposals.23 They 

noted that the evaluation team had ranked McDonnell Douglas and Lockheed 

significantly below the other two competitors, so they concentrated instead on 

North American and Grumman. After carefully reviewing the mission suitabil-

 19.

200
Spaceplane Reality

ity scores, the three men decided that the advantage went to North American. 

Since North American also presented the lowest probable cost, it was deemed 

the winner of the contract.24

After the stock market closed on July 26, 1972, NASA announced that the 

Space Transportation Systems Division of North American Rockwell25 had been 

awarded the $2.6 billion contract to design and build the Space Shuttle orbiter.26

The design was loosely based on the final MSC-040C configuration, and during 

its return to Earth, the vehicle would be capable of a 1,265-mile reentry maneuver on either side of its flightpath, thereby meeting Air Force requirements.27

The Space Shuttle represented the largest hypersonic vehicle ever designed, 

providing the first real test of experimental and theoretical knowledge gained 

during the X-15, ASSET, and PRIME programs. No real precedents existed to 

help establish the design requirements for such a vehicle, something the flighttesting of Dyna-Soar would have significantly helped. A primary challenge for 

the designers was the preflight prediction of the aerodynamic characteristics 

of the vehicle with sufficient accuracy to conduct the first orbital flight with 

a crew. This also required overcoming the unknowns involved in hypersonic 

wind tunnel testing, and developing reliable simulation techniques without 

having an extensive database to compare the results.28

In its Phase C response, Rockwell proposed using mullite shingles manufactured from aluminum silicate instead of the Lockheed-developed silica tiles 

since the technology was better understood and more mature. But the mullite 

tiles were heavier and potentially less durable. Given the progress Lockheed 

was making, NASA and Rockwell asked the Battelle Memorial Institute to 

evaluate both candidates’ systems—an evaluation the Lockheed product won. 

Nevertheless, the Lockheed material was not appropriate for the very high-temperature areas of the orbiter, such as the nose cap and wing leading edge. These 

 24.

201
Coming Home

would use a reinforced carbon-carbon (RCC) material originally developed by 

Ling-Temco-Vought (LTV), based on research for the Dyna-Soar program.29

Interestingly, NASA and Rockwell originally believed that the leeward side 

(top) of the orbiter would not require any thermal protection. However, in 

March 1975, the AFFDL conducted a briefing for Space Shuttle engineers on 

the classified results of the ASSET, PRIME, and boost-glide reentry vehicle 

programs, which indicated leeward-side heating was a serious consideration. 

The thermal environment was not particularly severe but easily exceeded the 

350 °F capability of the aluminum skin. To alleviate this concern, Rockwell 

would cover the upper surfaces of the wings and aft fuselage, along with the 

entire payload bay door (areas that never exceeded 650 °F) with an ablative 

elastomeric reusable surface insulation (ERSI). The ERSI, which consisted of 

a silicone resin with titanium dioxide and carbon-black pigments for thermal 

control, would be bonded directly to the aluminum airframe.

Between April 1978 and January 1979, a team from the AFFDL conducted 

a review of potential orbiter heating concerns and concluded that the OMS 

pod might have unanticipated problems. Air Force researchers conducted a 

series of tests at the Arnold Engineering Development Center between May 

and November 1979, and further tests were run at the Naval Surface Warfare 

Center shock tunnel in May 1980. The researchers discovered that the composite OMS pod skin would deflect considerably more than expected. The thin 

8-by-8-inch RSI tiles were relatively weak under bending loads, and engineers 

feared that the tiles might fracture and separate from the vehicle. Since the tiles 

had already been installed, a unique solution was developed in which each tile 

was diced while still attached to the pod by carefully cutting it into nine equal 

parts. The technique proved so successful that it has subsequently been used 

on other areas of the orbiter.30

The proposed thermal protection system at this point differed somewhat 

from the system ultimately used on the production orbiters. ERSI ablator 

covered the upper surfaces of the wings and payload bay doors, where temperatures were always less than 650 °F. Black LI-2200 high-temperature reusable surface insulation (HRSI) tiles protected the bottom of the fuselage and 

 29.

202
Spaceplane Reality

This 1979 structural test article of the Space Shuttle wing shows the reinforced carbon-carbon 

and tile structure. NASA.

203
Coming Home

wings, the entire vertical stabilizer, and most of the forward fuselage where 

temperatures exceeded 650 °F but were less than 2,300 °F. The tile thickness 

varied as necessary to limit the aluminum airframe’s maximum temperature to 

350 °F. Actually, earlier studies showed a combination of LI-2200 and LI-3000 

tiles would be required, although this was later reduced to entirely LI-2200, 

then to a combination of LI-900 and LI-2200.31 Further study showed some 

drawbacks to this protection system, and when Columbia was manufactured 

the ERSI would be replaced by a combination of LI-900 tiles and Nomex felt 

blankets. The HRSI tiles were used over a smaller percentage of the orbiter than 

originally anticipated as further analysis showed that entry heating would not 

be as severe as expected, mainly because of improved flight profiles.32

In the meantime, a problem had developed with the tiles themselves. As flight 

profiles were refined and aeroloads better understood, engineers began to question whether the tiles could survive the harsh conditions. By mid-1979 it had 

become obvious that the tiles in certain areas did not have sufficient strength to 

survive the tensile loads of a single mission. NASA immediately began an extensive search for a solution that eventually involved outside blue-ribbon panels, 

Government agencies, academia, and most of the aerospace industry. As Space 

Shuttle Program Deputy Director LeRoy Day recalled, “There is a case [the tile 

crisis] where not enough engineering work, probably, was done early enough in 

the program to understand the detail–the mechanical properties–of this strange 

material that we were using. It was, potentially, a showstopper.”33

As engineers spent more time examining the problem, they realized that 

the issue was not with the individual tiles but the manner in which the tiles 

were attached to the underlying structure. Analysis indicated that while each 

individual component—the tile, the felt strain-isolation pads (SIP) under the 

tile, and the two layers of adhesives—had satisfactory tensile strength, the 

components when combined as a system lost about 50 percent of their strength. 

Engineers largely attributed this to stiff spots in the SIP (caused by needling) 

that allowed the system strength to decline to as low as 6 pounds per square 

inch, instead of the baseline 13 pounds per square inch.34

 31.

204
Spaceplane Reality

The Space Shuttle structural test article undergoing tests in the Lockheed facility in Palmdale, CA. 

This airframe, STA-099, was subsequently refitted for orbital flight and became the Challenger 

(OV-099). NASA.

In October 1979, engineers developed a densification process that involved 

filling voids between fibers at the inner mold line (the part next to the SIP 

pad) with a special slurry mixture consisting of Ludox (a colloidal silica made 

by DuPont) and a mixture of silica and water. Since the tiles had been waterproofed during manufacture, the process began by applying isopropyl alcohol 

to dissolve the waterproof coating, then painting the back of the tile with 

Ludox. After air-drying for 24 hours, the tiles were baked in an oven at 150 °F 

for 2 hours, then re-waterproofed using Dow Coming’s Z-6070 methyltrimethoxysilane.35 The densified layer acted as a plate on the bottom of the tile, 

eliminating the effect of the local stiff spots in the SIP, bringing the total system 

strength back up to 13 pounds per square inch.36

Installing the tiles on the vehicle presented its own problems, and Rockwell 

quickly ran out of time to do so while Columbia was in Palmdale, CA. NASA 

 35.

205
Coming Home

needed to present the appearance of maintaining a schedule, and Columbia 

moving to KSC was a very visible milestone. So in March 1979, Columbia 

was flown from Palmdale to KSC on the 747 Shuttle Carrier Aircraft and 

was quickly moved into the Orbiter Processing Facility. Just over 24,000 

tiles had been installed in Palmdale, with 6,000 left to go. Unfortunately, by 

now it appeared that all of the tiles would need to be removed so that they 

could be densified.

The challenge revolved around how best to salvage as many of the installed 

tiles as possible while ensuring sufficient structural margin for a safe flight. The 

approach that engineers developed to overcome this almost insurmountable 

challenge was called the tile proof test. Technicians induced a stress over the 

entire footprint of each tile equal to 125 percent of the maximum flight stress 

experienced at the most critical point on the tile footprint.37

The device developed for the proof test used a vacuum chuck to attach to 

the tile, a pneumatic cylinder to apply the load, and six pads attached to surrounding tiles to react the load. Since any appreciable tile load might cause 

internal fibers to break, acoustic sensors placed in contact with the tiles were 

used to monitor the acoustic emissions for any internal fiber breakage. This 

testing proved the majority of the tiles were adequately bonded for flight, and 

only 13 percent failed the proof test; those were replaced with densified tiles.38

Engineers also developed two other techniques to strengthen tiles while 

they were still on the vehicle. The first involved relatively small, thick tiles that 

were usually located on the underside of the orbiter. As shock waves swept air 

over the tiles, they tended to rotate, inducing high stresses at the SIP bond. 

Technicians installed gap fillers that prevented the tiles from rotating, but 

this solution was not effective for small thin tiles, so engineers developed a 

technique in which the filler bar surrounding the SIP was bonded to the tiles. 

Technicians inserted a crooked needle into the tile-to-tile gap and posited 

RTV adhesive on top of the filler bar. This significantly increased the bonded 

footprint and decreased the effects of a shock-imposed overturning moment.

For the next 20 months, technicians at KSC worked 3 shifts per day, 7 days 

per week testing and installing 30,759 tiles. By the time the tiles were installed, 

proof tested, often removed and reinstalled, then re-proof tested, the technicians averaged 1.3 tiles per person per week. In June 1979, Rockwell estimated 

that 10,500 tiles needed to be replaced; by January 1980, over 9,000 of these 

had been installed, but the number remaining had ballooned to 13,100 as 

additional tiles failed the proof test or were otherwise damaged. By September 

 37.

206
Spaceplane Reality

During the flight of STS-1 in April 1981, Columbia shed some of its thermal protection system 

tiles, as shown in this orbital imagery. The vehicle returned safely. NASA.

1980, only 4,741 tiles remained to be installed, and by Thanksgiving, the 

number was below 1,000. It finally appeared that the end was in sight.39

Thermal Protection System

The thermal protection system protected the aluminum structure of the orbiter 

during the atmospheric portion of flight (both ascent and entry), maintaining 

skin temperatures at less than 350 °F. The chosen thermal protection system 

was completely passive in operation and, in theory, capable of unlimited reuse 

with only minor refurbishment. It provided a smooth aerodynamic mold line 

using lightweight materials with low thermal conductivity.40 These materials had to perform in temperature ranges from –250 °F in the cold soak of 

space to entry temperatures approaching 3,000 °F. The peak heating rates and 

 39.

207
Coming Home

longest exposures occurred during high cross-range entry maneuvers when 

equilibrium surface temperatures range from 3,000 °F, at stagnation points 

on the nose and leading edges of the wing and tail down to 600 °F on leeward 

surfaces of the wing.41

The thermal protection system was composed of RCC, four types of RSI 

tiles, two types of RSI blankets, thermal barriers, gap fillers, thermal windowpanes, and thermal seals.42

Reinforced Carbon-Carbon

Where temperatures exceeded 2,300 °F, such as the wing leading edge and 

the nose cap, the vehicle featured a composite composed of pyrolyzed carbon 

fibers in a pyrolyzed carbon matrix with a silicon carbide coating. RCC 

weighs 90 to 100 pounds per cubic foot. The operating range of RCC is from 

–250 °F to 3,000 °F. The carbon-carbon nose cap and wing leading edges were 

manufactured by LTV (now Lockheed Martin) of Grand Prairie, TX.43 The 

wing leading edge consisted of 22 RCC panels, and because of the wing profile 

changes from inboard to outboard, each RCC panel was unique.44 It should 

be noted that RCC was not an insulator; it was simply a structure that could 

withstand the appropriate aero and thermal loads. The underlying structure 

needed to be protected from extreme heat by other means.

RCC fabrication begins with a graphite-saturated rayon cloth impregnated 

with a phenolic resin. This impregnated cloth is layed up as a laminate and 

cured in an autoclave. After curing, the laminate is pyrolyzed to convert the 

resin to carbon. This is then impregnated with furfural alcohol in a vacuum 

chamber, then cured and pyrolyzed again to convert the furfural alcohol to 

carbon. This process is repeated three times until the desired carbon-carbon 

properties are achieved.45

 41.

208
Spaceplane Reality

To provide sufficient oxidation resistance to enable reuse, the outer layers 

of the RCC are converted to silicon carbide. The RCC is packed in a retort 

with a dry-pack material made of a mixture of alumina, silicon, and silicon 

carbide. The retort is placed in an argon-purged furnace with a stepped-timetemperature cycle up to 3,200 °F. A diffusion reaction occurs between the dry 

pack and carbon-carbon that converts the outer layers of the carbon-carbon 

to silicon carbide (whitish-gray color). The part is then sprayed with tetraethyl 

orthosilicate and sealed with a glossy overcoat. The RCC laminate is superior to 

a sandwich design because it is lightweight and rugged; it also promotes internal cross-radiation from the hot stagnation region to cooler areas, thus reducing stagnation temperatures and thermal gradients around the leading edge.

The RCC panels were mechanically attached to the wing leading edge with 

a series of floating joints to reduce loading on the panels caused by wing 

deflections. A T-seal between each panel allowed lateral motion and thermalexpansion differences between the RCC and the orbiter wing. The T-seals were 

also constructed of RCC.

Since RCC was highly radiative and provided no thermal protection, the 

adjacent aluminum structure had to be protected by internal insulation. The forward wing spar and nose bulkhead were protected by a series of captive and noncaptive RSI tiles that were installed onto removable carrier panels with Inconel 

601–covered Dyna-Flex (Cerachrome) bulk insulation (also called Incoflex).46

Beginning in 1998, engineers began modifying the leading-edge RCC to 

enable it to withstand larger punctures. The original design allowed a 1-inch 

hole in the upper surface of any panel. But on the lower surface, no penetrations 

were allowed on panels 5 through 13 because any hole generated by orbital 

debris (or other causes) would allow heat from the plasma flow during entry to 

quickly erode the 0.004-inch Inconel foil on the Incoflex insulators, exposing 

the leading-edge attach fittings and wing-front spar to the direct blast of hot 

plasma. The upgrade included additional insulation that was able to withstand 

a penetration of up to 0.25 inch in diameter in the lower surface of panels 9 

through 12, and up to 1 inch on panels 5 through 8 and 13. To achieve this, 

Nextel 440 fabric insulation was wrapped around the Incoflex insulators—one 

layer for panels 5 through 7 and 11 through 13, and two fabric layers for panels 

8 through 10 (which have the highest-potential heating environment).47

Despite the modification, an ascent debris strike on an RCC panel on the 

left wing of Columbia during STS-107 resulted in the destruction of the orbiter 

 46.

209
Coming Home

There were difficulties with burn-throughs on the reinforced carbon-carbon leading edge, as 

shown in this comparison of STS-2 and STS-3 panels in 1982. NASA.

during entry. This led to a reevaluation of the relative strength of the material, 

and subsequent testing revealed that it was not as robust as once thought, particularly to direct debris strikes. Nevertheless, it was one of the few materials 

that can withstand the environment it is subjected to, and it continued to be 

used, although with somewhat more inspection and analysis.

Reusable Surface Insulation Tiles

The RSI tiles used on the orbiter had three different densities of the base silica 

material. LI-900 weighed 9 pounds per cubic foot and comprises 85 percent of 

the orbiter’s tiles. Denser LI-2200 tiles weighed 22 pounds per cubic foot and 

were used around door and access panels, making up approximately 1 percent 

of the orbiter’s tiles. FRCI-12 tiles weighed 12.5 pounds per cubic foot and 

made up approximately 12 percent of the orbiter’s tiles. The secret to the tiles’ 

performance was the fact that they were 80 to 90 percent void (contain air) 

and minimized heat transfer within the tile itself.48

 48.

210
Spaceplane Reality

Black HRSI tiles were used on the entire underside of the vehicle where 

RCC was not used, including the following surfaces: the base heat shield; 

selected areas on the upper forward fuselage, such as around the forward fuselage windows; the portions of the OMS/RCS pods; the leading and trailing 

edges of the vertical stabilizer; the wing-glove areas; the elevon trailing edges; 

the area adjacent to the RCC on the upper wing; the interface with wing leading edge RCC; and the upper body flap. The HRSI tiles protected areas where 

temperatures were between 1,200 °F and 2,300 °F.

The HRSI tiles were made of low-density, high-purity, 99.8-percent amorphous silica fibers derived from common sand; they were 1 to 2 millimeters 

thick and were made rigid by ceramic bonding. A slurry containing fibers 

mixed with water was frame-cast to form soft, porous blocks onto which a 

colloidal silica binder solution was added. When it was sintered, a rigid block 

was produced that was cut into quarters and then machined to the precise 

dimensions required for individual tiles. HRSI tiles varied in thickness from 1 

to 5 inches, with the thickness determined by the heat load encountered during 

entry. Generally, the HRSI tiles were thicker at the forward areas of the orbiter 

and thinner toward the aft end. The HRSI tiles were nominally 6-by-6-inch 

squares, but they varied in size and shape in the closeout areas. The HRSI tiles 

withstood on-orbit cold soak conditions, repeated heating and cooling thermal 

shock, and extreme acoustic environments (163 decibels) at launch.49

The HRSI tiles were coated on the top and sides with a mixture of powdered tetrasilicide and sintered borosilicate reaction-cured glass (RCG) coating 

with a liquid carrier. The tiles were individually sprayed with 10 to 15 coats of 

the coating slurry to produce a final coating weight of 0.09 to 0.170 pounds 

per square foot. The coated tiles then were placed in an oven and heated to 

2,300 °F for 90 minutes. This resulted in a black, glossy coating that had a 

surface emittance of 0.85 and a solar absorptance of about 0.85. Then, the 

tiles were waterproofed using heated methyltrimethoxysilane. The silane vapor 

penetrated the tile, rendering the fibrous silica material hydrophobic.50

As an aside, an HRSI tile taken from a 2,300 °F oven could be immersed 

in cold water without damage. Also, surface heat dissipated so quickly that an 

uncoated tile could be held by its edges with an ungloved hand only seconds 

after removal from the oven while its interior still glowed red.

As the higher density HRSI tiles were replaced, new fibrous refractory 

composite insulation (FRCI) tiles were usually used instead. Researchers at 

NASA Ames developed this product later in the program, and it was more durable 

 49.

211
Coming Home

When an orbiter returned from a mission, technicians carefully inspected and replaced thermal 

protection system tiles as needed. Although the tiles required significant maintenance early in 

the program, after a few years they became very reliable, and late in the program only a handful 

of tiles were replaced after a mission. NASA.

and resistant to cracking than the earlier HRSI. The tiles were essentially similar 

to the original HRSI tiles, with the addition of a 3M Company additive called 

Nextal (AB312—alumina-borosilicate fiber). With an expansion coefficient 

10 times that of the 99.8-percent pure silica, Nextel activated a boron fusion, 

which effectively welded the pure silica fibers into a rigid structure during 

sintering. The resulting composite-fiber refractory material, composed of 

20 percent Nextal and 80-percent pure silica, had entirely different physical 

properties than the original 99.8-percent pure silica tiles.

The RCG (black) coating of the FRCI-12 tiles was compressed as it was 

cured to reduce the coating’s sensitivity to cracking during handling and operations. In addition to the improved coating, the FRCI-12 tiles were about 10 

percent lighter than the HRSI tiles for a given thermal protection. FRCI-12 

tiles have demonstrated a tensile strength at least three times greater than that 

of the HRSI tiles and a use temperature approximately 100 °F higher.

White low-temperature reusable surface insulation (LRSI) tiles were used 

in selected areas of the forward fuselage, midfuselage, aft fuselage, vertical 

stabilizer, upper wing, and OMS/RCS pods. These tiles protected areas where 

temperatures were between 750 °F and 1,200 °F, and they had a white surface 

coating to provide better thermal characteristics while on orbit.

212
Spaceplane Reality

LRSI tiles were of the same construction and had the same basic characteristics as the HRSI tiles but were cast thinner (0.2 to 1.4 inches). The tiles were 

produced in 8-by-8-inch squares and had a white RCG optical and moistureresistant coating applied 10 millimeters thick to the top and sides. The coating 

was made of silica compounds with shiny aluminum oxide to obtain optical 

properties, and it provided on-orbit thermal control for the orbiter. The LRSI 

tiles were treated with bulk waterproofing and were installed on the orbiter in 

the same manner as the HRSI tiles. The LRSI tile had a surface emittance of 

0.80 and a solar absorptance of 0.32.

A new high-temperature tile, known as toughened unipiece fibrous insulation (TUFI), was developed by NASA Ames in 1993 and was first flown on 

STS-59. The results were encouraging enough for TUFI tiles to be installed 

on an attrition basis to replace 304 HRSI tiles on the base heat shield around 

the Space Shuttle main engine and lower body-flap surface. TUFI was the 

first of a new type of composite known as functional gradient materials in 

which the density of the material is relatively high on the outer surface and 

becomes increasingly lower within the insulation. Unlike the HRSI tiles, in 

which the tetrasilicide and borosilicate glass coating received little support from 

the underlying tile, TUFI’s outer surface was fully integrated into the insulation, resulting in a more damage-resistant tile. TUFI tends to dent instead 

of shatter when hit. This tile material is also known as an alumina-enhanced 

thermal barrier (AETB-8), or TUFI/RCG-coated tiles.

Regardless of type, every tile was unique to a location on the orbiter, cut 

to size, thickness, and shape to conform to the contours of the vehicle and the 

necessary insulative properties. Using a commercial high-temperature paint 

called Spearex, the tiles were then permanently marked with a part number 

that designated its location on the orbiter (X/Y coordinate) and with a unique 

serial number that provided traceability by lot number to the materials and 

processes affecting the individual tile. Also, the part-number identification 

was located on the forward or leading edge of the tile to assist in the correct 

orientation of the tile at the time of installation. Because of discontinuities in 

the thin RCG coating, caused either by processing or flight damage, the tiles 

required re-waterproofing after each flight. Each tile was pneumatically injected 

with dimethylethoxysilane, which chemically reacted with the silica fibers of 

the base material and prevented water absorption into the tile.51

Since the tiles could not withstand airframe-load deformation, stress isolation was necessary between the tiles and the orbiter structure. This was provided 

by SIPs made of Nomex felt in thicknesses of 0.090, 0.115, or 0.160 inch. 

 51.

213
Coming Home

The SIP was bonded to the tiles, and then bonded to the orbiter skin using 

a silicon RTV adhesive. However, the SIP introduced stress concentrations 

at the needled fiber bundles. This resulted in localized failure in the tile just 

above the RTV bond line. To solve this problem, the inner surface of the tile 

was densified to distribute the load more uniformly. The densification process 

used a Ludox ammonia-stabilized binder. When mixed with silica slip particles 

it became a cement, and when mixed with water it dried to a finished hard 

surface. A silica-tetraboride coloring agent was mixed with the compound for 

penetration identification. The densification coating penetrated the tile to a 

depth of 0.125 inch, and the strength and stiffness of the tile and SIP system 

were increased by a factor of two.

The tiles thermally expanded or contracted very little compared to the orbiter 

structure, so it was necessary to leave 0.025- to 0.065-inch gaps between them 

to prevent tile-to-tile contact. Nomex felt-material insulation was required in 

the bottom of the gap between tiles. These filler bars were supplied in thicknesses corresponding to the SIPs, cut into strips 0.75-inch wide, and bonded 

to the orbiter structure. The filler bar was waterproof and temperature-resistant 

up to approximately 800 °F.

White blankets made of coated Nomex felt reusable surface insulation 

(FRSI) were used on the upper payload bay doors, portions of the midfuselage and aft-fuselage sides, portions of the upper wing surface, and parts of the 

OMS/RCS pods. The FRSI blankets protected areas where temperatures were 

below 700 °F. The FRSI contained the same Nomex material as the SIP used 

under the tiles. The FRSI blankets generally consisted of 3-by-4-foot, 0.16- 

to-0.40-inch-thick sheets, depending on location. A white-pigmented 92-007 

silicon elastomer coating functioned as aerodynamic-erosion protection and 

as a moisture barrier, and it acted as a passive cooling system by having low 

solar absorptance and high reflectance of solar radiation. The FRSI had an 

emittance of 0.8 and solar absorptance of 0.32 and covered nearly half of the 

orbiter’s upper surfaces. The density of the coated material was 5.4 pounds 

per cubic foot.52

Nomex felt is a noncombustible, heat-resistant aromatic polyamide (aramid) 

fiber that is 2 deniers in fineness, 3 inches long, and crimped. These were loaded 

into a carding machine that untangled the clumps of fibers and combed them 

to make a tenuous mass of lengthwise-oriented, relatively parallel fibers called 

a web. The cross-lapped web was fed into a loom, where it was lightly needled 

into a batt. Generally, two such batts were placed face-to-face and needled 

together to form felt. The felt then was subjected to a multi-needle pass process 

 52.

214
Spaceplane Reality

that passed barbed needles through the fiber web in a sewinglike procedure 

that compacted the transversely oriented fibers into a pad. This needle-punch 

operation was repeated as many times as required to produce a felt product 

with specified physical properties. The needled felt was calendered to the proper 

thickness by passing through heated rollers at selected pressures. The calendered material was heat-set at 500 °F to thermally stabilize the felt. The FRSI 

was bonded directly to the orbiter by RTV silicon adhesive, 0.20 inch thick.53

After the initial delivery of Columbia, an advanced flexible reusable surface insulation (AFRSI) was developed as a lightweight replacement for the 

thin, fragile LRSI tiles. The AFRSI blankets protected areas where temperatures were below 1,200 °F and aerodynamic loads were minimal. AFRSI was 

a quilted blanket consisting of a quartz fibrous batt insulation and a woven 

quartz fiber outer fabric quilted on 1-inch centers to a glass fiber inner fabric. 

The quilting was accomplished using a Teflon-coated quartz thread with three 

to four stitches per inch. The sewn quilted fabric blanket was manufactured 

in 3-by-3-foot squares with an overall thickness of 0.25 to 2 inches. The material had a density of 8 to 9 pounds per cubic foot and was used where surface 

temperatures ranged from 750 °F to 1,200 °F. The blankets were cut to the 

planform shape required and bonded directly to the orbiter by RTV silicon 

adhesive 0.20 inch thick.54

The first application of AFRSI was on the OMS pods of Columbia during 

STS-6. AFRSI blankets were used on Discovery and Atlantis to replace the 

majority of the LRSI tiles. The LRSI tiles on Columbia’s midfuselage, payload 

bay doors, and vertical stabilizer were replaced with AFRSI blankets during 

the Challenger standdown. Endeavour was delivered with an even greater use 

of AFRSI, and the other orbiters (except OV-102) initially migrated toward 

this configuration during normal maintenance as well as major overhauls. 

These blankets were initially used on the fuselage, upper elevens, upper wings, 

vertical stabilizer, and forward canopy surfaces. However, in an effort to save 

as much weight as possible for missions to the International Space Station 

(ISS), much of the AFRSI on the mid fuselage and aft fuselage, payload bay 

doors, and upper wing surfaces of the three ISS orbiters was subsequently 

replaced by the lighter FRSI during Orbiter Maintenance Down Period 

(OMDP). During Endeavour’s first OMDP, 1,472 pounds were saved by 

this modification. A change to the flipper doors that included FRSI saved 

another 520 pounds.55

 53.

215
Coming Home

As of its last flight, in 2011, Atlantis56 was protected with 18,490 HRSI-9 

(LI-900) tiles, 2,846 FRCI-12 tiles, 469 AETB-8 (TUFI) tiles, 234 Boeing 

Replacement Insulation (BRI)-18 tiles, and 3,254 thermal blankets of various 

types. Approximately 20 tiles were typically damaged on each flight, and the 

average postflight refurbishment consisted of about 200 tiles and 10 blankets. 

However, approximately half of those were replaced for reasons unrelated to their 

failure. These reasons included removing a tile (or blanket) to gain access to the 

structure or component it covers for maintenance, or replacing the tile (or blanket) with an improved version to solve some perceived (or anticipated) anomaly.

Thermal barriers prevented hot air and plasma from flowing through the 

various penetrations in the orbiter, such as the landing-gear doors and crew 

hatch. The thermal barriers were constructed of tubular Inconel 750 wire mesh 

filled with alumina silica fibrous insulation, or Saffil, wrapped into a ceramic 

sleeving and a Nextel AB312 ceramic alumina-borosilicate fabric outer cover. 

A black ceramic emittance coating was applied to the outer surface and an 

RTV silicone rubber was used to attach the barrier to the structure. Thermal 

barriers were designed for service in areas seeing temperatures up to 2,000 °F.57

Because the tiles were installed with a space between them, it was sometimes 

necessary to use gap fillers to ensure that heat did not seep between the tiles 

and impact the aluminum skin. Two types of gap fillers were used: the pillow/

pad and the Ames type. The pillow/pad gap filler was made of a ceramic fabric 

with alumina silica insulation with two layers of 0.001-inch-thick Inconel foil 

inside for support. The pad gap filler was stitched with quartz thread, and a 

black ceramic emittance coating was applied to the outer mold line. A coating 

of red RTV-560 was applied to the bottom surface as a stiffener to aid with 

installation into the tile-to-tile gaps. Approximately 5,000 pad-type gap fillers, 

each 0.200 inch thick, were used on the lower body flap, lower elevons, and 

vertical stabilizer. The Ames-type gap filler was made from a ceramic fabric 

that was originally impregnated with a black silicone coating, although this 

was subsequently replaced with a more durable ceramic coating.58

Two types of internal insulation blankets were used: fibrous bulk and multilayer. The bulk blankets were fibrous materials with a density of 2 pounds per cubic 

foot and a sewn cover of reinforced-Kapton acrylic film. The cover material had 

13,500 holes per square foot for venting. Acrylic-film tape was used for cutouts, 

 56.

216
Spaceplane Reality

patching, and reinforcements. Tufts throughout the blankets minimized billowing during venting. The multilayer blankets were constructed of alternate layers of 

perforated-Kapton acrylic-film reflectors and Dacron-net separators. There were 16 

reflector layers in all, with the two-cover halves counting as two layers. The covers, 

tufting, and acrylic-film tape were similar to that used for the bulk blankets.59

Flags and letters were painted on the orbiter with a Dow Corning-3140 

silicon-based material, colored with pigments. It was basically the same paint 

used to paint automobile engines, and it would break down in temperature 

ranges between 800 °F and 1,000 °F. Because of this, almost all markings were 

painted in relatively low-temperature areas of the orbiter.

Space Shuttle Entry

After completing its on-orbit tasks, the Space Shuttle crew oriented the orbiter 

to a tail-first attitude using the RCS, then used the two OMS engines to slow 

the orbiter for deorbit. The RCS turned the orbiter’s nose forward for an entry 

that began at 400,000 feet (defined as the entry interface), slightly over 5,000 

miles from the landing site. The orbiter’s velocity at entry was approximately 

17,500 mph with a 40-degree angle of attack. The forward RCS thrusters were 

inhibited by the onboard computers immediately before entry, and the aft RCS 

thrusters maneuvered the vehicle until a dynamic pressure of 10 pounds per 

square feet was sensed, which was when the orbiter’s ailerons became effective. 

The aft RCS roll thrusters were then deactivated. At a dynamic pressure of 20 

pounds per square foot, the orbiter’s elevators became effective and the aft 

RCS pitch thrusters were deactivated. The orbiter’s speed brake was used below 

Mach 10.0 to induce a more positive downward elevator-trim deflection. At 

approximately Mach 3.5, the aerodynamic rudder was activated, and the aft 

RCS yaw thrusters were deactivated at 45,000 feet.60

Entry guidance had to dissipate the tremendous amount of energy the 

orbiter possessed when it entered the atmosphere to ensure that the vehicle did 

not either burn up (entry angle too steep) or skip back out of the atmosphere 

(entry angle too shallow). During entry, excess energy was dissipated by the 

atmospheric drag on the orbiter. A steep trajectory gave higher atmosphericdrag levels, which resulted in faster energy dissipation. Normally, the angle of 

attack and angle of roll enabled the atmospheric drag of any flight vehicle to 

 59.

217
Coming Home

be controlled. However, for the orbiter, angle-of-attack variation was rejected 

because it created exterior surface temperatures in excess of the capabilities 

of the thermal protection system. The angle-of-attack schedule was loaded 

into the computers as a function of relative velocity, leaving only roll angle 

for energy control. Increasing the roll angle decreased the vertical component 

of lift, causing a higher sink rate and increased energy dissipation. Increasing 

the roll angle raised the surface temperature of the orbiter somewhat, but not 

nearly as drastically as an equivalent angle-of-attack variation.

If the orbiter was low on energy (current range-to-go greater than nominal), 

entry guidance would command lower-than-nominal drag levels. If the orbiter 

had too much energy (current range-to-go less than nominal), entry guidance 

would command higher-than-nominal levels to dissipate the extra energy to 

within the orbiter’s limits to withstand the additional surface heating. The goal 

was to maintain a constant heating rate until the orbiter was below 13,000 mph. 

During entry, the temperatures on the wing leading edge increased steadily for 

about 10 minutes, at which point the orbiter was traveling at Mach 23 and at 

230,000 feet with wing leading-edge temperatures of about 3,000 °F. While the 

orbiter was descending, superheated air molecules discharged light that could 

be seen as bright flashes through the crew cabin windows.61

Roll angle was also used to control cross range. Azimuth error is (1) the 

angle between the plane containing the orbiter’s position vector and headingalignment cylinder tangency point and (2) the plane containing the orbiter’s 

position vector and velocity vector. When the azimuth error exceeded a predetermined number, the orbiter’s roll angle was reversed.

The equilibrium-glide phase shifted the orbiter from the rapidly increasing 

drag levels of the temperature-control phase to the constant-drag phase. The 

equilibrium-glide flight was defined as flight in which the flightpath angle—the 

angle between the local horizontal and the local velocity vector—remained constant. Equilibrium-glide flight provided the maximum downrange capability 

and lasted until the drag acceleration reached 33 ft/sec squared. At that point, 

the constant-drag phase began. The angle of attack was initially 40 degrees 

but ramped down toward 36 degrees at the end of the phase. In the transition 

phase, the angle of attack continued to ramp down, reaching approximately 14 

degrees as the orbiter reached the terminal area energy management (TAEM) 

interface at approximately 83,000 feet and Mach 2.5.

The TAEM interface steered the orbiter to the nearest of two headingalignment circles (HAC), whose radii were approximately 18,000 feet and 

which were located tangent to and on either side of the runway centerline 

 61.

218
Spaceplane Reality

STS-4 sustained some damage to the thermal protection system, but the orbiter returned to 

Earth safely. NASA S82-28874.

on the approach end. In TAEM guidance, excess energy was dissipated with 

S-turns, and the speed brake was used to modify the drag coefficient as 

required. This decreased the ground-track range as the orbiter turned away 

from the nearest HAC until sufficient energy was dissipated to allow a normal 

approach and landing. The spacecraft slowed to subsonic velocity at approximately 49,000 feet, about 25 miles from the runway.

The approach-and-landing phase began at approximately 10,000 feet at an 

equivalent airspeed of 320 mph, roughly 8 miles from touchdown. Autoland 

guidance was initiated at that point to guide the orbiter to the –19-degree glideslope (about seven times as steep as a commercial airliner), aimed at a target 1 mile 

in front of the runway. The speed brake was used to control velocity, and at 1,750 

feet above ground level a preflare maneuver was started to position the spacecraft 

for a 1.5-degree glideslope in preparation for landing. The final phase reduced 

the orbiter’s sink rate to less than 9 ft/sec. Touchdown occurred approximately 

2,500 feet past the runway threshold at a speed of roughly 220 mph.

219
The National Aero-Space Plane served as a technology driver for the aerospace community 

throughout the mid-1980s, but the program never got to the flight stage. NASA.

220
CHAPTER 6

This chapter explores efforts to develop new reentry and landing concepts from 

the 1990s on. During this period, a series of ideas emerged that ranged from the 

DC-X powered landing concept to the return of a metallic heat shield for the 

National Aero-Space Plane (NASP) and X-33. The chapter includes a discussion 

of the decision to pursue development of the Crew Exploration Vehicle (CEV) 

using a capsule, a blunt-body, an ablative heat shield, and parachutes (or perhaps 

a Rogallo wing) to return to Earth (or perhaps land in the ocean) in 2005, as well 

as the decision to move toward a commercial piloted spacecraft that could be 

reusable and could possibly serve as a capsule. The recovery of scientific sample 

return missions to Earth (both the loss of Genesis and the successful return of 

Stardust) suggests that these issues are not exclusive to the human space flight 

community. While the effectiveness of the thermal protection system on the Space 

Shuttle has been demonstrated, its fragility and high maintenance costs remained 

an important concern for follow-on space reentry and recovery efforts. As a result, 

a large slate of new research programs on thermal protection systems for various 

vehicles—such as NASP, X-33, and X-34—followed in the 1980s and 1990s. 

Some of those efforts were exotic, but they eventually led full circle to a new version 

of ablative heat shields developed for the Genesis and Stardust missions, as well as 

the heat shield’s most advanced concept, expressed in the Constellation program.

The Rise (and Fall) of the X-30 

National Aero-Space Plane

Many aerospace engineers have long believed that the best solution to the world’s 

launch needs would be a series of completely reusable launch vehicles (RLVs). A 

debate has raged between those who believe RLVs are the only—or at least the 

best—answer and those who emphasize the continuing place of expendable launch 

vehicles (ELVs) in future space-access operations. This debate relates to estimates 

of the number of missions required: the more missions, the more effective a RLV 

system becomes. However, estimates of the number of future space flights have 

consistently been too high, thus favoring the ELV concepts. The same debate 

221
Coming Home

took place at the beginning of the Space Shuttle’s development. RLV advocates 

have been convincing in their argument that the only course leading to “efficient 

transportation to and from the earth” would be RLVs, and they have made the 

case repeatedly since the late 1960s.1

 Their model for a prosperous future in space 

is the airline industry, with its thousands of flights per year and its exceptionally 

safe and reliable operations. Since the advent of the Space Shuttle, NASA has been 

committed to advancing this model with Shuttle follow-on efforts, until a radical 

departure in 2005 with the short-lived Constellation program.

One especially important effort involved work on a hybrid air and spaceplane 

that would enable ordinary people to travel halfway around the globe in about 

1 hour. Such a concept was quite simple in theory but enormously complex in 

reality. It would require developing a passenger spaceplane with the capability to 

fly on wings from an ordinary runway like a conventional jet. Flying supersonic, 

it would reach an altitude of about 45,000 feet, at which point the pilot would 

start scramjet engines—a more efficient, faster jet engine with the potential to 

reach hypersonic speeds in the Mach 3 realm. These capabilities would take the 

vehicle to the edge of space for a flight to the opposite side of the globe, where 

the process would be reversed, and the vehicle would land like a conventional 

airplane. The vehicle would never reach orbit, but it would still fly in space, and 

the result would have been the same as orbital flight for passengers but for less 

time. It would even be possible, RLV supporters insisted, to build such a spaceplane that could reach orbit.2

One of the most significant efforts toward developing this reusable spaceplane was the NASP, a joint NASA/Air Force technology demonstrator begun 

in the early 1980s. With the beginning of the Ronald Reagan administration, 

and its associated military buildup, Tony DuPont, head of DuPont Aerospace, 

offered an unsolicited proposal to the Defense Advanced Research Projects 

Agency (DARPA) to design a hypersonic vehicle powered by a hybrid-integrated 

engine of scramjets and rockets. DARPA Program Manager Bob Williams liked 

the idea, and he funded it as a “black” program code-named “Copper Canyon” 

between 1983 and 1985. The Reagan administration later unveiled it as the 

 1. This was the argument made to obtain approval for the Space Shuttle. See 

222
Back to the “Ablative” Future

National Aero-Spaceplane, designated the X-30. Reagan called it “a new Orient 

Express that could, by the end of the next decade, take off from Dulles Airport 

and accelerate up to twenty-five times the speed of sound, attaining low-Earth 

orbit or flying to Tokyo within two hours.”3

 With this public announcement 

of NASP, the hypersonic aerospaceplane had returned.4

The NASP program initially proposed to build two research craft, at least 

one of which was intended to achieve orbit by flying in a single stage through 

the atmosphere at speeds up to Mach 25. The proposed X-30 would use a multicycle engine that shifted from jet to ramjet to scramjet speeds as the vehicle 

ascended, burning liquid hydrogen fuel with oxygen scooped and frozen from 

the atmosphere.5

 It never achieved anything approaching flight status.

NASP, like earlier hypersonic flight research projects, fell victim to budget 

cuts, but this time as a result of the end of the Cold War. With costs approaching $17 billion, some 500 percent over the projected budget, it was also a 

program 11 years behind schedule when canceled. With cost overruns and 

delays in scheduling, the project was less attractive than it had been originally. 

Accordingly, Congress canceled NASP in 1992, during fiscal year 1993 budget 

deliberations. In 1994, the program died a merciful death, trapped as it was in 

bureaucratic politics and seemingly endless technological difficulties.6

 3. Ronald Reagan, “State of the Union Address,” February 4, 1986.

223
Coming Home

Although the program never came close to building or flying hardware, NASP 

contributed significantly to the advancement of materials that were either capable 

of repeatedly withstanding high temperatures (on the vehicle’s nose and body) 

or capable of tolerating repeated exposure to extremely low temperatures (the 

cryogenic fuel tanks). By 1990, NASP researchers had realized significant progress in titanium aluminides, titanium aluminide metal matrix composites, and 

coated carbon-carbon composites that were all useful for thermal protection 

because of their resistance to superheating. Moreover, Government and contractor laboratories had fabricated and tested large titanium aluminide panels under 

approximate vehicle thermal conditions, and NASP contractors had fabricated 

and tested titanium aluminide composite pieces.7

 All of this research would have 

contributed to the realization of NASP’s capability to reenter the atmosphere, 

slow to landing speed, and recover from space flight.

Two areas of research were important in NASP’s thermal protection system. 

First, the vehicle’s engineers explored several categories of thermal protection 

systems, but the most significant area involved convectively (or actively) cooled 

structures—a method of limiting structural temperatures by circulating a coolant through the hot zones of the vehicle—for the fuselage. This technique 

had been extensively studied as far back as the 1950s, but NASP advanced 

it considerably, even though it never yielded a workable system.8 This active 

cooling approach applied fluids in a cooling loop to high heat-flux areas of the 

space vehicle. This, of course, required the vehicle to carry excess weight for 

the cooling system, and it was never feasible when fluids such as liquid metals 

(like sodium or potassium) were the coolant. This approach became more 

tenable when NASP pioneered development of a system that used the liquid 

hydrogen of the fuel system for this purpose. The weight of the fuel then served 

a dual purpose in the spaceplane.9 Ironically, the highest speeds possible in the 

combined-cycle engines under development for NASP were actually aided 

through the effective and imaginative use of active thermal management. As 

late as February 7, 1992, as the program was about to be canceled, a NASP 

structure was filled with liquid hydrogen (at –423 °F). The assembly then successfully endured bending and heating to 1,300 °F on the shell. This process 

 7. Schweikart, “The Quest for the Orbital Jet,” pp. III.37-III.38 and III.41-III.42.

224
Back to the “Ablative” Future

simulated atmospheric heating as high as Mach 16 and was an important success, but it was not sufficient for orbital flight.10

A major challenge, never truly resolved during the NASP program, involved 

curtailing the mass and weight of the active thermal-management system, 

which grew ever more complex and massive as speeds increased. Although 

much of the program was highly classified, this seems to remain a future 

research question.11

Second, in addition to convective cooling systems, the NASP program 

pursued research in advanced materials, including various composites and titanium-based alloys that retained structural integrity at temperatures sometimes 

in excess of 1,800 °F.12 NASP engineers pursued significant advances in metal 

matrix composites, consisting of several threads of research in advanced metal 

matrices and high-strength fiber composites. Advanced titanium and beryllium alloys had high strength and high temperature resistance while weight 

remained low. These lightweight materials possessed good thermal conductivity 

that would be critical in the successful completion of a single-stage-to-orbit 

(SSTO) vehicle.13

The X-33/VentureStarTM Program

NASA began its own RLV program after the demise of NASP in the mid1990s, and the Agency’s leadership expressed high hopes for the proposed 

X-33, a small suborbital vehicle that would demonstrate the technologies 

required for an operational SSTO launcher. This technology demonstrator 

was intended as the first of a projected set of four stages that would lead to a 

routine spacefaring capability. The X-33 project, undertaken in partnership 

 10.

225
Coming Home

An artist’s concept for the X-33. The X-33 program was created in 1997 to develop technologies that were intended to pave the way for a full-scale, commercially developed reusable 

launch vehicle. NASA 9906365.

with Lockheed Martin, had an ambitious timetable to fly by 2001. But what 

would happen after its tests were completed remains unclear. Even assuming 

complete success in meeting its R&D objectives, the time and money necessary to build, test, and certify a full-scale operational follow-on version remains 

problematic. Who would pay for such an operational vehicle also remains a 

226
Back to the “Ablative” Future

mystery, especially since the private sector had become less enamored with the 

joint project over the years and eased itself away from the venture.14

There is also an understanding that the technical hurdles have proven 

more daunting than anticipated, as was the case 30 years ago with the Space 

Shuttle and more recently with NASP. Any SSTO vehicle (and X-33 held 

true to this pattern) would require breakthroughs in a number of technologies, particularly in propulsion, materials, and thermal protection. When 

designers are to begin work on the full-scale SSTO, they might find that 

available technologies limit payload size so severely that the new vehicle 

provides little or no cost savings compared to old launchers. If this were to 

become the case, then everyone must understand that NASA would receive 

the same type of criticism that it had experienced over the Space Shuttle. 

Without question, this related to the inability to accurately predict the flight 

rate for space-access missions.15

This is not to say that an SSTO vehicle could never work or that the X-33 

should not have been pursued. It has always been NASA’s job to take risks and 

push the technological envelope. However, while the goal may be the development of a launch system that is significantly cheaper, more reliable, and more 

flexible than presently available, it is possible to envision a future system that 

cannot meet those objectives. This is all the more true in a situation in which 

breakthrough technologies had not emerged.16

Even so, NASA joined Lockheed Martin in an unusual partnership to 

develop the X-33, and despite the outcome of that particular effort it might 

prove a model worth returning to in the 21st century. NASA undertook this 

effort because over the past two decades the Space Shuttle, a partially reusable 

vehicle, taught political leaders that bringing down costs and reducing turnaround time were much greater challenges than originally anticipated. Even 

at the time, it was a conscious decision on the part of NASA and the White 

House to select a partially reusable Space Shuttle to minimize the development 

costs, fully recognizing the decision would have an adverse effect on long-term 

 14.

227
Coming Home

VentureStar® was to have been a follow-on to the smaller, experimental X-33 technology demonstrator, but cost, schedule, and technological problems led to its cancellation in April 2001. 

NASA MSFC-9906386.

operational costs.17 The X-33 program was intended to meet those objectives of 

cost effectiveness and reliability that the Shuttle had left unfulfilled. It was to 

have served as a technology demonstrator, but upon successful flight Lockheed 

Martin agreed to develop and fly a commercial version, VentureStarTM, in the 

first decade of the 21st century.

That cooperative agreement, furthermore, was at the heart of NASA’s efforts 

to initiate so-called new ways of doing business by serving as both procurement and management tools. As a new procurement instrument, the cooperative agreement was part of a fast-track managerial approach (and part of the 

larger “faster-better-cheaper” formula) that featured more expeditious acquisition procedures, streamlined bureaucracy, limited oversight, and lessened 

 17.

228
Back to the “Ablative” Future

investment by NASA. It also mandated that industry share the costs of the 

project so that both sides shared the risk.18

Given the problems experienced on the X-33 program (with delays of more 

than 1 year because of difficulties with critical elements such as the fuel tanks), 

some criticized the use of the cooperative agreement as the culprit in sidelining the project. That seems to put too fine a point on the issue, as perhaps the 

greatest advantage of that cooperative agreement was its fostering of shared 

responsibility for the funding of launcher development between the public and 

private sectors. The cooperative agreement also provided industry an important 

voice in launcher planning.

Before the X-33, the space industry had never expended significant resources 

in launcher development. The industry contribution to X-33 development was 

$442 million through 2000. In an era of declining Government space R&D 

budgets, the importance of that number cannot be underestimated. It seems 

obvious that although sizable Government investment in the development of 

future launchers will be required, there is a reason to pursue additional cooperative projects.19 If a new generation of launchers is to be developed in the first 

part of the 21st century, it appears that the Federal Government will have to 

take a leadership role in identifying R&D funds and working in partnership 

with industry. Perhaps this would be most effectively accomplished through 

additional cooperative agreements.

Of course, one may also legitimately criticize the overall X-33 effort as 

a self-deception that a single Government program—even one using a new 

type of partnership with industry—will be able to solve the problems of space 

access. In this case, like NASP of the 1980s, an ambitious program was created, hyped as the panacea for all space-access challenges, underfunded, and 

then ridiculed when it failed. Unrealistic goals, coupled with impossible political demands, left the X-33 program stunned and stunted at century’s end. 

As one space policy analyst concluded, “it continued to blur the line, which 

 18.

229
Coming Home

should be bright, between revolutionary, high-risk, high-payoff R&D efforts 

and low-risk, marginal payoff evolutionary efforts to improve systems.”20 The 

willingness to embark on this effort has been a mystery to many analysts as 

they review the linkage— which some called “foolish”—between the high-risk 

R&D of the X-33 and the supposedly operational VentureStar.

The X-33 program used direct technology developments from the NASP 

program for its thermal protection system and, in the process, sought to 

avoid a large-scale R&D program. Using metal and metal matrix composites 

in thermal protection materials, the X-33 sought to exploit those materials’ robustness and durability while sacrificing tolerance for extremely high 

temperatures. Project engineers refrained from the ceramic tiles of the Space 

Shuttle in favor of a metallic thermal protection system because of its intrinsic ductility and inherent resistance to water. Metallic thermal protection 

systems had a long history to this point. The notable metallic heat-sink 

materials include the copper nose tips on early Jupiter intermediate-range 

ballistic missiles, the Beryllium nose tips used on the Polaris submarinelaunched ballistic missiles, and the Inconel X used on the X-15.21 Even so, 

the X-33 engineering team insisted that the new thermal protection system 

would be pathbreaking: “The success of the X-33 will overcome the ballistic 

reentry TPS mindset. The X-33 TPS is tailored to an aircraft type mission 

while maintaining sufficient operational margins. The flight test program 

for the X-33 will demonstrate that TPS for the RLV is not simply a surface 

insulation but rather an integrated aeroshell system.”22

Approaching the X-33 thermal protection system presented engineers with 

significant challenges. Of course, the lifting body configuration of the X-33 

also offered some opportunities. The two principal objectives for the thermal 

protection system were to shield the primary airframe structure from excessive 

thermal loads and to provide an aerodynamic surface for the vehicle. Because 

the hypersonic environment for a lifting body is well understood, the X-33 

team chose an integrated standoff aeroshell design with minimal weight to 

reduce procurement and operational costs. As one research report noted:

 20.

230
Back to the “Ablative” Future

Today, metallic TPS found on the X-33 is a hybrid of materials 

sealed in a panel. For this type of thermal protection system, very 

thin foils act with a metallic honeycomb and fibrous insulation 

to form a TPS panel. The approach can provide the durable acreage coverage needed for reusable launch vehicles. The degree of 

maintenance on these metallic systems should be a fraction of 

the maintenance that the tiles require on the space shuttle. These 

metallic systems may require emissive coatings to help cool the 

vehicle on orbit. If the launch vehicle is orbital, acreage metallic 

TPS will need a coating with high emissivity in order to cool during orbit. The defunct National AeroSpace Program [sic] (NASP) 

developed such a coating tailored to titanium aluminide.23

The TPS developed for the X-33 consisted of diamond-shaped metallic panels 

approximately 46 centimeters along each side. Each of these panels incorporated 

a metallic honeycomb sandwich heat-shield outer panel with foil-encapsulated 

fibrous insulation attached to the inner side of the heat-shield panel. To put these 

shingles together on the vehicle, NASA used a metallic standoff rosette at each 

corner attached, in turn, to a composite support structure.

By 1998, much of the basic design work had been completed on the X-33 

thermal protection system, and tests were underway. As one report noted:

There are two basic panel types, with two material variations on 

the windward body. The primary panel type is made of Inconel 

617 with .006” inner and outer facesheets brazed to .0015” thick, 

3/16” cell core with a total thickness at .50”. Areas on the vehicle 

that are subjected to higher temperatures require .010” thick 

facesheets and .0035”, 3/16” cell core. The other primary panel 

type is a .50” achined isogrid made of Inconel 617, or MA 754 

where exposed to higher temperatures. The isogrid panels are used 

primarily at locations of high curvature where producibility or 

high stresses dictate their use. Isogrid panels are also used when a 

penetration, such as an RCS thruster or antenna[,] is located. The 

use of isogrid panels is minimized, as they are relatively heavier 

than honeycomb brazed panels....

 23.

231
Coming Home

The panel-to-panel seals consist of a primary shingle seal, which 

is an extension of the outer skin, and a secondary seal system. The 

secondary seal system has a ‘J’ hook and a leaf seal that contacts 

the ‘J’ seal. The secondary seal system has been designed to provide 

“fly home” capability if the primary seal fails.24

Of course, the thermal protection system was designed to perform the dual 

role of supporting the aerodynamic-pressure loads and providing thermal protection. The critical question was that, although metallic thermal protection 

system concepts had been demonstrated previously, could one be developed 

to protect the spacecraft at orbital reentry speeds and heating.25

The metallic thermal protection system of the X-33 originated through 

years of research sponsored by NASA and the Department of Defense as well as 

through its manufacture by Rohr Industries. Progressing from early standoff heat 

shields to multi-wall concepts, thermal protection systems by the latter stages of 

NASP used state-of-the-art prepackaged superalloy honeycomb sandwich panels. 

The X-33 advanced this effort and undertook detailed design and fabrication of 

these superalloy honeycomb thermal protection system elements, finding that 

they worked relatively effectively during evaluation.26 Modifications resulting 

 24.

232
Back to the “Ablative” Future

from tests of the system led to the outer honeycomb surface supported at four 

points by an underlying latticework that formed the aeroshell, supporting the 

vehicle’s aerodynamic shape.27 As one study of this effort noted:

Fibrous insulation, beneath the outer honeycomb sandwich panel, 

is encapsulated by foil attached to the perimeter of the panel. Seals 

between panels on the outer surface are intended to maintain the 

aerodynamic pressure across the outer, hot honeycomb sandwich. 

In contrast, the prepackaged superalloy honeycomb TPS is vented 

so that there is little pressure difference across the outer honeycomb sandwich when it is hot.28

Despite relative success with the X-33 thermal protection system, the 

X-33 was mired in seemingly inscrutable technological problems and bureaucratic challenges, and NASA lost faith in it and terminated the effort in 2001. 

Thereafter, NASA officials expressed a deeper understanding that the technical 

hurdles proved more daunting than anticipated, as was the case 30 years ago 

with the Space Shuttle, and more recently with the NASP.

Any successful SSTO vehicle (and the X-33 program reinforced this pattern) would require breakthroughs in a number of technologies, particularly 

in propulsion and materials.29 Some engineers referred to X-33 as being built 

from “unobtainium,” and they thought the United States should instead pursue 

more conventional space-access technologies. Without going that far, many 

engineers recognize four major challenges that complicate efforts to develop 

this technology:

1. Aerodynamics

2. Guidance and Control

3. Materials

4. Propulsion

In the first realm of aerodynamics, through a succession of projects and 

studies, researchers have overcome many of the roadblocks to create effective 

shapes for a hypersonic vehicle. Many of the aerodynamic questions are now 

 27.

233
Coming Home

satisfactorily understood. The same is true for the second challenge of guidance and control. Materials research remains an important aspect yet to be 

resolved as research continues on heat-resistant materials and composites that 

can reduce weight. But the biggest issue remains propulsion. There is, as yet, no 

fully functional scramjet engine (at least not in the public arena). Furthermore, 

scramjet engines require high dynamic pressure to operate effectively and are 

not ideally suited to the SSTO mission, which requires acceleration at low and 

zero dynamic pressure. These are all problems for future researchers to solve 

before realizing the dream of a true SSTO vehicle.30

X-34 and the Continuing Challenge of Reentry

Almost in parallel with the X-33 program, NASA undertook the X-34—also 

known as the Reusable Small Booster Program—to demonstrate certain technologies and operations useful to smaller reusable vehicles launched from an 

aircraft. Among those technologies were autonomous ascent, reentry, and landing; composite structures; reusable liquid-oxygen tanks; rapid vehicle turnaround; and thermal protection materials. The X-34 was considerably smaller 

and lighter than the X-33 and was supposed to be capable of hypersonic flight 

to Mach 8—but not nearly as capable as the X-33’s Mach 15. Consequently, 

the X-34 was to have been much less expensive and simpler to develop, operate, and modify.31

Design of the thermal protection system for this hypersonic flight vehicle 

required (just as it had for the X-33) determining the peak temperatures over 

the surface and the heating-rate history along the flight profile. Based on these 

determinations, project engineers employed a system that would work for the 

suborbital technology demonstrator containing advanced thermal protection 

systems capable of surviving subsonic flights through rain and fog. As one 

report stated:

A passive system, such as that employed on the X-34, takes advantage of the insulative properties of the TPS materials to hold 

a significant portion of the incident heat until it is radiated or 

convected away from the vehicle. The two critical parameters that 

 30.

234
Back to the “Ablative” Future

Artist concept of X-34 technology demonstrator, another second-generation RLV, in flight. The 

X-34 was a reusable-technology test that was to be capable of speeds up to Mach 8 and 

altitudes of 250,000 feet. NASA 9906366.

determine the TPS options suitable for a given vehicle and trajectory are the peak-heating rate and the integrated heating over the 

time of the flight profile. The former determines the maximum 

temperature environment, and thus the material options, and 

the latter determines the thickness distribution of the insulative 

TPS material required to protect the surface. Except for the nosecap and wing leading edges [where silicone impregnated reusable 

ceramic ablator (SIRCA) tile is used], the X-34 is protected with 

flexible-blanket or felt systems.32

There was very little subsequent application of this research effort because 

the project was canceled in 2001, and it did not progress to the flight stage.33

 32.

235
Coming Home

X-38 Recovery Parafoil

In a departure from the X-33/X-34 thermal protection system story, NASA 

also pursued parafoil development in the late 1990s as a dedicated method of 

returning the crew aboard the ISS to Earth. Although through 2009 the ISS 

could handle only a crew of three, when it was completed an international crew 

of up to seven was intended to be able to live and work in space for between 3 

and 6 months. Space agencies anticipated that a major problem in sending up a 

large crew, however, would be returning them safely to Earth in the event of an 

emergency. This led to a contentious debate over the value of the Space Station, 

as funding and program schedule again came under fire from critics. To ensure 

the crew size of seven, in 1994 NASA added to the program a U.S.-developed 

Crew Return Vehicle (CRV). Even so, it requested no additional funding to 

pay for the CRV. NASA carried the CRV as an overage in the budget for ISS 

until funding was finally allocated in the fiscal year 1999 budget submission 

to Congress. But that was only a small part of the problem.

As always, NASA engineers began the X-38 CRV project with optimism. 

The Agency had a record of excellence that had been forged during Project 

Apollo in the 1960s, and success after success had followed, including landings on Mars and voyages to the outer planets of the solar system. A culture of 

competence permeated NASA. Actor Robert Guillaume voiced the beliefs of 

many when in a sitcom he said, “You put an X anyplace in the solar system, 

and the engineers at NASA can land a spacecraft on it.”34 Nothing more effectively stated the public conception of NASA’s culture of competence than this 

public announcement. Unfortunately, that optimism and skill was misplaced 

in the case of the X-38, as costs and schedule combined to defeat the program.

The X-38 program really began at Dryden Flight Research Center in 1992, 

when Dale Reed, a veteran of the lifting body research of the 1960s and 1970s, 

began work for the JSC to test a CRV concept modeled on those earlier efforts. 

This proved successful, and in-house development of the X-38 concept began 

at JSC in early 1995. In the summer of 1995, early flight tests were conducted 

of the parafoil concept, dropping platforms with it from an aircraft at the 

Army’s Yuma Proving Ground, Yuma, AZ. In early 1996, NASA awarded a 

contract to Scaled Composites, Inc., of Mojave, CA, for the construction of 

three full-scale atmospheric test airframes. In September 1996, the first vehicle 

airframe was delivered to JSC, where it was outfitted with avionics, computer 

systems, and other hardware in preparation for flight tests at Dryden. The 

second vehicle was delivered to the Johnson Space Center in December 1996. 

 34.

236
Back to the “Ablative” Future

Artist’s concept of the X-38 Crew Return Vehicle. Intended to take the place of the Russian 

Soyuz capsule, the X-38 Crew Return Vehicle was pursued in the late 1990s as an escape 

vehicle for astronauts and cosmonauts aboard the International Space Station. NASA 9906387.

Testing progressed to an unpiloted space flight test in late 2000. About one 

hundred people quickly went to work on the project at Dryden and JSC.35

The wingless X-38 CRV, when operational, had been intended as the first 

reusable human spacecraft built in more than two decades. It was designed to 

fit the unique needs of a Space Station “lifeboat”—long term, maintenancefree reliability always in “turnkey” condition—ready to provide the crew a 

quick, safe trip home under any circumstances.36 The starting configuration 

for the X-38 was the SV-5, an Air Force effort developed under the START 

 35.

237
Coming Home

program. The SV-5 is the only lifting body shape that has been successfully 

demonstrated from orbital conditions (X-23 PRIME) to transonic and horizontal landing (X-24A). The NASA Flight Research Center supported the 

X-24A flight testing, which was treated much as an extension of the NASA 

lifting body program. During its development effort, the X-38 evolved the 

basic SV-5 shape.

Data from the aerodynamic studies contributed to the design and operational profile of the Space Shuttle and reemerged in the CRV program. When 

operational, the CRV was to serve as an emergency vehicle to return up to seven 

ISS crewmembers to Earth. It would be carried to the Space Station in the cargo 

bay of a Space Shuttle and attached to a docking port. If an emergency arose 

that forced the ISS crew to leave the Space Station, the CRV could be boarded, 

undocked, and after a deorbit engine burn, return to Earth much like a Space 

Shuttle. Not a true spacecraft in the traditional sense, the vehicle’s life support system could sustain a crew for about 7 hours. It used enhanced ceramic 

matrix components based on carbon-fiber-reinforced carbon, applied as nose 

and panel thermal protection systems and as hot-structure control surfaces 

(for the vehicle body), including the roller bearing assembly. Once the vehicle 

reached the denser atmosphere at about 40,000 feet, a steerable parafoil would 

be deployed to carry it through the final descent and landing. It would also be 

fully autonomous, in case the crew was incapacitated or otherwise unable to 

fly it, so it could return to the landing site using onboard navigation and flight 

control systems. Backup systems would, of course, allow the crew to pick a 

landing site and steer the parafoil to a landing.37

The X-38 design closely resembled the X-24A lifting body flown at Dryden 

between 1969 and 1971. That vehicle’s body shape, like the CRV, generated 

aerodynamic lift, which was essential to flight in the atmosphere. The 28 research 

missions flown by the X-24A helped demonstrate that hypersonic vehicles returning from orbital flight could be landed on conventional runways without power. 

 37.

238
Back to the “Ablative” Future

An X-38 test vehicle glides down under a giant parafoil toward a landing on Rogers Dry Lake 

near NASA’s Dryden Flight Research Center during its first free flight on November 2, 2000. The 

huge 7,500-square-foot parafoil enabled the vehicle to land in an area the length of a football 

field. NASA EC00-0317-41.

239
Coming Home

Interestingly, the horizontal landing capability demonstrated by the X-24A program was not incorporated into the X-38 program. Instead, it used a steerable 

parafoil concept for the actual landing. The three prototype X-38s used in the 

atmospheric flight-testing program were 24.5 feet long, 11.6 feet wide, and 

8.4 feet high—approximately 80 percent of the planned size of the CRV. The 

prototypes were designated V131, V132, and V131R. The V131 prototype was 

modified for additional testing beginning in the summer of 2000 and thereafter 

carried the designation V131R. A fourth prototype that was never completed, 

V133, was intended to incorporate the exact shape and size of the planned CRV. 

The prototype vehicles were manufactured by Burt Rutan and Scaled Composites 

in Mojave, CA. These vehicles had shells made of composite materials, such as 

fiberglass and graphite epoxy, and were strengthened with steel and aluminum 

at stress points; they weighed between 15,000 and 25,000 pounds.38

Another model that was never completed was a fully space-rated X-38 

CRV prototype, numbered V201, designed to test the concept under operational conditions. Its inner compartment, representing the crew area, would 

be a pressurized aluminum chamber. A composite fuselage structure would 

enclose the chamber, and the exterior surfaces would be covered with a 

thermal protection system designed to withstand the heat generated by air 

friction on atmospheric reentry. The thermal protection system would be 

similar to materials used on the Space Shuttle, but much more durable; it 

would use carbon and metallic-silica tiles for the hottest regions and flexible 

blanketlike material for areas receiving less heat during atmospheric reentry.39

Much of the technology for the X-38 was to be off the shelf to avoid lengthy 

R&D efforts. For instance, the flight control computer and the flight software 

operating system were commercially developed and used in many aerospace 

applications. Inertial navigation and global positioning systems, similar to units 

used on aircraft throughout the world, would be linked to the vehicle’s flight 

control system to steer the vehicles along the correct reentry path. Using global 

positioning already programmed into the navigation system, the flight control 

computer would become the autopilot that flies the vehicle to a predetermined 

landing site. Other avionics, flight control systems, materials, and aerodynamics used well-proven concepts and equipment. Finally, the U.S. Army originally 

 38.

240
Back to the “Ablative” Future

An artist’s depiction of NASA’s proposed X-38 Crew Return Vehicle reentering Earth’s atmosphere. 

The essential X-38 shape was derived from the earlier X-24A lift body tested at the Flight Research 

Center during the 1960s. The X-38 test vehicles were at 80 percent of full-size (28.5 feet long and 

14.5 feet wide) and weighed approximately 16,000 pounds, on average. In March 2000, Vehicle 132 

completed its third and final free flight in the highest, fastest, and longest X-38 flight to date. It was 

released at an altitude of 39,000 feet and flew freely for 45 seconds, reaching a speed of over 500 

mph before deploying its parachutes for a landing on Rogers Dry Lakebed. NASA ED97-43903-1.

developed the design of the parafoil that would deploy in the atmosphere 

and carry the X-38 to Earth. The European Space Agency (ESA) agreed to 

develop some of the X-38 systems, developing 15 subsystems or elements of 

the V201 X-38 spacecraft, scheduled for launch by Space Shuttle Columbia in 

September 2002. In addition, ESA provided the Guidance, Navigation, and 

Control (GNC) software for the parafoil phase of the V131R and V133 aerodynamic drop-test vehicles as well as for the supporting tests using a parafoil 

microlight aircraft and large drop pallets flying the full parafoil.40

 40.

241
Coming Home

The X-38’s thermal protection system development effort benefited significantly from ESA’s research and construction of the ceramic nose cap. As one 

study noted, “Over 10-years of experience in manufacturing carbon-based 

fiber-ceramic as well as the long-lasting competence in the range of system 

design of hot structures and thermal protection systems were converted into 

an innovative component concept. If the past work in this area had a rather 

experimental character (e.g. co-flight of ballistic capsule missions FOTON and 

CETEX on EXPRESS), the nose cap as a primary structure component had a 

mission-crucial task.”41 The study continued with the following:

During reentry into the earth’s atmosphere the nose cap of X-38 

experiences the highest thermal load of the entire thermal protection system due to its exposed location within the stagnation 

point of the vehicle. During the approx. 20 minutes reentry phase 

surface temperatures of up to 1,750°C were expected with a stagnation pressure of up to 10 to 150 hPa.

 The nose cap represents an absolute novelty regarding technological requirements. A component out of fiber-ceramic for thermal loads of up to approx. 1750°C had never been planned before 

with the claim of re-usability. The nose structure of the American 

Space shuttle is made of carbon/carbon (C/C) for example and 

the temperatures reach only scarcely 1500°C.

 The connection of the nose shell is made by 8 single fittings, 

which are made of fiber-ceramic as well respectively in the cooler 

range of a high temperature-steady (until approx. 1200°C) metal 

alloy (PM 1000 of Plansee, Austria). The special arrangement and 

design of the lever-like attachment system guarantees on the one 

hand a very good mechanical maximum load and on the other hand 

allows for an unrestricted thermal expansion of the structure, which 

can amount to 3mm with a medium diameter of the shell of 700 mm 

and the expected temperature level. In case of a rigid attachment the 

shell would be destroyed only by the thermal compressive forces. 

 For thermal isolation a multilevel flexible felt isolation made 

of oxide fibre-ceramic such as alumina, is intended between the 

nose shell and the sub-structure. The isolation reduces the maximum temperature at the aluminum sub-structure to about 100 

degrees Celsius over a thickness of approx. 45 mm. This isolation 

 41.

242
Back to the “Ablative” Future

system was developed and manufactured by ASTRIUM. The 

entire nose system has a mass of approx. 13.2 kg, whereby the 

nose shell alone accounts for 7 kg.42

The cap could withstand temperatures of up to 1,750 °C, and by the end 

of January 2001, the X-38’s nose cap had been delivered to JSC, where it languished when the program was canceled.

Although the CRV used proven technology for most of its systems, it quickly 

became mired in cost and schedule problems. The technical complexity of the 

development task, the international character of the effort, and the time phase 

for completion all conspired to push costs beyond acceptable bounds. Original 

estimates to build a capsule-type CRV amounted to more than $2 billion in 

total development cost. After paring down its efforts, the X-38 program team 

arrived at a cost estimate of $1.3 billion, but only with several development 

and acquisition assumptions yet to be verified. A review of the ISS program 

costs in the fall of 2001 noted that for the CRV and other elements, “There is 

inadequate current costing information associated with the non-U.S. components.” It also found that “Project interruptions will have cost impact on all of 

the elements under consideration.”43

By the spring of 2001 the CRV had been all but eliminated, not so much 

because of its cost overruns (although that did contribute), but because of 

huge overruns throughout the ISS program. Looking for places to trim the ISS 

program budget, the CRV proved an easy target. NASA Administrator Daniel 

S. Goldin testified to Congress in April 2001:

However, the U.S. CRV has a significant set of design activities to 

accomplish before we are ready to enter into a production contract. 

Just last year, NASA’s Integrated Action Team, focusing on program 

management excellence, concluded that technology risk reduction 

programs and design definition must be concluded before committing to production contracts to best insure that cost, schedule and 

technical targets can be realized. Given the magnitude of planned 

funding dedicated to CRV and the remaining definition work, 

funding allocated for the CRV production phase has been redirected to help resolve ISS core content budgetary shortfalls. NASA 

has initiated discussions with the European Space Agency (ESA) on 

 42

243
Coming Home

a role in the CRV project. Critical efforts such as X-38 atmospheric 

flight testing and some preliminary CRV design work and linkages 

with CTV under SLI will continue so as to maintain viable options 

for future CRV development. The planned space flight test of the 

X-38 is under review as part of the program assessment.44

When asked at the time if NASA was giving up the CRV, Goldin was 

emphatic that the CRV was necessary to the Station’s completion. As reported 

by Keith Cowing, “[Goldin] said that NASA was pursuing various options with 

the international partners on the program for these capabilities. As far as the 

CRV’s precursor, the X-38, Goldin said that there were funds to take that system 

up to and through orbital testing and that this would be done before NASA 

makes any commitment on funding the development of a CRV system.”45 But 

the newly arrived George W. Bush administration directed NASA to halt further 

development of the CRV. The CRV could have provided the ISS with a permanent replacement for the three-person Soyuz spacecraft currently being used as 

an interim CRV.46 NASA still sought to find ways to secure a CRV without direct 

Agency funding, particularly by having one or more of the international partners take it on, but the results yielded nothing. To this day, the ISS still relies on 

Russian-built Soyuz capsules, each of which could return a crew of three to Earth.

Bringing Deep Space Probes Back to Earth: 

Genesis and Stardust

With the diversion of the X-33, X-34, and X-38 lifting reentry efforts, NASA 

also pursued ablative technology for two small reentry vehicles for science 

sample collection and return to Earth. The first was the Genesis mission, an 

effort to collect particles from the solar wind and return them to Earth for 

laboratory study. Launched on August 8, 2002, atop a Delta II rocket, this 

spacecraft traveled to a point about 1 million miles from Earth, where it entered 

a halo orbit around the L1 Lagrangian point between Earth and the Sun. In 

 44.

244
Back to the “Ablative” Future

This graph shows the events that were to occur during the recovery of the Genesis sample 

return capsule in September 2004. The sample return capsule separated from the remainder 

of the spacecraft about 4 hours before it encountered Earth’s atmosphere. Friction initially 

slowed the capsule, followed by deployment of two parachutes, a small drogue chute, and then 

a larger main parachute. The parachutes malfunctioned, and two helicopters waiting to grab 

the capsule’s parachute in midair were unable to complete the recovery. The midair catch was 

to prevent the capsule from coming into contact with Earth materials that could contaminate its 

cargo of solar wind samples. Unfortunately, it did not work that way. NASA.

April 2004, after 2 1⁄2 years of collecting solar particles, Genesis left the L1 

halo orbit. After a flyby of the Moon, Genesis released its return capsule and 

returned to the L1 halo orbit. The Genesis return capsule entered Earth’s atmosphere on September 8, 2004, but the parachute failed to deploy, causing the 

capsule to impact the ground at a high velocity at the Utah Test and Training 

Range. Even so, the majority of the science goals were later met.47

The ablative heat shields of early space missions—especially Mercury, Gemini, 

and Apollo—were built specifically for the proposed use, were tailored to a wide 

 47.

245
Coming Home

array of missions scenarios, and were reengineered for each new flight regime. But 

since the Viking Lander program to Mars in the mid-1970s, NASA had ceased 

efforts to develop new ablative systems in favor of reusable TPS supporting the 

Space Shuttle. Accordingly, as one analysis commented:

As an example, the Pioneer Venus and Galileo missions employed 

fully dense carbon phenolic that was developed by the United States 

Air Force for ballistic missile applications. Over the past 30 years 

NASA adopted a “risk averse” philosophy relative to TPS, i.e., use 

what was used before, even if it was not optimal, since it had been 

flight-qualified. An unintended consequence was that the ablative 

TPS community in the United States slowly disappeared.

 NASA completed a number of relatively simple missions using 

TPS materials developed largely during the 1950s and 1960s. This 

began to change with the Stardust mission since it used a new ablative TPS because proven technology was not sufficient to ensure the 

safety of the spacecraft during reentry from deep space.48

As intended, the Genesis return to Earth mission was an extension of earlier 

efforts for a sample return from superorbital missions. One study laid out the 

manner in which the mission was to unfold as follows:

About four hours before Earth entry, the spacecraft reorients to 

the sample return capsule release attitude, spins up to 15 rpm and 

releases the capsule. Soon after release, the spacecraft reorients to 

joint its thrusters to Earth and performs a maneuver which will 

cause the spacecraft to enter Earth’s atmosphere, but break-up 

over the Pacific Ocean. Following release from the spacecraft, the 

Genesis sample return capsule experiences a passive, spin-stabilized 

aero-ballistic entry, similar to that of the Stardust mission. When 

the capsule has decelerated to 1.4 times the speed of sound, the onboard avionics system fires a mortar to deploy the drogue parachute. 

The drogue is a disk-gap-band design, with heritage dating to the 

Viking program, and an extensive history of supersonic applications. It serves both to increase the deceleration of the capsule, and 

to stabilize it through the transonic phase. As the capsule descends 

into the airspace of the Utah Test & Training Range (UTTR), 

 48.

246
Back to the “Ablative” Future

The Genesis return capsule entered Earth’s atmosphere on September 8, 2004, but the parachute failed to deploy, causing the capsule to impact the ground at a high velocity at the Utah 

Test and Training Range. This is the Genesis return capsule after impact. Despite the apparent 

damage, the majority of the science goals were later met. NASA.

recovery helicopters are directed to fly toward the intercept point. 

The capsule’s ballistic path is designed for delivery within an 84 × 30 

km footprint with subsequent reduction to a 42 × 10 km helicopter 

zone.... The first helicopter on-site will line up and match descent 

rate, then execute a Mid-Air Retrieval (MAR) capture. If a pass is 

aborted for any reason the pilot can line up and repeat.49

If only it had worked as intended. Instead, Genesis crash-landed in the 

Utah desert.

While this study is about return to Earth, the story of the Genesis and 

Stardust thermal protection systems was informed by a major effort to develop 

the entry vehicle that Galileo carried into Jupiter. Undertaken in the late 1980s 

and early 1990s, development of the Galileo probe’s thermal protection system 

proved taxing. Bernard Laub concluded the following:

The Galileo probe to Jupiter was the most challenging entry mission ever undertaken by NASA. The probe employed a 45 deg 

 49.

247
Coming Home

blunt cone aeroshell and it entered the Jovian atmosphere at a 

velocity of ≈ 47.4 km/s. The forebody TPS employed fully dense 

carbon phenolic ( = 1450 kg/m3

) that, at the time, was the best 

ablator available. The entry environment was very severe and estimates of the peak heating (combined convective and radiative) 

were on the order of 35 kW/cm2 with a total integrated heat load 

of ≈ 200 kJ/cm2

.

Because of this experience, NASA engineers concluded that using new 

materials to replace the carbon phenolic ablative thermal protection systems 

was appropriate for future missions.50

The heat shield for the Genesis return capsule used a carbon-bonded carbon 

fiber (CBCF) insulating material as a base, an approach used with success in several 

other NASA missions. The Genesis thermal protection system also included phenolic-impregnated carbonaceous ablator (PICA) and toughened unipiece fibrousreinforced oxidation-resistant composite (TUFROC). As one study noted:

CBCF utilized in the above mentioned TPS systems is an attractive substrate material because of its low density and high porosity, superior thermal performance, and compatibility with other 

components. In addition, it is low cost because of the commercial 

market it also serves. However, the current CBCF manufacturing 

process does not produce materials engineered to the specifications NASA desires to put in place. These emerging and highly 

innovative TPS designs require material manufactured to specification.... The benefits derived include significantly improved 

flexibility for the TPS design engineer, as well as, more cost efficient CBCF derived TPS fabrication.51

As finally flown, the Genesis return capsule was an amalgam of existing 

technologies and newer concepts. Its backshell was made of SLA-561V “Mars” 

ablator in an aluminum-honeycomb substructure. Its main structure possessed 

a carbon-carbon skin with a continuous surface (multilayer fabric layup).52 As 

reported by NASA:

 50.

248
Back to the “Ablative” Future

The heat shield is made of a graphite-epoxy composite covered 

with a thermal protection system. The outermost thermal protection layer is made of carbon-carbon. The capsule heat shield 

remains attached to the capsule throughout descent and serves as 

a protective cover for the sample canister at touchdown. The heat 

shield is designed to remove more than 99 percent of the initial 

kinetic energy of the sample return capsule.

 The backshell structure is also made of a graphite-epoxy composite covered with a thermal protection system: a cork-based material called SLA-561V that was developed by Lockheed Martin for 

use on the Viking missions to Mars, and have been used on several 

missions including Genesis, Pathfinder, Stardust and the Mars 

Exploration Rover missions. The backshell provides the attachment 

points for the parachute system, and protects the capsule from the 

effects of recirculation flow of heat around the capsule....

 The parachute system consists of a mortar-deployed 2.1-meter 

(6.8-foot) drogue chute to provide stability at supersonic speeds, 

and a main chute 10.5 by 3.1 meters (about 34.6 by 12.1 feet).

 Inside the canister, a gas cartridge will pressurize a mortar tube 

and expel the drogue chute. The drogue chute will be deployed at an 

altitude of approximately 33 kilometers (108,000 feet) to provide 

stability to the capsule until the main chute is released. A gravityswitch sensor and timer will initiate release of the drogue chute. 

Based on information from timer and backup pressure transducers, a small pyrotechnic device will cut the drogue chute from the 

capsule at about 6.7 kilometers altitude (22,000 feet). As the drogue 

chute moves away, it will extract the main chute. At the time of 

capture, the capsule will be traveling forward at approximately 12 

meters per second (30 miles per hour) and descending at approximately 4 meters per second (9 miles per hour).53

On September 8, 2004, during reentry, the thermal protection system worked 

well, but the parachute system failed and the Genesis return capsule streaked 

into the ground at the Utah Test and Training Range, southwest of Salt Lake 

City. A mortar should have fired at 100,000 feet, releasing a drogue parachute 

to slow and stabilize the vehicle. Then, at 20,000 feet the drogue would have 

pulled out the main parafoil, and the spacecraft would slow to 9 mph, whereupon 

 53.

249
Coming Home

helicopters could capture the capsule from midair using a hook. However, the 

helicopter missed it and the capsule plowed into the ground. The 5-foot-wide, 

420-pound capsule burrowed into the desert floor at a speed of 193 mph. In 

the end, NASA scientists were able to recover some of the solar-wind particles 

from the Genesis sample return capsule, and they continue to analyze results 

from this data.54

The Stardust sample return mission turned out somewhat better, at least 

in the effectiveness of the reentry and recovery system. Stardust was launched 

in 1999 and was recovered in 2006, and it was the first U.S. space mission 

dedicated solely to returning extraterrestrial material from beyond the Moon. It 

collected samples from Comet Wild 2 and from interstellar dust. The Stardust 

return system had six major components: a heat shield, back shell, sample 

canister, sample collector grids, parachute system, and avionics. The canister 

was sealed in an exterior shell that protected the samples from the heat of 

reentry. Scientists believe the material Stardust returned could date from the 

formation of the solar system, and scientific studies of the samples may alter 

our understanding of the universe. One major discovery is that ice-rich comets, 

the coldest and most distant bodies in the solar system, contain fragments of 

materials that make up the terrestrial planets.

The Stardust thermal protection system worked well during the mission. 

NASA realized that it required new technologies for reentry as its planetary 

exploration ramped up in the early/mid-1990s.55 Accordingly, it pursued “development of two new lightweight ablators, Phenolic Impregnated Carbon Ablator 

(PICA) and Silicone Impregnated Reusable Ceramic Ablator (SIRCA). Owing 

to its performance in the ~ 1 kW/cm2 heating environment and low heat shield 

mass, PICA enabled the Stardust Sample Return Mission. SIRCA has been used 

on the backshell of Mars Pathfinder and the Mars Exploration Rover missions.”56

Furthermore, as one report elaborated:

 54.

250
Back to the “Ablative” Future

Artist’s rendering of the Stardust spacecraft. The spacecraft was launched on February 7, 1999, 

from Cape Canaveral Air Force Station aboard a Delta II rocket. The primary goal of Stardust was 

to collect dust and carbon-based samples during its closest encounter with Comet Wild 2—

pronounced “Vilt 2” after the name of its Swiss discoverer. NASA PIA03183.

251
Coming Home

PICA was developed by NASA Ames in the early-mid 90s, is 

fabricated by Fiber Materials, Inc. (FMI) and was employed as the 

forebody TPS on the Stardust Return Capsule. It is currently the 

baseline forebody TPS for the Orion Crew Exploration Vehicle 

(CEV) and is being fabricated as the forebody TPS for the Mars 

Science Laboratory (MSL), scheduled for launch in mid-2009. 

PICA is a low density carbon-based ablator. Under both the CEV 

TPS Advanced Development Program and MSL [Mars Science 

Laboratory], an extensive data base has been developed. The failure modes are well-understood (upper heat flux limits of ≈ 1500 

W/cm2 and pressure of 1.0-1.5 atm) and validated design models 

have been developed. For Stardust, the PICA heat shield was fabricated as one piece. But that will not be possible for larger vehicles 

leading to a tiled design (such as the MSL design) that introduces 

significant design and fabrication complexities.57

PICA has been flight-qualified and used in the past, and along with investments in new thermal protection system materials, it has offered acceptable, 

cost-effective results.58

Because of the high velocity of reentry for Stardust, the vehicle’s forebody 

thermal protection systems had to use ablative materials, even as the afterbody’s blunt-cone aeroshell’s heating was much less hazardous to the spacecraft. 

Because of the stresses of forebody heating, NASA engineers used a complex 

analysis to arrive at the proper materials for its ablative heat shield:

To meet the requirements for the Stardust mission, one of a family 

of lightweight ceramic ablator materials developed at NASA Ames 

Research Center was selected for the forebody heat shield of the 

Stardust sample return capsule. This material, phenolic impregnated carbon ablator (PICA), consists of a commercially available 

low density carbon fiber matrix substrate impregnated with phenolic resin.... The Stardust program resulted in intensive material 

development, modeling, and testing efforts to provide a heat shield 

 57.

252
Back to the “Ablative” Future

NASA’s Stardust sample return capsule successfully landed at the U.S. Air Force Utah Test and 

Training Range on January 15, 2006. The capsule contained cometary and interstellar samples 

gathered by the Stardust spacecraft. NASA PIA03669.

for the high convective heating conditions expected during Earth 

entry while under constraints of limited time and funding.59

As it turned out, the Stardust reentry system worked well, dissipating about 

90 percent of the total energy via the bow-shock heating of the atmospheric 

gases. Additionally, the lessons learned from this effort fed into the next major 

project NASA undertook for reentry and recovery from space.

The Orion Capsule as the Next 

Generation Piloted Vehicle

As the United States entered the first decade of the 21st century, its piloted 

space flight vehicle, the Space Shuttle, had been in use for 20 years, and the 

technology was being superseded by more advanced concepts. Even so, NASA 

 59.

253
Coming Home

had not been successful in replacing the vehicle with a viable successor despite 

having tried to do so with NASP, the X-33, the X-34, and other projects that 

did not reach the hardware stage. This situation reached crisis proportions in 

the aftermath of the tragic Columbia accident, on February 1, 2003, when the 

vehicle and its crew were lost during reentry. With the intention of refocusing 

the space agency’s human space flight efforts, on January 14, 2004, President 

George W. Bush announced the “Vision for Space Exploration,” which was 

aimed at human exploration of the Moon and Mars. As stated at the time, the 

fundamental goal of this vision was to advance U.S. scientific, security, and 

economic interests through a robust space exploration program. In support of 

this goal, the President announced that the Nation would do the following:

•	

•	

•	

•	

Implement a sustained and affordable human and robotic program 

to explore the solar system and beyond;

Extend human presence across the solar system, starting with a 

human return to the Moon by the year 2020, in preparation for 

human exploration of Mars and other destinations;

Develop the innovative technologies, knowledge, and infrastructures 

both to explore and to support decisions about the destinations for 

human exploration; and

Promote international and commercial participation in exploration 

to further U.S. scientific, security, and economic interests.

In so doing, the President called for completion of the ISS and retirement 

of the Space Shuttle fleet by 2010. Resources that would have been expended 

in keeping the Shuttle program operating could then go toward creating the 

technologies necessary to return to the Moon and eventually to Mars. By 

2009, however, it had become clear that this program was unsustainable, and 

after a review of the program by a blue-ribbon panel led by former Lockheed 

Martin Chief Executive Officer Norm Augustine the new administration of 

President Barack Obama terminated the program in favor of another approach. 

Accordingly, the “Vision for Space Exploration” followed the path of the 

aborted Space Exploration Initiative, which was announced with great fanfare 

in 1989 but derailed in the early 1990s.60

 60.

254
Back to the “Ablative” Future

In the meantime, NASA had begun efforts to build a new spacecraft to carry 

out the expanded space exploration mission announced in the “Vision for Space 

Exploration.” This was intended to allow humanity to move beyond Earth toward 

multiplanetary activities. Named the Constellation program, it called for the 

reuse of as much existing Space Shuttle technology as possible to create a new 

Ares I crew launch vehicle. The vehicle was to consist of a Space Shuttle solid 

rocket booster as a first stage and an external tank as the beginning point for a 

second stage. A piloted space capsule, Orion, was to sit atop this system. A second 

rocket, the proposed Ares V cargo launch vehicle, would provide the heavy-lift 

capability necessary to journey back to the Moon or to go beyond. Ares I was 

intended to carry a crew of up to six astronauts to low-Earth orbit in the Orion 

crew exploration vehicle, with the capability for expanding its use to send four 

astronauts to the Moon. Ares V was intended to serve as the Agency’s primary 

vehicle for delivery of large-scale hardware and cargo into space.

The Orion spacecraft, which began development as the Crew Exploration 

Vehicle (CEV), was only one part of the Constellation program’s fleet of launchers and other vehicles designed to take humans out of Earth’s orbit. According 

to NASA specifications:

Orion borrows its shape from space capsules of the past, but takes 

advantage of the latest technology in computers, electronics, life 

support, propulsion and heat protection systems. The capsule’s 

conical shape is the safest and most reliable for re-entering the 

Earth’s atmosphere, especially at the velocities required for a direct 

return from the moon.

 Orion will be 16.5 feet in diameter and have a mass of about 25 

tons. Inside, it will have more than 2.5 times the volume of an Apollo 

capsule. The spacecraft will return humans to the moon to stay for 

long periods as a testing ground for the longer journey to Mars.61

Conceived by NASA and built by Lockheed Martin, the Orion was originally intended to be all things to all people. It could fly to the ISS but just as 

easily travel to and from the Moon. Over time, NASA had to back away from 

that requirement in favor of more modest Earth-orbital capabilities, with the 

potential for a hardier translunar version of Orion in the future.62

 61.

255
Coming Home

The Ares V (left) and Ares I (right) were part of an ambitious plan to 

replace the Space Shuttle with Shuttle-derived hardware that would 

carry out the “Vision for Space Exploration.” Part of the Constellation 

program, announced in 2005, it was canceled in 2010 in favor of a 

commercially derived piloted space launcher. NASA.

256
Back to the “Ablative” Future

The thermal protection system for Orion went through several considerations until finally, in September 2006, NASA awarded a contract to Boeing 

for its development. As Bill Christensen wrote in Space.com at the time:

NASA Ames Research Center has awarded the $14 million contract for the development of a phenolic impregnated carbon ablator (PICA) heat shield. The PICA shield was first used on the 

Stardust interplanetary spacecraft launched February 7, 1999 

to study the composition of the Wild 2 comet. It successfully 

reentered Earth’s atmosphere January 15, 2006. The capsule was 

traveling at 28,900 miles per hour (46,510 kilometer per second), 

the fastest reentry speed ever achieved by a man-made object....

 The ability to survive high speed reentry is an important consideration for the Orion spacecraft. Orion is intended to perform 

lunar-direct returns, which result in considerably higher speeds; the 

spacecraft will need to withstand about five times more heat than experienced by spacecraft returning from the International Space Station.

 The best protection against high heat flux is an ablative heat 

shield. The extreme heat of reentry causes the material to pyrolize—

the chemical decomposition of a material by heating in the absence 

of oxygen. As the PICA chars, melts and sublimates, it creates a cool 

boundary layer through blowing, protecting the spacecraft.63

To determine how best to proceed with development of the new heat shield, 

NASA undertook study over a 3-year period with Boeing and other entities 

aimed at ensuring success; its Orion Thermal Protection System Advanced 

Development Project considered eight different candidate materials.

The two final candidates for the Orion thermal protection system were Avcoat 

and phenolic-impregnated carbon ablator, both of which had proven track records 

in space reentry and recovery operations. NASA announced on April 7, 2009:

Avcoat was used for the Apollo capsule heat shield and on select 

regions of the space shuttle orbiter in its earliest flights. It was put 

 63.

257
Coming Home

The Orion Crew Exploration Vehicle undergoes a test in 2010. This basic vehicle survived the 

cancellation of the Constellation Program and morphed into the Multi-Purpose Crew Vehicle, 

which was still called Orion. NASA JSC-2010E042235.

back into production for the study. It is made of silica fibers with 

an epoxy-novalic resin filled in a fiberglass-phenolic honeycomb 

and is manufactured directly onto the heat shield substructure and 

attached as a unit to the crew module during spacecraft assembly. 

PICA, which is manufactured in blocks and attached to the vehicle after fabrication, was used on Stardust, NASA’s first robotic 

space mission dedicated solely to exploring a comet, and the first 

sample return mission since Apollo.

 “NASA made a significant technology development effort, 

conducted thousands of tests, and tapped into the facilities, talents and resources across the agency to understand how these 

materials would perform on Orion’s five-meter wide heat shield,” 

said James Reuther, the project manager of the study at NASA’s 

Ames Research Center at Moffett Field, Calif. “We manufactured 

full-scale demonstrations to prove they could be efficiently and 

reliably produced for Orion.”

 Ames led the study in cooperation with experts from across 

the agency. Engineers performed rigorous thermal, structural and 

environmental testing on both candidate materials. The team 

then compared the materials based on mass, thermal and structural performance, life cycle costs, manufacturability, reliability 

and certification challenges. NASA, working with Orion prime 

258
Back to the “Ablative” Future

contractor Lockheed Martin, recommended Avcoat as the more 

robust, reliable and mature system.

 “The biggest challenge with Avcoat has been reviving the technology to manufacture the material such that its performance is 

similar to what was demonstrated during the Apollo missions,” 

said John Kowal, Orion’s thermal protection system manager at 

Johnson. “Once that had been accomplished, the system evaluations clearly indicated that Avcoat was the preferred system.”

 Even with this work, NASA was unsure what direction to 

go in its decision-making process. “While Avcoat was selected as 

the better of the two candidates,” NASA engineers said, “more 

research is needed to integrate it completely into Orion’s design.”64

Another problem quickly emerged once NASA began concentrating on 

Avcoat as a possible Orion thermal protection system material. As a proprietary material, the recipe for Avcoat was hard to track down; and since Avcoat 

had not been made for some 40 years, those familiar with the formula were 

few and far between. At the same time, the Orion spacecraft would have to 

withstand approximately five times greater reentry heating than the Apollo 

missions endured. So even if Avcoat could be duplicated, it had to be stronger 

and thicker. “We can handle the initial operating system of (reentry from) lowEarth orbit,” said James Reuther, head of Orion’s Thermal Protection System 

Advanced Development Project for NASA’s Ames Research Center. “[O]n the 

lunar side it’s a much greater challenge. We need a single heat-shield material 

for the lunar environment and reentry.... We’re at greater risk there building 

a single system for both from scratch.”65

To help with the Avcoat research, NASA came to the Smithsonian Institution’s 

National Air and Space Museum to analyze parts of the heat shield from the 

Apollo spacecraft that was flown 40 years ago. “ ‘We started working together 

at the end of June [2009] to track down any Apollo-era heat shields that they 

had in storage,’ said Elizabeth (Betsy) Pugel of the Detector Systems Branch at 

NASA Goddard. ‘We located one and opened it. It was like a nerd Christmas 

for us!’ ” They examined the material, the connections to the spacecraft, and 

 64.

259
Coming Home

the properties of the Avcoat.66 The challenges were significant, even with what 

was learned through examining the Apollo heat shields. The Orion capsule was 

expected to withstand as much as 5,000 °F during a return to Earth from the 

lunar mission, and the experience of the Space Shuttle was only about 2,300 °F.67

Despite the challenges, work continued apace, and virtually every NASA 

Center had some role in it. In August 2008, NASA and Lockheed Martin engineers were successful in bonding the backshell to the demonstration spacecraft 

and began testing the PICA ablator to determine compression loads, vibration 

parameters, and reaction to acoustical patterns.68 NASA engineers at Langley 

Research Center added their expertise and undertook tests on PICA and Avcoat 

to discover the thermal and structural behavior of the ablators attached to a 

honeycomb backing structure.69 Meanwhile, NASA Glenn Research Center 

engineers developed especially strong bonding to affix the ablator to the spacecraft.70 Among many other efforts, a team from NASA’s Johnson Space Center, 

Ames Research Center, and Engineering Research Consultants Inc. undertook 

research to “improve the Orion (new launch vehicle) vehicle’s thermal protection system (TPS) properties”71 by adding new elements to the Avcoat formula, 

perhaps even carbon nanotubes. This report noted:

Small additions of carbon nanotubes to the phenolic polymer 

have potential to greatly improve char strength. Char is formed 

as a thin layer as heat from entry pyrolizes PICA’s leading surface. 

Aerodynamic heat and pressure from lunar and martian returns 

 66.

260
Back to the “Ablative” Future

Engineers performed a series of tests in the 20-foot vertical spin tunnel in 2010 that measured 

the aerodynamics of a 6.25-percent model of the Orion crew module with deployed drogue 

parachutes. The purpose of the test was to compare how the model performed in simulated 

flight versus a free-flight flight test in the tunnel. NASA.

261
Coming Home

will push the ablator to its limits. If the char is too weak, pressure from hot internal gases can cause the char to spall or break 

off, thus exposing virgin PICA. This causes uneven heating and 

disturbances to the boundary layer. Carbon nanotubes increase 

char strength by adding an additional fibrous structure to the 

otherwise randomly orientated pyroloisis region.72

Of course, this effort is far from complete and may never enter into the 

Orion thermal protection system. Further study is ongoing.

In the end, the Orion thermal protection system will be a passive ablator that will work much like the system used on Apollo. The higher reentry 

speeds and harsher flight regime for a lunar reentry have been put off a bit 

with concentration on the Earth orbital reentry and recovery. It will consist of 

the following elements:

For the CEV CM, spacecraft protection is the TPS that includes 

ablative TPS on the windward (aft) side of the vehicle, reusable surface insulation for the external leeward (central and forward) TPS, 

and internal insulation between the pressurized structure and OML. 

There are a number of potential materials available for use in the CEV 

CM protection system and the eventual TPS materials selected will 

be the result of a rigorous trade study based on performance and 

cost. Some of these materials may include carbon-carbon, carbonphenolic, AVCOAT, Phenolic Impregnated Carbonaceous Ablator 

(PICA), PhenCarb-28, Alumina Enhanced Thermal Barrier-8 

(AETB–8))/TUFI, Advanced Flexible Reusable Surface Insulation 

(AFRSI, LI-900 or LI-2200, CRI, SLA-561S, cork, and many others.

 TPS mass for the present CEV CM concept is scaled from an 

analysis conducted for a vehicle of the same base diameter but lower 

sidewall angle and higher mass at Entry Interface (EI). A 5.5-m, 

28-deg sidewall concept with a total mass of approximately 11,400 

kg requires an aft TPS mass of 630 kg and forward TPS mass of 

180 kg. The assumed TPS materials for this analysis were PICA 

for the aft side and a combination of LI-2200, LI-900, AFRSI, 

and Flexible Reusable Surface Insulation (FRSI) at equal thicknesses for the central and forward side. The maximum heating rate 

for the TPS is driven by ballistic entry trajectories at lunar return 

speeds (11 km/s), and TPS thickness is sized by the total integrated 

 72.

262
Back to the “Ablative” Future

heat load of a skip-entry trajectory. For the lighter 5.5-m, 32.5-deg 

CM, the 630-kg aft TPS mass from the larger, heavier concept has 

been retained to provide additional margin, while the central and 

forward TPS mass has been scaled based on the lower surface area. 

The current CEV CM mass, including external TPS, is 9,301 kg 

at atmospheric EI for the nominal lunar mission.73

It remains to be seen how this will work out, but the Orion program has 

“invested considerable resources in developing analytical models for PICA and 

AVCOAT, material property measurements that is essential to the design of the 

heat-shield, in arc-jet testing, in understanding the differences between different arc jet facilities...and in integration of and manufacturing [the] heat shield 

as a system.” Since 2005, significant strides have been made in understanding 

the problem and working toward its resolution. In the process, researchers have 

advanced the state of the art in thermal protection system technology, even as 

Orion engineers understood that much more remained to be done.74

One other issue affecting the recovery from space was the development of 

a new parachute system that would carry Orion home. These relatively simple 

systems have flown reliably in all but a handful of cases since the beginning 

of the space age, although there was a problem with the Genesis parachute. 

As envisioned, the CEV Parachute Assembly System (CPAS) would rely on 

proven methods of past experience:

The CPAS will be deployed in subsonic controlled flight with the 

angle of attack and sideslip within 40 degrees of trim, the vehicle pitch 

and yaw rates less than 40 deg/sec and the vehicle roll rate less than 80 

deg/sec. The CPAS will be initiated at altitudes ranging from 4,000 

to 40,000 ft AGL [above ground level]. The vehicle weight at drogue 

deploy is assumed to be 17,176 lbs. The drogues will be deployed 

at dynamic pressures ranging from 19 to 115 psf and be required to 

stabilize and decelerate the vehicle prior to release and deployment of 

the mains. The vehicle weight at touchdown is assumed to be 14,400 

lbs. The CPAS shall nominally deliver no greater than 26 ft/sec rate of 

descent and no greater than 33 ft/sec with one main failed using an air 

density of 0.00182526 (representing a three sigma dispersed hot day 

 73.

263
Coming Home

at White Sands Missile Range). The CPAS shall stabilize the vehicle 

to within +/- 5 degrees of the desired hang angle, recognizing that 

the hang angle for land and water landings will likely be different.75

The configuration was to be similar to Apollo’s, with dual drogue chutes, 

both sufficient to deploy the main parachutes. In addition, there would be three 

main parachute systems, any two of which could land the spacecraft safely. It 

would also employ pilot parachutes to deploy each main parachute individually; this was especially attractive because of the bell-shaped geometry of the 

spacecraft—like Apollo—which used a similar system.76

A New President Changes the Paradigm

Despite this success, by the end of the administration of President George W. 

Bush it had become clear that the Constellation program was not progressing 

as well as intended. A 2008 report summarized attitudes toward the program 

during the last year of the Bush administration as follows:

Congress has been debating the Vision, including its impact on 

the shuttle and on U.S. human access to space. Some Members 

wanted to terminate the shuttle earlier than 2010 because they feel 

it is too risky and/or that the funds should be spent on accelerating 

the Vision. Others want to retain the shuttle at least until a new 

spacecraft, the Crew Exploration Vehicle (CEV), is available to 

take astronauts to and from the ISS. The CEV is now planned for 

2015 at the earliest, leaving a multi-year gap during which U.S. 

astronauts would have to rely on Russia for access to the ISS.77

Even before the presidential election of 2008, however, it had become 

highly uncertain that the Constellation program would continue. Virtually 

 75.

264
Back to the “Ablative” Future

no political will existed in the Bush administration to make it a reality, and 

certainly not much additional funding was forthcoming.

After President Barack Obama took office in January 2009, he set about 

organizing his administration. During this organization process, the President 

received several reports from a variety of sources, including his own transition team for NASA, stating that the Constellation program was very much 

over budget and behind schedule. There was a much-reported confrontation 

between NASA Administrator Michael D. Griffin and the head of the transition team (and future Obama appointee as NASA Deputy Administrator) 

Lori B. Garver at a December 2008 event at NASA Headquarters. Garver and 

Griffin spoke quietly about the situation but were overheard. Garver used the 

metaphor of purchasing a vehicle: “Mike, I don’t understand what the problem 

is. We are just trying to look under the hood.” Griffin’s response was defensive. 

“If you are looking under the hood, then you are calling me a liar,” Griffin 

replied. “Because it means you don’t trust what I say is under the hood.”78

Administrator Griffin publicly voiced his complaints about those expressing 

concern about the Constellation program, in April 2009, when delivering a 

speech before the National Space Club:

I’ve grown impatient with the argument that Orion and Ares 1 are 

not perfect, and should be supplanted with other designs. I don’t 

agree that there is a better approach for the money, but if there were, 

so what? Any proposed approach would need to be enormously better to justify wiping out four years worth of solid progress. Engineers 

do not deal with “perfect”. Your viewgraphs will always be better 

than my hardware. A fictional space program will always be faster, 

better, and cheaper than a real space program.79

Impatience aside, a cacophony of criticism from many quarters required 

investigation and perhaps alteration of the Constellation program.

What was obvious among space-policy decision makers working with the 

new administration was that the technological and budgetary challenges of the 

Constellation program had exploded since its baseline in 2005 and that the 

Obama administration had to take action. The President appointed a blue-ribbon 

panel to review the program and make recommendations. The original desire 

 78.

265
Coming Home

was that a new NASA Administrator would be nominated before January 20, 

2009, so that this person could commission a review. When the Administrator’s 

selection was delayed, the White House decided to go ahead with the review. In 

the spring of 2009, President Obama tapped longtime aerospace official Norm 

Augustine to head the study effort. The President’s announcement stated:

The “Review of United States Human Space Flight Plans” is to 

examine ongoing and planned National Aeronautics and Space 

Administration (NASA) development activities, as well as potential alternatives, and present options for advancing a safe, innovative, affordable, and sustainable human space flight program 

in the years following Space Shuttle retirement. The panel will 

work closely with NASA and will seek input from Congress, the 

White House, the public, industry, and international partners as 

it develops its options. It is to present its results in time to support 

an Administration decision on the way forward by August 2009.

The President’s mandate emphasized a four-part agenda: “1) expediting a 

new U.S. capability to support use of the International Space Station; 2) supporting missions to the Moon and other destinations beyond low-Earth orbit; 

3) stimulating commercial space flight capabilities; and 4) fitting within the 

current budget profile for NASA exploration activities.”80

Augustine’s panel’s report was submitted to the White House in the fall of 

2009. It validated the growing concerns that the Constellation program would 

not be sustainable because of the pressures of budget, technology, and time. 

Specifically, Augustine’s panel concluded in its summary report:

•	

•	

•	

Human exploration beyond low-Earth orbit is not viable under the 

FY 2010 budget guideline.

Meaningful human exploration is possible under a less constrained 

budget, ramping to approximately $3 billion per year above the FY 

2010 guidance in total resources.

Funding at the increased level would allow either an exploration 

program to explore Moon First or one that follows a Flexible Path of 

exploration. Either could produce results in a reasonable timeframe.81

 80.

266
Back to the “Ablative” Future

The final report offered more detail and went into a sustained discussion 

of something hinted at in the summary: the possibilities for commercial crew 

and cargo support for the International Space Station.

A key discussion in the final report involved the harnessing of private-sector, 

especially entrepreneurial, firms in supporting the International Space Station, 

instead of relying on the Constellation program. The report noted that a major 

gap would occur between the retirement of the Space Shuttle and the first use of 

any Constellation hardware: “Under current conditions, the gap in U.S. ability 

to launch astronauts into space will stretch to at least seven years. The Committee 

did not identify any credible approach employing new capabilities that could 

shorten the gap to less than six years.” This would be true even with increased 

funding for NASA’s program. The panel noted that a $3 billion-per-year increase 

for fiscal years 2010 to 2014 could return the Constellation program to health. 

But there was another option. “As we move from the complex, reusable Shuttle 

back to a simpler, smaller capsule, it is appropriate to consider turning this transport service over to the commercial sector,” the panel concluded. “This approach 

is not without technical and programmatic risks, but it creates the possibility of 

lower operating costs for the system and potentially accelerates the availability of 

U.S. access to low-Earth orbit by about a year, to 2016. If this option is chosen, 

the Committee suggests establishing a new competition for this service, in which 

both large and small companies could participate.”82

The response to this report from the space community was immediate. Some 

administration officials urged that the President cancel Constellation. Edward 

Crawley, a Massachusetts Institute of Technology professor and a member of the 

Augustine panel, remarked that Ares I was suffering from technical issues that could 

only be overcome with more money and time. “It was a wise choice at the time,”83

said Crawley, when asked about originating the program in 2005. “But times have 

changed,” he added. “The budgetary environment is much more tight, and the 

understanding of the cost and schedule to develop the Ares I has matured.” Others 

were supportive of continuing Constellation. Constellation Program Manager Jeff 

Hanley defended the program and argued that the panel did not “take into account 

the improvements we have made in our schedule quality and risk posture through 

deployment of reserves and the reduction of program content.”84

 82.

267
Coming Home

In the midst of this turmoil, NASA made a test flight of the Ares I-X on 

October 28, 2009. There was a 2-minute powered flight segment, and the 

entire mission lasted about 6 minutes from launch until splashdown of the 

rocket’s booster stage nearly 150 miles downrange. The rocket reached nearly 

3 g’s and a speed of Mach 4.76 (not quite hypersonic speed). The dummy 

upper stage returned to the recovery area under parachutes at a suborbital 

altitude of 150,000 feet after the separation of its first stage, a four-segment 

solid-rocket booster. As this was not a test of the reentry and recovery system, 

the simulated upper stage, Orion crew module, and launch-abort system were 

not recovered. The results of this test were mixed; some criticized that the 

first stage was a surplus (shelf-life expired) four-segment Shuttle solid-rocket 

booster and not the proposed five-segment Ares first stage.85

Based on these responses, President Obama proposed on February 1, 2010, 

with more details added in a Presidential speech on April 15, a radical new path 

for future U.S. human space flight efforts. Central to this was the termination of 

the Constellation program as a single entity, the continuation of certain technology developments such as the Orion space capsule, the continuation of operations on the International Space Station until at least 2020, and the fostering 

of private-sector solutions to support operations in low-Earth orbit. Since this 

declaration on February 1, 2010, numerous high-profile space flight advocates 

weighed in on both sides of the debate. In April 2010, Apollo astronauts Neil 

Armstrong, Gene Cernan, and James A. Lovell, Jr., famously sent the President 

a letter warning that the proposed changes to human space flight “destines our 

nation to become one of second- or even third-rate stature.” Proponents of the 

plan, among them Apollo astronaut Buzz Aldrin, counter that the President’s 

approach will return NASA to its roots as an R&D organization while private 

firms operate space systems. Turning low-Earth orbit over to commercial entities 

could then empower NASA to focus on deep space exploration, perhaps eventually sending humans to Mars or elsewhere.86

The debate has largely been over maintaining a traditional approach to 

human space flight with NASA dominating the effort, owning the vehicles, 

and operating them through contractors. That was the method whereby the 

United States went to the Moon; it has proven successful for over 50 years of 

 85.

268
Back to the “Ablative” Future

NASA’s Ares I-X demonstration vehicle on LC-39B at the Kennedy Space Center, FL, on October 

26, 2009. The flight test of Ares I-X, offered an opportunity to test and prove flight characteristics, 

hardware, facilities, and ground operations associated with the Ares I. NASA JSC2009E225491.

269
Coming Home

human space exploration. Then there are those from the “new space” world 

that emphasize allowing private-sector firms to seize the initiative and pursue 

entrepreneurial approaches to human space flight. Advocates of the more traditional approach believe that the other side will sacrifice safety; those who 

support the entrepreneurial approach say that those who support a more robust 

Government-led program are advocating large, over-budget, underachieving 

space efforts. It remained unclear how much, if any, of this new initiative 

that the U.S. Congress will approve; in 2011, there still was no resolution. 

Meanwhile, NASA began moving forward with the “program of record” and 

at the same time planning for programs to replace it.

NASA’s Commercial Crew and Cargo Program

During the next year, NASA pursued efforts to replace access to low-Earth orbit, 

previously provided by the Space Shuttle, through a multiphase space technology 

development program known as Commercial Crew Development (CCDev). 

Intended to stimulate development of privately operated crew vehicles to lowEarth orbit, CCDev’s first phase offered a token sum of $50 million during 2010 

to five American companies for R&D into human space flight concepts and 

technologies in the private sector. In its second phase, with contracts of $269 

million awarded to four firms in April 2011, the objective was to move toward 

the establishment of one or more orbital space flight capabilities on which 

NASA could purchase cargo and eventually transport crews into space.87 The 

concepts, which ranged from lifting body to capsule spacecrafts, were awarded 

to Blue Origin, Sierra Nevada Corp., Space Exploration Technologies (SpaceX), 

and Boeing. As announced by Ed Mango, NASA’s commercial crew program 

manager: “The next American-flagged vehicle to carry our astronauts into space 

is going to be a U.S. commercial provider. The partnerships NASA is forming with industry will support the development of multiple American systems 

capable of providing future access to low-Earth orbit.”88

The spacecraft nearest to being ready for flight might well be the SpaceX 

Dragon capsule launched atop the Falcon 9 rocket. With its successful suborbital 

 87.

270
Back to the “Ablative” Future

test flight on December 8, 2010, the SpaceX entry into this competition appeared 

destined for an early operational date. The reentry and recovery technology on 

Dragon were virtually identical to those used by the Orion vehicle, especially 

the PICA heat-shield technology pioneered at NASA. Its three-canopy parachute 

recovery system was also modeled on those used in earlier programs.89

The second awardee, garnering the largest award of $92.3 million, was 

Boeing. Its CST-100 crew capsule had been pursued as a support vehicle for 

the ISS. Very close in design to the Orion spacecraft of the Constellation program (but without the deep space capability), Boeing’s vehicle was intended 

to house crews as large as seven and could be attached to ISS for more than 

half a year before requiring relief. For reentry and recovery, Boeing intended 

to use three parachutes for landing and its Lightweight Ablator for the heat 

shield. In every case, Boeing has emphasized its long history in working 

on these technologies and the legacy reentry and recovery systems that the 

company intended to use in the CST-100. Boeing officials have projected 

flight for this vehicle in 2015.90

Another recipient of the CCDev award, but receiving only $22 million, 

was Blue Origin, a startup entrepreneurial firm that proposed developing a 

biconic orbital capsule launched atop an Atlas V. Information about this project 

is limited, and Blue Origin is reticent to speak in public about its activities. 

Dan Rasky of NASA Ames remarked: “I joke with people that if you want 

to see what a billionaire’s clubhouse looks like, go visit Blue Origin.”91 Little 

information is available as yet about the Blue Origin capsule’s thermal protection and recovery systems. In Phase 2A of the effort, the company would 

develop “a pressurized ‘Crew Return Vehicle’ capsule with TPS, solid escape 

system and a landing mechanism (whether that’s VTVL rocket or parachutes 

or both). Partially funded by CCDev. Give the solid escape system sufficient 

 89.

271
Coming Home

capability abort at liftoff and sufficient TPS return from Mach 10 trajectories 

(and structure to beef up TPS to orbital velocities).”92

 Finally, the CCDev program awarded an $80 million contract to the Sierra 

Nevada Corp. to build a commercial Space Transportation System based on NASA 

HL-20 and launched on an Atlas V. This Dream Chaser lifting body spacecraft 

would utilize Virgin Galactic’s carrier aircraft as a platform for atmospheric drop 

tests as early as 2012. As a lifting body, this vehicle is intended “to land on a 

runway and be reusable. Its carbon composite airframe looks large but because of 

the composite structure it weighs only 27,100 pounds. For all of the spaceplane’s 

sleekness, as one commentator noted, it could handle only 1500K cargo and only 

2 crew. This is a good looking vehicle, but it is not a mini Shuttle. It is a minuscule 

Shuttle.” 93 Its thermal protection system, at present, is unclear but will probably 

use a combination of RCC panels, ceramic tiles, and composites.

The Air Force and the X-37B

As NASA engaged in this highly public effort to develop a new human space 

flight capability to succeed the Space Shuttle and support the ISS, the U.S. Air 

Force quietly undertook its own spaceplane development effort: the X-37B. 

This program had originated as a NASA effort in August 1998, when a research 

announcement solicited proposals for “Future-X”—a flight demonstrator 

designed to validate emerging hypersonic technologies leading toward reductions in the cost of space access. The announcement specifically called for the 

development of propellant tanks, thermal protection systems, avionics, and 

structures—especially thorny technology issues that required more capable 

systems if efficiencies in space access were to be significantly advanced. NASA 

contracted with the Boeing Company of Seal Beach, CA, in July 1999, for 

a 4-year cooperative agreement to develop what became known as the X-37 

advanced-flight demonstrator. Not a large effort, the 4-year cooperative agreement eventually amounted to something over $500 million, with a 50/50 

cost-sharing ratio between Government and industry.94

 92.

272
Back to the “Ablative” Future

The X-37 represented the third in a succession of efforts to develop and test 

technologies required to build a spaceplane at NASA (after the X-33 and the 

X-34). This X-37 was intended to build on those earlier efforts and eventually 

reach orbital flight regimes. To be launched shrouded aboard an expendable 

launcher, the X-37 was a 120-percent scale derivative of the Air Force’s unpowered X-40, developed in the Air Force Space Maneuver Vehicle program, which 

flew seven successful unpowered approach and landing tests at Dryden Flight 

Test Center in 2001. The X-37 technology demonstrator effort was also intended 

to feed into NASA’s Orbital Space Plane program. Had the X-37 flown, NASA 

at one point intended to deploy it from the Space Shuttle, but in the context of 

changing plans for Shuttle replacement the program fell by the wayside.95

As DARPA was wrapping up airdrop tests on the Approach and Landing 

Test Vehicle (ALTV), the Air Force’s Rapid Capabilities Office (AFRCO) began 

contracting for the X-37 program. The X-37, now the X-37B, had its first drop 

test on April 7, 2006, at Edwards AFB, and flew an orbital test mission on April 

22, 2010. Launched inside a shroud atop an Atlas V rocket, the X-37B orbited 

Earth until December 3, 2010, at which point it became publicly known 

because of its success as a robotically landed spaceplane. This was the first test of 

the vehicle’s heat shield and hypersonic aerodynamic-handling characteristics. 

The launch of a second X-37B on March 5, 2011, and its successful operation 

in orbit suggest that this program shows great promise to, in the words of the 

Secretary of the Air Force, complete “risk reduction, experimentation, and 

operational concept development for reusable space vehicle technologies, in 

support of long-term developmental space objectives.”96

Developing an integrated thermal protection system for the X-37B was 

one of the key objectives of the program. This effort included the following:

•	

•	

−

−

−

Flight demonstration of an integrated thermal protection system.

Level 1 Requirement: provide maturation and validation of thermal 

protection system within the confines of the X-37 reentry heating 

environments, which include:

Leading edge >2,950 °F.

Acreage thermal protection system at 2,400 °F.

High-temp gap fillers and seals to support 2,950 °F leading 

edge and 2,400 °F acreage.

 95.

273
Coming Home

−

−

−

The X-37B Orbital Test Vehicle waits in the encapsulation cell at the Astrotech facility in Titusville, 

FL, on April 5, 2010. Half of the Atlas V 5-meter fairing is visible in the background. U.S. Air Force.

Durability/Re-Usability of thermal protection system better 

than existing Shuttle.

Thermal protection system components 10 times more durable 

than current tile in windward high-temperature environments.

New thermal protection system enables adverse weather flight 

conditions.97

The X-37B used a hot-structure control surface developed by Boeing–

Huntington Beach/Seal Beach with major support from GE Power Systems 

Composites, and with assistance from the Air Force Research Laboratory and 

NASA Langley Research Center. The effort required the design, analysis, manufacture, and testing of two flaperons and ruddervators for the X-37 reentry 

vehicle. From there, research and testing of full-acreage thermal protection 

systems followed, and they were proven in the first orbital flight. The thermal 

protection system used silica tiles impregnated with toughened unipiece fibrous 

insulation. This technology has been well-tested on the Space Shuttle since 

1994, when it was first used on STS-59. These tiles were hardier than earlier 

 97.

274
Back to the “Ablative” Future

Shuttle thermal protection systems elements and were used extensively on the 

X-37’s underside. These TUFI tiles are less susceptible to impacts since the 

surface material permeates the underlying insulation.

A public report on the X-37 program included the following description 

of the new developments in its thermal protection system:

The X-37’s most notable thermal advance is on the wing leading edge. On the shuttle, that vulnerable area was covered with 

reinforced carbon-carbon; the X-37 uses a different material, 

called TUFROC, for Toughened Uni-piece Fibrous Reinforced 

Oxidation-Resistant Composite. TUFROC (pronounced “tough 

rock”) was developed at NASA’s Ames Research Center in 

California by a group led by David Stewart, who has worked on 

thermal protection systems since the shuttle program.

 Stewart explains that during reentry, heat is generated not 

just by friction of the vehicle against the atmosphere, but also 

by atoms on the surface recombining. In the shuttle’s case, the 

carbon-carbon oxidizes. As the name implies, the new material 

resists oxidative damage. The surface of the shuttle’s tiles heats up 

very fast because the insulator’s high-density coating is very thin. 

TUFROC’s surface material is thicker, and therefore takes longer 

to heat up. And the new material will reduce weight, which will 

enable the spaceplane to carry more payload.98

Compared to the X-20 Dyna-Soar by project officials, the X-37 appears to 

hold promise as a future spaceplane that might satisfy both robotic and perhaps 

after modification, human missions.

Summary

Quite a lot of activity in the last decade has affected the course of reentry and 

recovery from space. After giving up on the winged, reusable Space Shuttle 

and its unique thermal protection system and runway landing capability, on 

January 14, 2004, President George W. Bush pressed the reset button by mandating that NASA focus on a new Moon/Mars exploration agenda using a 

 98.

275
Coming Home

The U.S. Air Force’s X-37B landed after several months in Earth orbit on December 3, 2010, 

at Vandenberg AFB, CA. Ironically, the runway at Vandenberg had been extensively modified to 

accommodate the Space Shuttle. U.S. Air Force.

capsule called the Crew Exploration Vehicle, thereby overturning any other 

initiative.99 That program remained in place only until 2009, when President 

Barack Obama overturned it in favor of a commercial effort to replace the 

Space Shuttle. Although the new commercial vehicle might also be a capsule, by 

2011 NASA’s CCDev program was supporting yet another design for a winged 

spaceplane “very reminiscent of the Shuttle’s design intended to ferry crews 

to and from the ISS.”100 Additionally, the CCDev program also supported 

three capsule concepts, and in every case the reentry and recovery systems 

 99.

276
Back to the “Ablative” Future

were reminiscent of earlier heat shields and parachute systems. Although only 

the Dragon capsule had flown, albeit on a suborbital trajectory, the challenges 

of reentry and recovery were not major areas of concern for these programs. 

Instead, the programs appeared intent on using legacy technologies pioneered 

by NASA in some cases as far back as the Apollo program. Thus far, it appears 

that only the Boeing thermal protection system, developed for other systems in 

the recent past, pushed the envelope of knowledge about reentry. As reported 

in 2010: “Both capsules and space planes have their advantages, and neither 

has a spotless safety record. But it will be interesting to see which mode NASA 

eventually selects for the next generation of ISS missions.”101

The X-37B is an entirely different type of program. Pursued as a technology demonstrator that may well lead to an operational orbital spaceplane, the 

Air Force has been diligent in pressing the frontiers of new technologies for 

thermal protection systems. By all accounts (although the details are classified 

for national security purposes), this system has been quite effective. As reported 

at the time of the first test flight:

The success of OTV-1 on its first orbital flight, USA-212, was 

nothing short of spectacular, accomplishing every mission objective with a precision that stunned the aerospace world. Despite 

blowing out a tire on landing, which caused some minor damage 

to the underside of OTV-1, the spacecraft is in excellent shape and 

being refurbished for a second flight sometime in late 2011. The 

second X-37B, OTV-2, is presently being prepared for flight, with 

an announced liftoff scheduled for March 4, 2011, again onboard 

an Atlas V 501 booster. Plan on seeing a regular stream of these 

launches, especially if the USAF pursues the program further.102

With this large number of new space flight vehicles under various stages 

of completion, one may well be successful in providing future human space 

access to low-Earth orbit and the International Space Station, as well as safe 

reentry and recovery from space.

101. Ibid.

102. John D. Gresham, “X-37B Orbital Test Vehicle,” DefenseMediaNetwork, February 28, 2011, 

available online at http://www.defensemedianetwork.com/stories/x-37b-orbital-test-vehicle/, 

accessed June 10, 2011.

277
This image from July 2008 shows the aeroshell for NASA’s Mars Science Laboratory. During the 

journey from Earth to Mars, the aeroshell encapsulates the rover, and it shields the rover from the 

intense heat of friction with the upper Martian atmosphere during the initial portion of descent. The 

aeroshell has two main parts: the backshell, which is on top in this image and during the descent, 

and the heat shield, which is on the bottom. The heat shield in this image is an engineering unit 

used for testing. The heat shield to be used in flight will be substituted later. The heat shield has a 

diameter of about 15 feet. By comparison, the heat shields for the Mars Exploration Rovers Spirit 

and Opportunity were 8.5 feet, and the heat shields for the Apollo capsules that protected astronauts returning to Earth from the Moon were just less than 13 feet. NASA PIA11430.

278
CHAPTER 7

The success of atmospheric reentry missions is constrained by the design of the 

thermal protection systems of the aerospace vehicles involved. During reentry 

into Earth’s atmosphere, space vehicles must operate in intense thermal-stress 

regimes that require an effective thermal protection system to ensure survival 

of the craft. Indeed, the thermal protection system may be the most important 

system used on a spacecraft that must enter an atmosphere and land. Next most 

significant, of course, is the landing system, whether it be a parachute, rocket, 

glide system, or Rogallo wing.

One recent study made the observation about the criticality of these systems 

as follows:

For vehicles traveling at hypersonic speeds in an atmospheric environment, TPS is a single-point-failure system. TPS is essential to 

shield the vehicle structure and payload from the high heating 

loads encountered during [reentry].... Minimizing the weight and 

cost of TPS, while insuring the integrity of the vehicle, is the 

continuing challenge for the TPS community.1

The modern thermal protection system originated in the decade after World 

War II as part of the ballistic missile program. Required to ensure the reentry 

of nuclear warheads, the thermal protection system represented a fundamental 

technology for these missile systems. The ablative thermal protection system 

protected the vehicle through a process that lifted the hot shock-layer gas away 

from the vehicle, dissipated heat absorbed by the ablative material as it burned 

away, and left a char layer that proved remarkably effective as an insulator that 

blocked radiated heat from the shock layer. 

 1. Ethiraj Venkatapathy, Christine E. Szalai, Bernard Laub, Helen H. Hwang, Joseph L. Conley, 

279
Coming Home

This ablative thermal protection system chronology demonstrates the evolution of the capability 

from the start of the space age to the present. NASA.

Throughout the 1960s and much of the 1970s, engineers involved in ablative thermal protection system development were quite active. The succession 

of programs, both human and robotic, kept them involved in making ever 

more capable thermal protection systems. This changed in the mid- to late 

1970s, however, as R&D and testing of ablative thermal protection system 

materials significantly declined for both the nuclear missile programs and the 

robotic probes that NASA pursued. At the same time, human programs that 

had used ablative systems were completed, and NASA moved toward a nonablative, reusable thermal protection system for the Space Shuttle program. This 

transition involved significant R&D, but by the early 1980s the systems had 

been resolved, and only intermittent upgrades would follow. Accordingly, the 

ablative thermal protection system community experienced a serious decline in 

capability. A return to an ablative thermal protection system for robotic entry 

probes in the last 20 years has required a reconstitution of the knowledge base 

lost in the 1970s. As a result, in the 1990s NASA invested in the development 

of two new lightweight ablators: phenolic-impregnated carbon ablator (PICA) 

and silicone-impregnated reusable ceramic ablator (SIRCA). It was PICA that 

enabled the Stardust Sample Return Mission. Additionally, SIRCA was used on 

the backshell of the Mars Pathfinder and the Mars Exploration Rover missions. 

280
Summary

These efforts reached a serious crescendo when NASA began developing the 

Orion crew capsule to replace the Space Shuttle. In essence, we have come full 

circle in the context of thermal protection systems, albeit with a spiral upward 

in capability and flexibility, and much of the resulting work is critically based 

on the efforts of earlier systems dating back to the Apollo era.2

A key realization coming from the cyclical process of ablative thermal protection system efforts was that despite the desire to use off-the-shelf materials, 

the knowledge base had atrophied so significantly that considerable effort had 

to be undertaken to reconstitute that knowledge base. Moreover, engineers 

working on thermal protection systems found that the knowledge gained in 

other programs was profound. For example, several engineers have concluded 

that without the investment in the 1990s in PICA thermal protection systems, 

the Mars Science Laboratory would have fallen even further behind schedule.3

In essence, the space thermal protection system community worked most effectively when it coordinated efforts to codevelop materials and technology that 

could be used in multiple projects and programs. As one study concluded: 

“The very important lesson learned here is that it is wise to have at least two 

viable candidate TPS materials in place for mission projects because although 

the selected TPS sufficed for a previous project, it may not be adequate for the 

next, even if the entry environments are only slightly more severe.”4

The process of moving from ablative heat shields in the 1960s to the Space 

Shuttle thermal protection system in the 1970s and to the various types of 

thermal protection systems currently available has represented something of a 

back-to-the-future approach to space flight. It raises important questions about 

the history of technology and its usually presumed progress. This history looks 

more circular than progressive when considering spacecraft return and recovery 

technology throughout the space age. It also illustrates the inexact nature of 

space technology and the lack of a clear line of development as the technology 

has looped back and forth, jumped ahead, and backtracked to earlier, proven 

 2. Bernie Laub, “Ablative Thermal Protection: An Overview,” presentation at the 55th Pacific Coast 

281
Coming Home

Current capabilities: TPS for sample return missions Density

Forebody 

Heat 

Shield

Supplier

Flight 

Qual or 

TRL

Potential Limit Entry velocity, 

km/s Other 

Potential 

Missions Heat flux, 

W/cm2

Pressure, 

atm <13 >13

FOREBODY HEAT SHIELD

Low-Mid

PICA FMI Stardust ~1200 <1 SR, CEV, 

Mars

Avcoat Textron Apollo ~1000 <1 Venus 

(aerocapture)

ACC LMA/C-Cat Genesis >2000 >1 MSR, CEV, 

Mars

BPA Boeing TRL 3–4 ~1000 ~1 Venus 

(aerocapture)

PhenCarb 

Family ARA TRL 5–6 (1000–4000) >1 MSR, CEV, 

Venus, Earth

High

3DQP Textron DOD 

(TRL4) ~5000 >1 SR, Venus

MSR, Venus, 

Jupiter, 

Saturn, 

Neptune

Heritage 

Carbon 

phenolic

Several 

capable, 

none active

Venus, 

Jupiter

(10,000–

30,000) >>1

 Fully Capable Potentially capable, qual needed Capable but heavy Not capable

NASA.

concepts.5 The history also suggests that very little is fully defined in space 

technology, especially as older concepts are resurrected and brought to the fore 

years after their abandonment.

Ultimately, this situation suggests one firm lesson: Thermal protection system 

technologies for many space missions are unique to NASA, challenging, and cross 

cutting. Developing these systems requires specialized efforts, but they end up 

offering multiple uses.

Moreover, since the thermal protection system forms the external surface 

of any space vehicle, it must be designed for all the environments experienced 

by the vehicle. As one researcher concluded:

 5. This is very effectively laid out in John Law, “Technology and Heterogeneous Engineering: The Case 

282
Summary

The primary function of the TPS is to protect the vehicle and 

its contents from aerodynamic heating, so it must be sized to 

keep internal temperatures within acceptable limits. In addition, the TPS must maintain its structural integrity so that it 

provides an acceptable aerodynamic surface during all portions 

of flight through the atmosphere. The TPS panels must therefore be designed to withstand aerodynamic pressure and drag 

loads, acoustic and dynamic loading from the engines at launch, 

dynamic pressures that can cause panel flutter, thermal expansion mismatches between the TPS and underlying structure, and 

strains induced by primary vehicle loads on the underlying structure. The TPS must also have an acceptable risk of failure after 

low-speed impacts during launch and hypervelocity impacts from 

orbital debris in space. Of course, all of these functions must be 

accomplished while eliminating any unnecessary mass.6

The requirements of these systems might evolve considerably over the various design efforts of the spacecraft, and the development of an adequate thermal protection system is always challenging and time consuming. Generic 

solutions do not seem to work, and only with specific applications for unique 

situations combined with rigorous testing in all of the envisioned elements of 

the spacecraft’s thermal protection system has the current generation of ablative 

heat shields emerged for use indefinitely into the future.

Some have suggested that development of a new thermal protection system 

material may only be accomplished through a careful balance between thermal 

performance and thermal structural integrity. “Regardless of whether the heat 

shield design is a tiled system (PICA), or a monolithic system (Avcoat), thermalstructural capabilities are critical. Detailed thermal response must be understood for the integrated system not just for acreage TPS material. Penetrations 

and closeouts require significant work and are difficult [to] manage...due to 

changing requirements.”7

 For all of the success enjoyed in reconstituting an 

ablative heat shield capability since the 1990s, many analysts warn that future 

neglect of thermal protection system development—as was the case in the 

1970s and 1980s—will again necessitate “an expensive, high risk, critical path 

approach to recover. Without the fortuitous timing of the CEV TPS ADP 

 6. M.L. Biosser, R.R. Chen, I.H. Schmidt, C.C. Poteet, and R.K. Bird, “Advanced Metallic Thermal 

283
Coming Home

PICA heat shield effort, MSL would have had no TPS options to meet their...

launch window.”8

 Indeed, researchers seem to agree that at present, “NASA/

US [is] short of efficient, robust TPS materials for future exploration missions: 

high mass Mars entry, outer planets, Venus, extra-Lunar Earth return.”9

There has been a similar circular trajectory in the history of spacecraft recovery systems. The first landing systems employed parachutes, followed by the 

aborted development of paraglider systems, the testing of lifting bodies, and the 

Space Shuttle’s winged design that allowed runway landings. Finally, spacecraft 

landing systems have returned to the use of parachutes. For the longest time, 

NASA explored ways to avoid landing its astronaut missions in the sea, which 

requires the U.S. Navy to perform overtaxing water rescues. Furthermore, the 

astronauts, who were all pilots, found the method objectionable. This was 

eloquently communicated in 1965 when artist Wen Painter, at NASA’s Flight 

Research Center, drew a powerful cartoon depicting two astronauts sitting in 

a Gemini capsule bobbing in the water while the Navy steams around in the 

distance, juxtaposed with a second image of a sleek lifting body landing on a 

runway with astronauts walking across the tarmac. The caption read: “Don’t 

be rescued at sea; fly back in style.” This cartoon captured magnificently what 

some saw as a key difference between space-capsule splashdowns at sea and 

spaceplane landings on a runway.

As soon as opportunity permitted, NASA moved to a winged spacecraft that 

allowed a fine degree of control over landing and ensured that splashdowns 

were a thing of the past. As one analyst wrote: “The ‘golden age’ had the astronauts flying in a spacecraft that was largely a closed system. The Shuttle, on 

the other hand, allows for a significant amount of control, but lacks the cachet 

and interest of its predecessor programs. The irony of the situation is that the 

Shuttle lands like an airplane in an era when test pilots are no longer the crème 

de la crème of astronaut candidates.”10 As the Space Shuttle was retired in 2011, 

NASA is wrestling anew with these landing systems. The simplest approach 

remains using parachutes deployed to slow the spacecraft and gently return 

it to Earth. With new aerodynamic systems capsules, landings might be well 

controlled and, whether landing on soil or water, should be quickly recovered. 

Orion, Dragon, and other possible future piloted spacecraft all feature this 

approach to landing. While considerable effort remains to be made on these 

 8. Ibid.

284
Summary

new reentry and recovery systems, this seems to be the direction of space policy 

for the foreseeable future.

In the context of the larger story of reentry and recovery from space, some 

five distinct lessons emerge that are worthy of discussion. These may be encapsulated as stated in a study of failures concerning the Space Shuttle, but they 

are applicable across a broad spectrum of activities:

1. Operating Experience: People and organizations need to learn valuable lessons from internal and external operating experience to avoid 

repeating mistakes and to improve operations.

2. Mission and External Influences: Budget and schedule pressures must 

not override safety considerations to prevent unsound program decisions.

3. Normalizing Deviations: Routine deviations from an established 

standard can desensitize awareness to prescribed operating requirements and allow a low-probability event to occur.

4. Technical Inquisitiveness: To ensure safety, managers need to 

encourage employees to freely communicate safety concerns and differing professional opinions.

5. Focus on Planning and Prevention: Safety efforts should focus 

more on planning and preventive actions rather than investigations 

and corrective actions resulting from accidents or events.11

In the end, the focus on thermal protection and landing systems cannot be 

deemphasized. As a recent study reported: “NASA’s ambitious exploration vision 

requires TPS innovations. Many future missions require TPS materials and/or 

concepts not currently available or, in some cases, new versions of old materials. 

New TPS materials, ground test facilities, and improved analysis models are 

required and will take some time to develop[.] Advances and improved TPS 

capabilities will benefit an array of missions (and enable some).”12

 11. 

285
Three canopies in the sunset, Apollo returns to Earth. NASA S69-36594

286
Roger D. Launius is senior curator in 

the Division of Space History at the 

Smithsonian Institution’s National Air 

and Space Museum in Washington, DC, 

where he was division chair from 2003 

to 2007. Between 1990 and 2002, he 

served as chief historian of the National 

Aeronautics and Space Administration. A 

graduate of Graceland College in Lamoni, 

IA, he received his Ph.D. from Louisiana 

State University, Baton Rouge, in 1982. 

He has written or edited more than 20 

books on aerospace history, including, 

most recently, Globalizing Polar Science: 

Reconsidering the International Polar and 

Geophysical Years (Palgrave Macmillan, 

2010); Smithsonian Atlas of Space Exploration (HarperCollins, 2009); Robots 

in Space: Technology, Evolution, and Interplanetary Travel (Johns Hopkins 

University Press, 2008); Space Stations: Base Camps to the Stars (Smithsonian 

Books, 2003; 2nd ed. 2009); Imagining Space: Achievements, Predictions, 

Possibilities, 1950–2050 (Chronicle Books, 2001); and others. He is a fellow 

of the American Association for the Advancement of Science, the International 

Academy of Astronautics, and the American Astronautical Society, and he is 

an associate fellow of the American Institute of Aeronautics and Astronautics. 

He also served as a consultant to the Columbia Accident Investigation Board 

in 2003 and presented the prestigious Harmon Memorial Lecture on the history of national security space policy at the United States Air Force Academy 

in 2006. He is frequently consulted by the electronic and print media for his 

views on space issues and has been a guest commentator on National Public 

Radio and all the major television network news programs.

287
Coming Home

Dennis R. Jenkins has worked as a NASA 

contractor for the past 32 years, mostly on the 

Space Shuttle program in a variety of engineering and management roles. After supporting the first few Space Shuttle launches at 

the Kennedy Space Center (KSC), he spent 5 

years activating the Vandenberg Launch Site 

in California before the facility was closed following the Challenger accident. Returning to 

KSC, he supported the recovery of the Shuttle 

accident and a variety of special projects. 

During the late 1990s, he was the ground systems lead for the X-33 program. Afterward, he 

managed a variety of upgrade projects at KSC. 

He spent 2003 on the staff of the Columbia 

Accident Investigation Board (CAIB), 2004 as staff to the President’s commission on the Future of Human Space Flight, and 2005 as the Verville Fellow 

at the National Air and Space Museum (NASM), after which, he returned to 

KSC. In 2010, he began participating in the Orbiters on Display Working 

Group that planned the delivery of the Space Shuttle orbiters to their display 

sites at the NASM; the Intrepid Sea, Air & Space Museum; the California 

Science Center; and the KSC Visitor Complex.

288
Numbers in bold indicate pages with illustrations.

A

A3-4/SV-5 lifting body, 135, 137, 140, 153

A3 lifting body, 133, 135

A-4 (V-2) missiles, 8–9, 10, 11–12, 16

Abbott, Ira H., 163, 164n95

ablative thermal protection systems

advantages of, 27

Apollo program, x, 47, 70–72, 74–79, 77, 

78, 257–58, 259–260, 278, 280, 282

applications for, ix

Army use of, 28–29

blunt-body reentry designs, viii, ix, 128

concept of, 27

coordination of development, 281

decline in development and capabilities, 280

development of, viii, ix, 280–84

elastomeric ablator, 64, 72, 191–92

film-return capsules from CORONA program, 

32

Gemini program, ix–x, 47, 62, 64–65, 67, 

69, 69n78, 72, 280

HL-10 lifting body shingle structures, 151

hot structures and, 191–92

lifting body/lifting reentry concepts, 147–48, 

154–55, 158–59, 159n86

M2 lifting body program, 147–48

Mars missions, 248–49, 250, 252, 278, 

280, 280, 281, 282, 283–84

materials for, 27, 29, 47, 54–55, 64–65, 72, 

140–41, 191, 200, 280

Mercury program, ix, 2, 47, 50–55, 53, 56, 

58–62, 61, 64–65, 70, 72

Orion CEV, x, 221, 252, 257–260, 262–63, 

281, 282, 283–84

PRIME vehicles, 140–41, 158

pyrolysis, char formation, and loss of ablator, 

27, 56, 58, 60, 61, 64, 67, 71, 76, 78, 

260, 262

refurbishment of, 160, 191–92

removable panels, 191, 192

return to use of, 280–84

safety record of use of, 2, 2n4

scientific mission sample return capsules, ix, 

245–46, 247–48, 250, 252–53

semiballistic capsules, 47

shock-layer gases and, 279

skip-gliding (dynamic soaring) and, 9

Space Shuttle program thermal protection, 

199

Stardust vehicle, 250, 252–53, 257, 258, 

280, 280, 282

testing of, 28–29

warheads (reentry vehicles) applications, 

28–31, 30, 52, 279

weight of, 65, 72, 74

X-15 aircraft, 158, 191–92, 196

See also specific types

accidents

during launch into space, 81

prevention of, 285

during reentry and return from space, vii, 81, 

209–10

acid-ammonia rocket facility (Langley), 29

acrylic film-tape, 216–17

289
Coming Home

active cooling concepts

BoMi skip-glider, 161

boost-gliders, 129

development of, 9, 195

hot structures and, 9

hypersonic flight and, 129, 130

HYWARDS program, 168

limitations of, 168, 195

National Aero-Space Plane, 224–25

types of, 27–28

for warheads (reentry vehicles), 28

water-wall (double-wall) cooling systems, 28, 

40, 71, 161, 165

weight of, 165, 224–25

Adam Opel automobiles, 5, 5n9

adhesives, room temperature vulcanizing (RTV), 

196, 206, 214, 215, 216

advanced flexible reusable surface insulation 

(AFRSI), 215, 216n56, 262

Advanced Research Projects Agency (ARPA), 

49–50

Aerobee sounding rocket, 94

aerodynamics

research and technological developments, 

vii–viii

shock layer, boundary layer, and 

aerodynamic heating, 6

smooth aerodynamic surfaces and reentry 

heating, 126, 169

thermal protection characteristics and, 126, 

169, 283

X-33 VentureStar program, 233–34

Aerospace Corporation, The, 132, 133, 135

Aerospace Test Pilot School, 155

Aerothermodynamic Elastic Structural Systems 

Environmental Tests (ASSET) vehicles. See

ASSET (Aerothermodynamic Elastic Structural 

Systems Environmental Tests) vehicles

aerothermodynamic structural vehicle (ASV), 

138–140, 138n31

aerothermoelastic vehicle (AEV), 138–39

AFFDL. See Air Force Flight Dynamics Laboratory 

(AFFDL)

Agena rocket stages, 91

ailerons, 143n43, 217

aircraft

design and shapes of, 16–17

jet-assisted technologies, 5

metal alloy construction, 22, 23n52

rocket-power-assisted technologies, 5

rocket-powered hypersonic aircraft, 9, 

21–22, 39

thrust, acceleration, and lift, 14

Air Force, U.S. (USAF)

ablative thermal protection system 

applications, 28–29, 52

Air Research and Development Command, 

49–50

Blue Gemini program, 67n75

boost-glider research, 163–64

CORONA program, 31, 32

Dyna-Soar program contract, 173, 173n119

FIRST program, 94

hypersonic flight research, 21

lifting body/lifting reentry programs, 132, 

137–142, 152–59

M-103 project, 135

Man In Space Soonest (MISS) program, 

48–50, 48n33, 49n37

Manned Glide Rocket Research System, 44

Manned Orbiting Laboratory (MOL) program, 

47, 67, 67n75

Mercury program role, 50

Mk series warheads, 19n40, 25, 25n58, 

25n61, 28, 29, 31

MX-2276 project (BoMi skip-glider), 

160–64, 162, 163n93, 165–66, 170

Rapid Capabilities Office, 273

reentry vehicles, 25, 25n61

Scientific Advisory Board Aircraft Panel, 21

290
Index

Space Maneuver Vehicle program, 273

Space Shuttle requirements, 186–87, 188, 

189, 198, 201

SR-126 (Rocket Bomber, RoBo), 164–66, 

170, 171

SR-131 (HYWARDS), 166–69, 166n102, 

170

X-17 research, 21

X-24A lifting body, 133, 153–55, 154

X-37/X-37B vehicle, 272–75, 274, 276, 277

Air Force Flight Dynamics Laboratory (AFFDL)

boost-glider concepts, 137–38, 202

establishment o, 133n21

L/D definition, 125n2

lifting body concepts, 132, 133, 135, 152, 

155–56, 157

Space Shuttle design, 188–89, 202

airframe insulation, 168–69

Aldrin, Edwin E. “Buzz,” Jr., 68, 69, 77, 268, 

268n86

Allegheny Ballistic Laboratory, 74

Allen, H. Julian

ablative thermal protection systems, 

development of, viii

active cooling concepts, 28

Ames, role at, 17

blunt-body reentry techniques, development 

of, viii–ix, 17–20, 18, 19n38, 33

boost-glider designs, 129

hypersonic flight vehicle designs, 127, 128

HYWARDS program, 167

name and nickname of, 17n36

Alpha Draco Model 122B boost-glider, 137

altitude

equilibrium altitude, 127–28

X-plane altitude records, 42, 42n19

alumina-enhanced thermal barrier (AETB-8), 

213, 216, 262

aluminosilicate and aluminosilicate-chromia, 

194–95

aluminum

BoMi skip-glider, 161

M2-F2 lifting body, 144

PRIME vehicle, 140

riser connectors, 79

Space Shuttle orbiter, 202, 207, 209

aluminum oxide, 71

American-Marietta Corporation, 173n120

Ames Research Center/Ames Aeronautical 

Laboratory

Allen role at, 17

Atmospheric Entry Simulator, 20

boost-glider research, 165

free flight tunnel research at, 20

full-scale 40-by-80 foot wind tunnel, 143, 

145, 148, 149, 153–54

hypersonic flight vehicle designs, 127

hypersonic tunnel research, 165

HYWARDS program, 166–69

lifting body concepts, 131, 132

M1 lifting body, 133–34

M2-F1 lifting body testing, 143

Orion CEV thermal protection development, 

257, 260

PICA development, 252, 257–59, 260

research and technological developments 

at, viii

skip-gliding and reentry heating research, 

9, 44

TUFROC development, 275

Unitary Plan Wind Tunnel, 18

X-24A (SV-5P) testing, 153–54

Antares rocket motors, 74

antipodal bomber/glider, ix, 7n15, 8n17, 81–82

Apollo program

AS-201 test, 74–75

AS-202 test, 75–76, 77

ablative heat shields, x, 47, 70–72, 74–79, 

77, 78, 257–58, 259–260, 278, 280, 282

Apollo 1 spacecraft, 75

291
Coming Home

Apollo 4 mission, 76, 77, 78

Apollo 5 mission, 76n96

Apollo 6 mission, 76–77

Apollo 8 mission, 78, 79

Apollo 9 mission, 121

Apollo 10 mission, 78

Apollo 11 mission, 77, 78, 78

Apollo 13 mission, 78–79, 123

Apollo 15 mission, 79, 120, 122

Apollo 16 mission, 119

approval of, 62, 62n66, 69

ballistic technologies and, 3

capsule design and development, x, 19n40, 

69–70, 125, 185

Command Module, 72, 75, 76–77, 76n96, 

77, 122–23, 123, 185

FIRE, Project, 72, 74

heat of reentry, 70, 76

landing accuracy, 125

L/D ratios, 125, 130, 170n111

length of reentry, 70, 76

lifting reentry vehicle shapes, 69, 71, 73

Lunar Module, 76, 76n96

Moon-landing missions, 77, 117, 185

number of flights, 81, 81n1

parachute reentry and recovery, x, 79, 83, 

113, 117–123, 119, 121, 123, 264, 286

parasail concepts, 101, 122

radome, 196

safety record of, 2, 2n4

semiballistic/semilifting capsules, 47, 70

subscale model testing, 74

success of, 117, 185

water landings and recovery, 47, 77, 78, 79, 

117–18, 119, 122–23, 123

Apt, Milburn G., 37

arc-jet facilities (Langley and Lewis), 29, 58

Ares I/Ares V crew launch vehicles, 255, 256, 

267, 268, 269

Armstrong, Johnny, 42n19, 156–57

Armstrong, Neil

Apollo 11 mission, 77

Dyna-Soar program, 178n132

human space flight programs, support for, 

268

Parasev flight testing, 101

side vision, landing with, 177–78

X-15 program role, 43

Army, U.S., 28–29, 240

Arnold Engineering Development Center, 165, 

202

ARPA (Advanced Research Projects Agency), 

49–50

ASSET (Aerothermodynamic Elastic Structural 

Systems Environmental Tests) vehicles

aerothermodynamic structural vehicle (ASV), 

138–140, 138n31

aerothermoelastic vehicle (AEV), 138–39

goals of program, x, 138

materials development and testing, 

139–140, 181

nose cap testing, 177

Space Shuttle design and, 43, 142, 201

START program, 140, 153

wings and reentry technique, 15

ASV (aerothermodynamic structural vehicle), 

138–140, 138n31

Atlantis, 215, 216

Atlas intercontinental ballistic missiles (ICBMs)

A/B/C missiles, 25n59

ablative thermal protection applications, 29, 

52

Apollo program model testing, 74

B/C missiles, 25, 28

Big Joe flight tests, 55–56, 57

blunt-body reentry designs, throw weight 

of, 132

CCDev program, 271, 272

design and shapes of, 16–17, 19, 19n40

development of, 16–17

292
Index

D missiles, 25n59, 28, 29

E/F missiles, 25n59, 28, 29

Mk series warheads, 19n40, 25, 28

piloted spacecraft launch vehicle, 45, 185

PRIME vehicle launch vehicle, 142

X-40 launch vehicle, 273

Atmospheric Entry Simulator (Ames), 20

Attainability of Celestial Bodies, The (Die 

Erreichbarkeit der Himmelskörper, Hohmann), 

14

Augustine, Norm, 254, 266

Aurora 7 Mercury capsule, 61

automobiles, rocket-powered, 5

Avcoat ablator, 71–72, 257–260, 262, 263, 

280, 282, 283

Avco Corporation/Avco Manufacturing 

Corporation

Apollo program heat shield development, 

71–72, 74

Mk series warheads, 29, 31

thermal protection systems for warheads, 

development of, 28, 29, 52

azimuth error, 218

B

B-52C aircraft, 182

B-70/XB-70/XB-70A bomber, 135, 152n66

ballistic coefficient of reentry vehicles, 20, 

20n41, 25, 26, 29, 31, 132

ballistic missiles and technologies

advantages of, 128

blunt-body theory and, 16–26, 19n40, 128

capsule designs for piloted spacecraft, 

45–46

development of ballistic missiles, 9–13, 10, 

13

Dyna-Soar program, 171

flight profile of, 127

g-loads during reentry, 35, 45, 130

heat of reentry and, viii, 70, 126, 128, 130

heat sink applications, 22–25, 26, 128, 230

intermediate-range missiles, 25, 25n59, 

28–29, 138n31, 230

limitations of, 128

materials for missiles, 12

reentry and recovery technologies and, viii, 

1–2, 3, 16–26, 19n40, 45–46, 51–52, 

70, 279

space flight and, 12

trajectory of ballistic reentry, 16, 130

See also Atlas intercontinental ballistic 

missiles (ICBMs); warheads (reentry 

vehicles)

ball nose for flow angle (attitude) measurements, 

42, 43

Becker, John V.

human space flight vehicles, 170

hypersonic flight research, 9, 21–22, 38, 

39, 43

HYWARDS program, 166

winged vehicle designs, support for, 44–45, 

132, 186

X-15 development, ix

X-15 program role, 39–40

Bell Aerosystems hydrogen peroxide engines, 

150–52

Bell Aircraft

boost-glider program, 140, 165

Dyna-Soar program, 171

fluid cooling systems, viii, 28

hypersonic bomber concepts, 164

skip-glider designs, 160

water-wall (double-wall) cooling systems, 28, 

40, 71, 161, 165

Bell Bomber-Missile (BoMi) skip-glider (MX-2276 

project), 160–64, 162, 163n93, 165–66, 170

Bell Brass Bell boost-glide concept, 160–64

Bell spaceworthy orbital Bomber-Missile 

(S-BoMi), 165–66

Bell X-1/XS-1 vehicle, 17, 37, 37n5, 144

293
Coming Home

Bendix Products Aerospace Division, 122, 178

beryllium

ballistic missile nose tips, 230

capsule afterbody shingle insulation, 58, 65, 

67, 69, 70

heat absorption testing, 24

heat sinks, viii, 23–24, 25, 35, 52, 54, 59

limitations and risks of use of, 52

NASP program, 225

PRIME vehicle, 140

B.F. Goodrich, 54

Big Joe flight tests, 54–56, 57, 85

Bikle, Paul F., 44, 143, 144

Billeter, Orion D., 146

Bisplinghoff, Raymond, 179

Blue Gemini program, 67n75

Blue Goose, 188

Blue Origin, xi, 270, 271–72

Blue Scout boosters, 138

blunt-body reentry designs

ablative thermal protection applications, viii, 

ix, 128

active cooling protection for, 28

Apollo program capsules, x, 19n40

ballistic coefficient of, 20, 20n41

ballistic missiles and technologies and, 

16–26, 19n40, 128

design and shapes of, 20–22

development of, viii–ix, 17–22, 19n38, 33

Gemini program capsules, ix–x, 19n40, 

62–64

g-loads during reentry, 35

heat of reentry and, viii–ix, 17, 19n40

human space flight vehicles, 170

lifting reentry concept, 62–64

Mercury program capsules, ix, 19n40, 170

Orion CEV capsules, x, 221

shadowgraphs of, 63

shock waves, reentry heat, and, 17, 19, 63

Space Shuttle program, x, 187–88

success of, ix

temperatures experienced by, 22, 35

testing of, 20–22

trajectory of reentry, 35

X-15 research airplane, 35

Boeing Company

boost-glider program, 140

CST-100 and CCDev program, xi, 270, 271

Delta IV EELV program, 180n139

Dyna-Soar program, 171, 173, 173n121, 

175, 177, 181

hypersonic bomber concepts, 164

Space Shuttle program, 198

thermal protection system development, 

257, 271

X-37/X-37B vehicle, 272, 274

Boeing Replacement Insulation (BRI)-18 tiles, 

216

Bomber-Missile (BoMi) skip-glider (MX-2276 

project), 160–64, 162, 163n93, 165–66, 170

boost-gliders

active cooling, 129

Brass Bell boost-glide concept, 160–64

concept and design of, 127–29

heat of reentry, 128–29, 163

HYWARDS (System 610A) program and 

vehicle (System 455L), 166–69, 166n102

launch vehicle, glider affects on, 144

McDonnell Alpha Draco Model 122B boostglider, 137

radiative cooling, 128–29

research and technological developments, 

163–64, 165

usefulness of design, 163–64, 165

See also Dyna-Soar (X-20) program

Borman, Frank, 67

Bossart, Karel, 16–17, 19n40

boundary layer

ablative thermal protection and, 27, 257, 

262

294
Index

heat of reentry and, 16

shock layer, aerodynamic heating, and, 6

smooth aerodynamic surfaces and, 169

Brass Bell boost-glide concept, 160–64

Bredt, Irene, 7, 8n18, 9, 127n6, 131n16, 160, 

161

Brown, Clinton E., 21

Brown, Harold, 179

budgets, 285

Bumper 8 rocket, 13

Bush, George W., 244, 254, 264–65, 275–76

C

C-47 (R4D/DC3) aircraft, 144

C-119 aircraft, 33, 101

C-130 aircraft, 83, 84, 86, 109

Cape Canaveral

Apollo program model testing, 74

Big Joe flight tests, 55

Bumper 8 rocket launch, 13

Gemini-B and MOL program, 67

Jupiter missile launch from, 28

orbital lifting body flights, 160

Stardust mission, 251

X-17 research, 21

capsules

Apollo program, x, 19n40, 69–70

Atlas launch vehicles, 45

attitudes toward and preference for winged 

vehicles, 92–94, 113–14, 124, 129, 130, 

185–86, 284

ballistic technologies and, 3, 45–46

blunt-body reentry designs, ix–x, 19n40, 

63, 170

Gemini program, ix–x, 19n40, 62–63

g-loads and human tolerances, 45

laminar flow and, 46

limitations of, 3

Mercury program, ix, 19n40, 50–62, 53, 

55, 170

midair recovery techniques, ix

nonlifting capsules, support for, 44–45

Orion CEV, x, 221

parachute reentry and landings, ix, 82–83

piloted spacecraft designs, 44–47, 47

reconnaissance program capsules, ix, 35

recovery of, ix (see also water landings and 

recovery)

return to use of, xi

safety record of, 2, 2n4

semiballistic/semilifting capsules, 46–47, 

70, 130

shadowgraphs of reentry designs, 63

single-use capsules, 3, 47, 67

size of, 45

Titan launch vehicles, 45

weight of, 45

carbon (graphite)

ablative thermal protection systems, 31, 191

Apollo program heat shields, 70

ASSET program and development of, 140

carbon-bonded carbon fiber (CBCF) 

insulating material, 248

carbon-fabric ablative material, 31

carbon fiber–reinforced resins, 141, 191, 

248, 249, 262, 280, 282

heat absorption testing, 24

heat shield applications, 195

heat sinks, 23–24

limitations of, 24

nanotubes, 260, 262

siliconized graphite, 177

See also phenolic-impregnated 

carbonaceous ablator (PICA)

carbon-carbon

applications for, 282

ASSET program and development of, 140

coated carbon-carbon composites, 224

effectiveness of, 159n86

Genesis capsule, 248–49, 280, 282

295
Coming Home

heat shield applications, 194

nose caps, 173n119, 199, 208

Orion CEV thermal protection system, 262

Space Shuttle reinforced carbon-carbon 

(RCC), x, 155, 195, 199, 200, 201–2, 203, 

208–10, 210

caret-wing wave rider concepts, 152, 152n67

Carpenter, Scott, 60, 61

CBS600 alloy, 193

CCDev (Commercial Crew Development) 

program, xi, 270–72, 276–77

Central Intelligence Agency, 31, 32n74

Cerachrome (Dyna-Flex insulation), 165, 

176–77, 177n128, 209

ceramic thermal protection systems

Apollo program, 70, 71

blunt-body reentry design heat shields, ix

characteristics of, 168

foamed materials, 71

HL-10 lifting body, 151

HYWARDS program, 168

Mercury capsule paint, 58

reusable surface insulation (RSI), 195–96, 

198–200, 200n20, 201–2, 208, 210–17

Space Shuttle thermal protection system, x, 

2, 155, 159, 195

Cernan, Gene, 268

Challenger, 81, 180n139, 205

Chamberlin, James A., 99n31, 105, 107

Chance Vought and Dyna-Soar program, 171, 

173, 173n119, 177

China

number of space flights, 81, 81n1

parachute reentry and recovery, 82

chromium and chromium alloys

Inconel X, 23n52, 41n16

thoria dispersed nickel-chromium (TD NiCr), 

193

Clarke, Arthur C., 11n23, 185

cobalt and cobalt alloys

heat shield applications, 71, 193, 194

heat sink applications, 23

L-605 alloy, 140, 168

Collier’s space travel articles, 36, 96

Collins, Michael, 77, 102n35

Columbia

accident and loss of crew, vii, 81, 209–10

cause of accident, vii

debris from, vii

debris strikes, 209–10

flight testing before maiden voyage, 185

thermal protection system, 204, 215, 

216n56

thermal protection system, damage to, vii, x, 

207, 209–10, 219

tile installation, 205–7

Columbia Accident Investigation Board (CAIB), vii

columbium

AEV experiments, 139

ASSET program and development of, 140

D-36 columbium, 176, 177

Dyna-Soar construction, 176–77, 181

F-48 columbium alloy, 71

heat shield applications, 193, 194

HL-10 lifting body shingle structures, 151

Inconel X, 23n52, 41n16

silicide coating, 177, 181

Commercial Crew Development (CCDev) 

program, xi, 270–72, 276–77

commercial spacecraft, 221, 266, 268

Comparative Analysis of the Performance of 

Long-Range Hypervelocity Vehicles, A (Allen, 

Eggers, and Neice), 127

composite materials, 140. See also specific 

composites

conical piloted reentry vehicle with rotors, 

182–83

Constellation program

Ares I/Ares V crew launch vehicles, 255, 

256, 267, 268, 269

296
Index

cancellation of, 258, 264–65

nonreusable, water recovery concepts for, 3

Orion CEV, 275–76

Orion Crew Exploration Vehicle (CEV), x, 221, 

252, 255–264, 258, 261

spacecraft development and design, 186, 

222, 255

thermal protection system, 221

control surfaces

aeroelastic panel flutter, 139

AEV experiments, 139

Dyna-Soar program, 176

X-24A (SV-5P) lifting body, 153

Convair

Atlas ICBM development, 16–17

Dyna-Soar program, 171

hypersonic bomber concepts, 164

ILRV design, 186

Space Shuttle program, 190

warhead design and shape testing, 17

Convair Pye Wacket missile, 135

Cooper, Gordon, 62

copper

heat absorption testing, 24

heat sinks, viii, 22, 23–24, 35

Mk I/Mk II warhead heat sinks, 25

Copper Canyon program, 222. See also National 

Aero-Space Plane (NASP, X-30)

CORONA satellite reconnaissance program, ix, 

31–33, 33, 35, 82, 135

corrosion-resistant steels (CRES), 41

Cosenza, Charles, 188

Coupled Mode landing devices, 114

coupled roll-spiral mode, 147, 147n55

Crawley, Edward, 267

Crew Exploration Vehicle (CEV). See Orion Crew 

Exploration Vehicle (CEV)

Crew Recovery Vehicle. See X-38 Crew Return 

Vehicle (CRV)

Crews, Albert H., 178n132

Crossfield, A. Scott, 37, 39–40, 40n12

cross-range capabilities

Apollo program, 170n111

DC-3 (MSC-001) concept, 188

Dyna-Soar program, 178–79

film-return capsules and vehicles, 135

heat of reentry and, 207–8

L/D ratios and, 46, 170n111

lift and, 46

PRIME vehicle, 140, 142

runway landing at launch site, 132, 187, 198

Space Shuttle program, 132, 186–87, 188, 

189, 190, 198, 201

D

D-558-2 Skyrocket, 37

Dana, William H., 147, 149–150, 158, 178n132

DARPA (Defense Advanced Research Projects 

Agency), 222–23, 273

Day, LeRoy, 204

DC-3 (MSC-001) concept, 184, 187–88

DC-3 (R4D/C-47) aircraft, 144

DC-325 ablative material, 64, 280

DC-X powered landing concept, x, 221, 229n18

Decoupled Mode landing devices, 114

Defense, U.S. Department of (DOD)

CORONA program, ix, 31–33, 33, 35, 82

Dyna-Soar program, 173, 179–181, 186

human space flight program, role in, 50

Man In Space Soonest program, 49–50

parachute reentry and recovery, ix

SCORE project, ix, 82

Space Shuttle program, approval of, 198

thermal protection system development, 232

TRLs, 282

See also ASSET (Aerothermodynamic 

Elastic Structural Systems Environmental 

Tests) vehicles; PRIME (Precision Reentry 

Including Maneuvering reEntry) vehicles

297
Coming Home

Defense Advanced Research Projects Agency 

(DARPA), 222–23, 273

Delta IV EELV program, 180n139

Delta rockets, 138n31, 251

delta-wing concepts

advantages of, 126

double-delta design, 199

hypersonic flight and, 189

HYWARDS program, 166

lifting body/lifting reentry concepts, 130

runway/land landings and recovery, x

Space Shuttle program, x, 188–89, 190, 

198, 199–200

thermal protection systems for, 126

Dennis, David H., 134

deviations, normalization of, 285

Discoverer program, 31–32. See also CORONA 

satellite Reconnaissance program

Discovery, 215

Disney, Walt, 34

Dornberger, Walter R., viii, 12n25, 160, 165

double-wall cooling system. See water-wall 

(double-wall) cooling systems

Douglas Aircraft

Apollo program model heat shroud, 74

Dyna-Soar program, 171

hypersonic bomber concepts, 164

Project RAND, 14–15

Douglas D-558-2 Skyrocket, 37

Douglas F5D Skylancer, 177–78

Douglas R4D (C-47/DC3) aircraft, 144

Douglas X-2 aircraft, 22, 37

Dow Corning 3140 paint, 217

Dow Corning DC-325 ablative material, 64, 280

Dow Corning Z-6070 methyltrimethoxysilane, 

205

Dragon capsule spacecraft, xi, 270–71, 277, 

284–85

Draper, Alfred C., Jr.

delta-wing Space Shuttle design, 188–89

L/D definition, 125n2

lifting body/lifting reentry concepts, 155, 156

lifting body/lifting reentry programs, 138, 

152

Dream Chaser lifting body, xi, 272

Dryden, Hugh L., 48, 49

Dryden Flight Research Center/Flight Research 

Center

Gemini program reentry and landing 

concepts, ix–x, 82

HL-10 testing, 144, 148–49, 153

HSTV testing, 104–7, 106

lifting body programs, x, 139

M2 lifting body testing, 143, 145, 153

orbital lifting bodies, 160

Parasev flight testing, 101–3, 103

X-24A (SV-5P) testing, 153, 154, 238, 240

X-38 Crew Return Vehicle (CRV), 236–37

X-40 vehicle testing, 273

dry landings. See runway/land landings and 

recovery

Duke, Charles M., Jr., 119

DuPont, Tony, 222

DuPont Aerospace, 222

Dupont Ludox, 205

Dyna-Flex insulation (Cerachrome), 165, 

176–77, 177n128, 209

dynamic soaring (skip-gliding), 7–8, 9, 44, 127, 

128, 129–130, 163–64

Dyna-Soar (X-20) program

approval of, 62n66

booster, interface with, 182, 182n44

boost-gliders and design of, 138

cancellation of program, 140, 172, 

179–181, 182, 186

construction materials, 176–77, 178, 178, 

181

contractors, 171, 173, 173n119, 177

cross-range capabilities, 178–79

298
Index

design and characteristics, 138, 172, 173, 

175–78, 178, 180, 189

development and construction of, 44, 171, 

173, 175, 181, 185

engines for, 171

funding for, 170, 179, 182

g-loads and human tolerances, 35

goals of program, x, 180–81

graphite (carbon) construction, 24

guidance and control, 171

heat shield, refractory metal, x, 70, 71, 

176–77, 178, 186, 192, 195

hypersonic boost-glide concept (Step I), 

170–71

insulation, 40, 176–77, 177n128, 180

intelligence-gathering, limited-bombardment 

vehicle (Step II), 171

launch vehicle, glider affects on, 144

L/D range, 176, 176n127, 178–79

legend and myths, 171, 173

length of reentry, 70

lifting reentry concept, x, 70

naming of, 170, 170n110

nose caps, 173n119, 177, 181

pilots, 178n132

radiative cooling, 129n12, 176, 179, 192

reconnaissance and bombardment vehicle 

(Step III), 171

reusability of vehicles, 70, 177, 178, 186

runway landings, x, 178–79, 181–82

schedule for, 173, 182

Space Shuttle design and, 182, 186, 201

System 620A designation, 173n120

technical problems with, 181–82

temperature-sensitive paint for testing, 175

testing of, 175

Titan launch vehicles, 173, 174, 180n139, 

182

water-wall (double-wall) insulating technique, 

40

windows, 177–78

wing design, 175

wings and reentry technique, 15

E

Edwards, George G.

M1 lifting body, 133–34

tail designs, 134

Edwards Air Force Base (AFB)

dry lakebed landings, 63, 94, 108, 182, 239

orbital lifting body flights, 160

rocket-powered research at, 158, 158n83

X-24A (SV-5P) testing, 153, 154, 155

X-24B (SV-5P-2) lifting body testing, 156–58

X-24B (SV-5P-2) testing, 156–58

X-38 testing, 239

X-40 vehicle testing, 273

See also Dryden Flight Research Center/

Flight Research Center

Eggers, Alfred J., Jr.

ablative thermal protection systems, 

development of, viii

active cooling concepts, 28

Ames, role at, 17

blunt-body reentry techniques, development 

of, viii–ix, 17, 19–20, 19n38, 33

boost-glider designs, 129

hypersonic flight vehicle designs, 127

HYWARDS program, 167

lifting reentry vehicles, preference for, 130

M1 lifting body, 133–34

M2 lifting body program, 143

skip-glider designs, 130

Eglin Air Force Base (AFB), 160

Ehricke, Krafft A., 160

Eisenhower, Dwight D., 31, 50, 169–170

elastomeric ablator, 64, 72, 191–92

elastomeric reusable surface insulation (ERSI), 

202, 204

El Kabong test program capsule, 101

299
Coming Home

Endeavour, 215, 216n56

Engineering Research Consultants Inc., 260

engines

for Dyna-Soar program, 171

hydrogen peroxide engines, 150–52

lifting body/lifting reentry concepts, 164

scramjet engines, 222, 234

spaceplane engines, 222

Space Shuttle landing engines, 156, 198

X-33 VentureStar program, 233, 234

See also rocket engines

Enos the Chimp, 60

entry, 8n16

environments, natural and induced, 190, 

282–83

epoxy-based composites, 31, 191, 249, 258

equilibrium altitude, 127–28

equilibrium-glide flight, 218

ESA-3560/ESA-3560HF/ESA-5500M3 ablators, 

140–41, 200

ethylene-air jet facility (Langley), 29

European rocket enthusiast organizations, 9

European Space Agency (ESA), 241–42, 243–44

expendable launch vehicles (ELVs), 221–22

F

F5D Skylancer, 177–78

Fabrication of Inflatable Reentry Structures for 

Test (FIRST) program, 94

Faget, Maxime A.

Apollo program capsules, 70

blunt-body reentry designs, 132, 170, 

187–88

DC-3 concept, 184, 187–88

Mercury program capsules, 50–51, 70

nonlifting capsule designs, 45, 46

Polaris Task Group role, 29

Space Shuttle design, 187–89, 198

warhead research, 29

Falcon 9 heavy lifter, xi, 270

faster-better-cheaper strategy, 228–29

felt reusable surface insulation (FRSI), 214–15, 

235

Feltz, Charles H., 39–40, 42

fiberglass, 29, 54–55, 64, 65, 72

Fiber Materials, Inc., 252

fibrous refractory composite insulation (FRCI), 

211–12, 216

film cooling systems, 27–28

film-return capsules and vehicles, ix, 32, 33, 82, 

132, 135, 152

FIRE (Flight Investigation of the Reentry 

Environment), Project, 72, 74

FIRST (Fabrication of Inflatable Reentry 

Structures for Test) program, 94

flat-bottomed vehicles

heat transfer and, 130

human space flight vehicles, 170

HYWARDS program, 166

lifting body/lifting reentry concepts, 130, 

131–32, 134

Flax, Alexander H., 179

Fletcher, James, 200, 200n23, 201n24

flexible reusable surface insulation (FRSI), 262

Flight Investigation of the Reentry Environment 

(FIRE), Project, 72, 74

Flight Research Center. See Dryden Flight 

Research Center/Flight Research Center

fluid cooling systems. See liquid cooling systems

flying, advantages of over falling, 116

flying bathtub, 145

full-scale 40-by-80 foot wind tunnel (Ames), 

143, 145, 148, 149, 153–54

Funk, Ben I., 107

G

Galileo program, 246, 247–48, 280, 282

Gardner, Trevor, 164

Garver, Lori B., v, 265

Gehman, Harold W., Jr., vii

300
Index

Gemini program

ablative heat shields, ix–x, 47, 62, 64–65, 

67, 69, 69n78, 72, 280

approval of, 62

ballistic technologies and, 3

capsule afterbody shingle insulation, 65, 

67, 69

capsule design and development, ix–x, 

19n40, 62–64, 125, 185

Gemini II mission, 47, 65, 67

Gemini III mission, 66, 67

Gemini VII mission, 67, 69

Gemini XII mission, 68, 69

Gemini-B capsule, 47, 67, 67n75, 180

guidance and control systems, 62–64

landing accuracy, 62–64, 125

L/D ratios, 64, 125, 130

lift and maneuvering capability during 

reentry, 62–64, 65

Mark II program, 62

mission requirements and research goals, 

62, 90–91

number of flights, 81, 81n1

parachute reentry and recovery, x, 68, 82, 

91–92, 91, 99, 109, 111, 113

paraglider landing system, 94–114, 97, 100, 

110, 116, 238n37

parasail concepts, 99–101, 111–12, 115, 

122

Parasev (Paraglider Research Vehicle), 94, 

101–4, 103

Rogallo wing paraglider recovery system, 

ix–x, 62–63, 82, 91, 92–94, 93, 95, 111

safety record of, 2, 2n4

semiballistic capsules, 47

Test Tow Vehicle, 94–95

Titan launch vehicles, 66, 91, 173n120

water landings and recovery, ix–x, 47, 

63–64, 66, 82, 91, 92

General Dynamics. See Convair

General Electric

Apollo program vehicle, 73

Dyna-Soar program, 171

Mercury capsule heat shields, 54, 58–59, 

60, 72

Mk 12 warhead, 31, 31n73

Mk III warheads, 29

Mk I/Mk II warhead heat sinks, 25, 28, 52

Power Systems Composites, 274

thermal protection systems for warheads, 

development of, 28, 29, 52

Genesis program, xi, 221, 244–250, 245, 247, 

280, 282

Gentry, Jerauld R., 149, 155

German Society for Space Travel (Verein für 

Raumschiffahrt, VfR), 9, 11, 14, 36

Germany

rocket engine research and designs, 5, 9, 11

rocket enthusiast organizations, 9, 11, 14, 

36

V-2 missiles, 8–9, 10, 11–12, 16

Gilruth, Robert R., 21, 29, 62, 102n35, 107

Glenn, John, 60

Glenn L. Martin Company

boost-glider program, 140

ceramic thermal protection systems, 71

Dyna-Soar program, 171

hypersonic bomber concepts, 164

merger to form Martin Marietta Corporation, 

173n120

PRIME vehicle heat shields, 140–41

Titan launch vehicles, 173, 173n120, 174

See also Lockheed Martin; Martin Marietta 

Corporation

Glenn Research Center/Lewis Research Center

arc-jet facility, 29

Mercury boilerplate capsule manufacturing, 

54

Orion CEV thermal protection development, 

260

301
Coming Home

research and technological developments 

at, viii

gliders, 164. See also boost-gliders; lifting body/

lifting reentry concepts; skip-gliding (dynamic 

soaring)

g-loads

ballistic reentry and, 35, 45, 130

blunt-body reentry and, 35

contour couches for astronauts, 85–86

Dyna-Soar vehicle, 35

human tolerances, 35

L/D ratios and, 126, 131

Mercury program, 51, 56

semiballistic reentries and, 47

X-15 research airplane, 35

Goddard, Robert H., xii, 2, 4–5, 6

Goddard Research Center, 259

gold, 22, 58, 65

Goldin, Daniel S., 243–44

Goodyear Aircraft Corporation, 85, 178

Gordon, Henry C., 178n132

graphite. See carbon (graphite)

gravity

energy needed to escape, 1

speed needed to escape, 3, 14

Griffin, Michael D., 265

Grissom, Virgil I. “Gus,” 52, 67, 101

Grumman, 198–99, 200

Guadalcanal, 121

guidance and control

Apollo program model, 74

coupled roll-spiral mode, 147, 147n55

Dyna-Soar program, 171

Gemini program, 62–64

hypersonic gliders, 137–38

M2-F2 pilot-induced oscillations, 147

research and technological developments, 

vii–viii

Space Shuttle program, 217–19

X-33 VentureStar program, 233, 234

X-38 program, 241

Guillaume, Robert, 236

H

HAC (heading-alignment circles), 218–19

Haise, Fred W., Jr., 123

hang gliders and hang gliding, 93, 95–96

Hankey, Wilbur L., 1n2, 132

Hanley, Jeff, 267

hardened compacted fibers (HCF) materials, 

194–95, 200, 200n20

Hastelloy, 193

Haynes 25 alloys, 71

Haynes 188 alloy, 193

HCF (hardened compacted fibers) materials, 

194–95, 200, 200n20

heading-alignment circles (HAC), 218–19

heat of reentry

Apollo program, 70, 76

ballistic coefficient and, 20, 20n41

ballistic technologies and, viii, 70, 126, 128, 

130

blunt-body reentry designs, shock waves, 

and, 17, 19, 63

blunt-body reentry designs and, viii–ix, 17, 

19n40

boost-gliders, 128–29, 163

boundary layer heat and, 16

cross-range capabilities and, 207–8

direct reentry, 9

early approaches to dissipation of, viii–ix

hot jet research facility research on, 47

hot structures and, 165

hypersonic flight and, 21–22, 38

HYWARDS program, 166–69

kinetic energy and, 16, 16n34

leeward-side heating, 202, 208

length of reentry and, 47, 70, 76, 126

lifting body/lifting reentry concepts, 70, 126, 

130, 145

302
Index

Mercury program, 51–52, 56, 70, 130

research and writings about, 14–15

return speed and, 1

semiballistic reentries and, 47

shadowgraphs of reentry designs, 63

skip-gliding (dynamic soaring) and, 7–8, 9, 

129–130, 163

slender, streamlined designs and, 16–17, 

19n40, 20, 31, 63

Space Shuttle orbiter, 207–8, 217–19

superheating, viii, 224

temperature-sensitive paint, 175

temperatures of, 16n34, 22

X-15 aircraft, 40

X-plane research goals, 39

See also thermal protection systems (TPS)

heat pipes, 27

heat sinks

ballistic missile applications, 22–25, 26, 

128, 230

concept of, viii, 26

heat absorption testing, 24, 24n53

human space flight and, 35

limitations of, viii, 35

materials for, viii, 22–25, 22nn48–49, 35

skip-gliding (dynamic soaring) and, 9

warheads (reentry vehicles) applications, 25, 

26, 28, 35, 52, 59

heat transfer

flat-bottomed shapes and, 130

hypersonic flight and, 6, 8

shock layer and, 6, 279

Hello, Bastian “Buzz,” 153, 153n70

Hetzel, Charles “E.P.,” 109

High Speed Tunnel (Langley), 145, 149

high-temperature reusable surface insulation 

(HRSI), 202–4, 203, 211–12, 213, 216, 

216n56

High Temperature Tunnel (Langley), 72

HL-10 lifting body

characteristics and design of, 133, 148–49

development of, 133, 134–35

horizontal ladder designation, 135

hydrogen peroxide engines, 150–51

NB-52 carrier aircraft launch, 149

orbital lifting body based on, 159–160

Space Shuttle design and, 150, 151–52

supersonic flight in, 149–150

testing of, 144, 148–152, 153

thermal protection system, 150, 151

Hoag, Peter C., 149, 151

Hoey, Bob, 42n19, 46n30, 114, 116n66, 

134n23, 157n81, 159n82

Hohmann, Walter, viii, 14–15, 14n28

Hohmann Transfer Orbit, 14n28

Holmes, Brainerd, 93

hot jet research facility, 47

hot structures

ablative thermal protection systems and, 

191–92

BoMi skip-glider, 161

concept of, 9, 26

development of, 9

heat of reentry and, 165

HYWARDS program, 168

materials for, 24n53, 195–96

radiative cooling and, 9, 26

titanium, 161

X-15 aircraft, 23n52, 24n53, 26, 40–42, 

168, 191–92, 195–96

human space flight

ablative thermal protection applications, ix

accidents and loss of crew, vii, 81, 209–10

beginning of program, 33

Collier’s space travel articles, 36, 96

commercial piloted spacecraft, 221, 266

crew escape system, backup, 97, 98

design and shapes of vehicles for, 170

exploration beyond low-Earth orbit, 266

funding for, 266

303
Coming Home

Gilruth role in, 21, 29

g-load tolerances, 35

heat sink technology and, 35

launch vehicles, 185–86

lifting reentry as more suitable than ballistic 

reentry, 130

Mercury program contributions of to piloted 

space flight, 44

number of flights, 81, 81n1

programs and initiatives, review of, 264–68, 

270

reentry requirements for humans, 1–2

simplicity of, 35–36

spaceplanes and early visions for, 2–6, 15, 

81–82, 185–86, 222

success of U.S. programs, 185

warhead researchers and role in program 

for, 29

X-15 program contributions to, 42, 43–44, 

43

X-plane research goals and, 39, 170

See also space flight

Huxman, Joseph D., 146

hydrogen peroxide engines, 150–52

hypersonic bombers, 164

hypersonic flight

analysis of hypersonic airflow, 6

blunt-body reentry design research, 20–22

characteristics of, 6

conference on, 137

delta-wing concepts and, 189

design and shapes of vehicles for, 16–17, 

127–130

equilibrium altitude and, 127–28

heat of reentry and, 21–22, 38

heat transfer and, 6, 8

intercontinental missile development and, 

38–39

L/D research, 152

lifting body/lifting reentry concepts, 137

pressure generation during, 6, 8

propulsion for, 38, 39

research and technological developments, 

8–9, 38–39

rocket-powered hypersonic aircraft, 9, 39

shock waves, shock layer, and, 6, 41, 43

spaceplanes and, 5–6

V-2/WAC Corporal rocket vehicle testing, 12

wind tunnel research at Langley, 21, 38, 39

X-15 program research, 43–44

Hypersonic Glide Rocket Weapons System, 170. 

See also Dyna-Soar (X-20) program

hypersonic gliders, 137–38, 176n127

hypersonic lift-to-drag (L/D) coefficient, 125–26, 

125n2, 127–28

hypersonic tunnel (Ames), 165

Hypersonic Weapons and Research and 

Development Supporting Systems (HYWARDS, 

System 610A) program and vehicle (System 

455L), 166–69, 166n102, 170

HYWARDS (Hypersonic Weapons and Research 

and Development Supporting Systems, System 

610A) program and vehicle (System 455L), 

166–69, 166n102, 170

I

ICBMs. See Atlas intercontinental ballistic 

missiles (ICBMs); Titan intercontinental ballistic 

missiles (ICBMs)

impact theory, 131

Inconel alloys

heat shield applications, 193

Inconel 601–covered bulk insulation, 209

Inconel 718 riser connectors, 79

Inconel 750 wire mesh, 216

Inconel X characteristics, 23n52, 41n16

Inconel X heat absorption testing, 24, 24n53

Inconel X heat sink applications, 23–24

Mercury capsule afterbody application, 54

304
Index

X-15 Inconel X hot structures, 22, 23n53, 

24n53, 26, 41–42, 168, 230

Incoflex, 209

inertial measurement unit, 42

Inflatable Micrometeoroid Paraglider (IMP) 

project, 94

inquisitiveness, technical, 285

insulated structures, 26, 40, 40n15

insulation

airframe insulation, 168–69

Dyna-Soar program, 40, 176–77, 177n128, 

180

function of, 195

insulation types

ceramic-fiber batt, 26

Dyna-Flex (Cerachrome), 165, 176–77, 

177n128, 209

foam, vii

Micro-Quartz (Q-Fiber Felt), 177, 177n128

Min-K, 65

Thermoflex/Thermoflex RF, 58, 64, 65, 67, 

169

See also reusable surface insulation (RSI); 

shingle insulation

Integral Launch and Reentry Vehicle (ILRV), 

186–87

intercontinental missiles

Dyna-Soar guidance system, 171

hypersonic flight research and development 

of, 38–39

Soviet ICBM program, 170

types of, 171n114

See also Atlas intercontinental ballistic 

missiles (ICBMs); Titan intercontinental 

ballistic missiles (ICBMs)

intermediate-range missiles, 25, 25n59, 28–29, 

138n31, 230

International Geophysical Year, 169

International Space Station (ISS)

CCDev program, 271

completion of, 254

Crew Recovery Vehicle (X-38), 236–244, 

237, 239, 241

felt reusable surface insulation (FRSI), 215

funding for, 236, 243–44

Orion CEV for crew transportation, 257, 264

Soyuz spacecraft for crew transportation, 

244

support for program, 266, 268

vehicles for access to, 277

Intrepid, 66

Iwo Jima, 123

J

JC-130B Hercules, 141, 142

John F. Kennedy Space Center. See Kennedy 

Space Center (KSC)

Johnson, Lyndon Baines, 32–33

Johnson Space Center/Manned Spacecraft 

Center

DC-3 concept, 184

Orion CEV thermal protection development, 

260

parasail concepts, 77, 122

Space Shuttle program, 186

X-38 Crew Return Vehicle (CRV), 236–37

Josephs, Lyman, 153

Jupiter missiles, 13, 28–29, 52, 54, 230

Jupiter missions, 247–48, 282. See also Galileo 

program

K

Kapton acrylic film, 216, 217

Kennedy, John F., 69, 186

Kennedy Space Center (KSC)

Ares I-X vehicle, 269

Columbia tile thermal protection tile 

installation, 206–7

Space Shuttle launch site, 198

Space Shuttle program, vii, 199

305
Coming Home

Kenyon, George C., 133–34

Killian, James R., 49

King, Jay L., 146

Knacke, Theodor W., 84–85, 118–120

Knight, William J. “Pete,” 42, 178n132

Korycinski, Peter F., 166

Kowal, John, 259

Kubrick, Stanley, 185

L

L605 alloy, 193

Lake Champlain, 65

laminar flow, 46

land, landings on. See runway/land landings and 

recovery

landing systems

challenges related to, vii–viii, xi

Coupled Mode landing devices, 114

DC-X powered landing concept, x

Decoupled Mode landing devices, 114

importance of, 279

research and technological developments, 

vii–viii, ix, xi, 284–85

targeted and pinpoint landings, 62–64, 93, 

113–14, 116, 116n66, 125

See also parachute reentry and recovery; 

paraglider/parawing landing systems; 

runway/land landings and recovery; water 

landings and recovery

Langley Research Center/Langley Memorial 

Aeronautical Laboratory

acid-ammonia rocket facility, 29

Apollo capsule heat shield requirements, 

71–72

arc-jet facility, 29, 58

boost-glider research, 165

ethylene-air jet facility, 29

FIRE, Project, 72, 74

High Speed Tunnel, 145, 149

High Temperature Tunnel testing, 72

HL-10 lifting body (see HL-10 lifting body)

Hypersonic Tunnel testing, 21, 38, 39

HYWARDS program, 166–69

lenticular vehicle, 133, 135

Little Joe capsules, 89

Mercury boilerplate capsule manufacturing, 

54

Mercury capsule heat shield specifications, 

58–59

Mercury capsule heat shield testing, 56, 58

Mercury landing and recovery system 

testing, 83–84

orbital lifting bodies, 160

Orion CEV thermal protection development, 

260

paraglider landing system development, 82, 

91, 95–96, 108

research and technological developments 

at, viii, ix

Supersonic Pressure Tunnel testing, 41

Thermal Structures Tunnel testing, 72

Unitary Plan Wind Tunnel testing, 72

wind tunnel testing at, 37, 72

X-37/X-37B vehicle development, 274

launch into space

accidents during, 81

force and power for, 1, 13–14

speed needed to escape gravity, 3, 14

See also space flight

launch vehicle, glider affects on, 144, 182, 

182n143

L/D ratios. See lift-to-drag (L/D) ratios

Lea, R.H., 159

lenticular vehicle, 133, 135, 139, 143

Lewis Research Center. See Glenn Research 

Center/Lewis Research Center

Ley, Willy, 7n14, 36

LI-900 tile insulation, 196, 204, 210, 216, 

216n56, 262

LI-1500 tile insulation, 196, 200

306
Index

LI-2200 tile insulation, 196, 202–4, 203, 210, 

216n56, 262

LI-3000 tile insulation, 204

lift

capsule lift and maneuvering capability 

during reentry, 62–64, 65, 130

capsules and semiballistic reentry, 46–47

concept of, 46

undersurface designs to increase, 152, 

152n55

lifting body/lifting reentry concepts

ablative heat shields, 147–48, 154–55

ablative thermal protection systems, 

158–59, 159n86

advantages of, 69, 125–27, 131–33

Air Force and Air Force–sponsored research, 

132, 137–142, 152–59

Apollo program vehicles, 69, 71, 73

blunt-body lifting reentry, 62–64

conference on, 137

designs, research, and technological 

developments for, x, 130, 131–37, 136

(see also specific lifting bodies)

Dyna-Soar program, x, 70

engines, 164

flat-bottomed shapes, 130, 131–32, 134

Flight Research Center programs, x

gliders, 164

heat of reentry, 70, 126, 130, 145

human space flight vehicles, 170

internal volume of, 132–33, 159

launch vehicle, glider affects on, 144, 182, 

182n143

L/D ratios, 131–32, 133–34, 135, 152, 

170n111

length of reentry, 70, 126

limitations of, 44–45, 69, 126

orbital lifting bodies, 159–160

piloted lifting bodies, 142–159, 145, 146, 

148, 154, 157

preference for, 130

reusability of vehicles, 142, 159–160, 186

rocket-powered, 144, 164

runway landings, 125

supersonic flight in, 149–150

support for designs, 44–45, 125, 126–27, 

130

unpowered landing concept, x, 156–57

lift-to-drag (L/D) ratios

accuracy of recovery and, 46, 93, 116, 125

Apollo program, 125, 130, 170n111

cross-range capabilities and, 46, 170, 

170n111

Dyna-Soar program, 176, 176n127, 178–79

equilibrium altitude and, 127–28

Gemini program, 64, 125, 130

glider vehicles, 127–28

g-loads and, 126, 131

hypersonic gliders, 176n127

hypersonic L/D coefficient, 125–26, 125n2, 

127–28

hypersonic L/D research, 152

HYWARDS program, 167, 168

lifting body concepts, 131–32, 133–34, 

135, 152, 170n111

Mercury program, 130

PRIME vehicle, 142

reentry and, 46

Space Shuttle, 187

unpowered landings and, 156

X-24B (SV-5P-2) lifting body, 155–59, 157

Ling-Temco-Vought (LVT), 74, 202, 208

liquid cooling systems, viii, 19, 27–28, 40, 

40n15, 165, 224–25. See also water-wall 

(double-wall) cooling systems

liquid lithium, 165

liquid sodium, 165

Little Joe flight tests, 85, 86, 87, 89

Lockheat, 196

Lockheed JC-130B Hercules, 141, 142

307
Coming Home

Lockheed/Lockheed Missiles and Space 

Company

ceramic insulation development, 195

Dyna-Soar program, 171

hypersonic bomber concepts, 164

ILRV design, 186

launch vehicle booster program, 180n139

Polaris Mk series warheads, 29

reusable surface insulation (RSI), 195–96, 

198–200, 200n20, 201–2

Space Shuttle program, 198, 199–200

X-17 research, 21

Lockheed Martin

carbon-carbon nose caps, 208

Orion CEV, 258–59, 260

X-33 VentureStar program, 226–230

Lohman, Robert L., 159

Love, Eugene S. “Gene,” 114, 134

Love, Michael V., 156, 157

Lovell, James A., Jr.

Apollo program missions, 78–79, 123

Gemini program missions, 67, 68, 69

human space flight programs, support for, 

268

Low, George, 200, 200n23

low-temperature reusable surface insulation 

(LRSI), 212–13, 215

Ludox, 205, 214

Lyndon B. Johnson Space Center. See Johnson 

Space Center/Manned Spacecraft Center

M

M1 lifting body, 133–34

M1-L lifting body, 134, 139, 143

M2 lifting body program

ablative heat shields, 147–48

air tows, 143–44, 145

car tows, 143

flight testing, 142–43

M2b Cadillac lifting body, 134

M2-F1 lifting body, 136, 139, 143–44, 145, 

153

M2-F2 lifting body, 144–47, 146n53, 146, 

153

M2-F3 lifting body, 133, 147–48, 148, 149

NB-52 carrier aircraft launch, 144, 145, 146

orbital lifting body research and, 160

Titan launch vehicles, 144, 147

M-103 project, 135

MA-25S ablator, 191

magnesium-base alloys, 41

Maja, Valtteri, 116

Man In Space Soonest (MISS) program, 48–50, 

48n33, 49n37

Manke, John A., 148, 149, 156, 157

manned capsule concept, 63

Manned Glide Rocket Research System, 44

Manned Orbiting Laboratory (MOL) program, 47, 

67, 67n75, 179n137, 180, 180n139

Manned Spacecraft Center. See Johnson Space 

Center/Manned Spacecraft Center

Mark II program, 62. See also Gemini program

Marshall Space Flight Center (MSFC), 186

Mars missions

ablative thermal protection systems, 

248–49, 250, 252, 278, 280, 280, 281, 

282, 283–84

Mars Exploration Rover (MER) program, 249, 

250, 278, 280, 280

Mars Science Laboratory (MSL), 252, 278, 

281, 284

Pathfinder program, 249, 250, 280, 280

Viking Lander program, 246, 249, 280

Vision for Space Exploration initiative, 

254–55, 275–76

Martin Company. See Glenn L. Martin Company

Martin Marietta Corporation

ILRV design, 186

merger and formation of, 173n120

orbital lifting bodies, 159–160

308
Index

Space Shuttle program, 190, 198

See also Glenn L. Martin Company

Martin Marietta X-24A (SV-5P) lifting body, 133, 

153–55, 154, 156, 158, 238, 240

Martin Marietta X-24B (SV-5P-2) lifting body, 

152, 155–59, 157

Martin SV-5/A3-4 lifting body, 135, 137, 140, 

153

Martin X-23A (SV-5D) Precision Reentry 

Including Maneuvering reEntry (PRIME) 

vehicles, x, 140–42, 153, 238

Mattingly, Thomas K., II, 119

McCurdy, Richard, 200, 200n23

McDivitt, James A., 121

McDonnell

aerothermodynamic structural vehicle (ASV), 

138–140, 138n31

aerothermoelastic vehicle (AEV), 138–39

boost-glider program, 140

Dyna-Soar program, 171

Gemini capsule heat shields, 72

Mercury capsule manufacturing, 58–59, 60

Mercury landing system testing, 85

S-3 silicon elastomeric ablator, 64, 72

McDonnell Alpha Draco Model 122B boostglider, 137

McDonnell Douglas

ILRV design, 186

Space Shuttle program, 190, 198, 200

McLellan, Charles H., 39, 130

McMillan, Brockway, 179, 179n137

McNamara, Robert, 173, 179, 180–81, 186

Mercury program

ablative heat shields, ix, 2, 47, 50–55, 53, 

56, 58–62, 61, 64–65, 70, 72

Air Force role in, 50

approval of, 62, 62n66

ballistic reentry of capsule, 46, 51–52, 70

ballistic technologies and, 3

beryllium heat sinks, 52, 54

Big Joe flight tests, 54–56, 57, 85

boilerplate capsule manufacturing and flight 

tests, 54–56, 55, 57

capsule afterbody shingle insulation, 58, 

65, 70

capsule design and development, ix, 19n40, 

50–62, 53, 55, 170, 185

capsule testing, 51, 58

Carpenter flight, 60

Carpenter flight (Aurora 7), 61

contour couches for astronauts, 85–86

contributions of to piloted space flight, 44

Cooper flight, 62

Enos flight, 60

flight testing, 54–56, 57, 59–60

Glenn flight, 60

g-loads, 51, 56

heat of reentry, 51–52, 56, 70, 130

honeycomb energy-absorption system, 85, 

86

landing and recovery system testing, 83–84, 

84, 85–90, 87, 89

Little Joe flight tests, 85, 86, 87, 89

MISS program, 50

number of flights, 81, 81n1

parachute reentry and recovery, ix, 51, 59, 

80, 82, 83–90, 84, 92, 99, 113

pig drop tests, 85

Reef, Project, 88

research goals, 44

safety record of, 2, 2n4

Schirra flight, 60, 62

water landings and recovery, ix, 51, 59, 82, 

83, 90

metal alloys

aircraft construction, 22, 23n52

characteristics of, 22, 22nn48–49, 23, 25, 

25n62

heat absorption testing, 24, 24n53, 41

heat shield applications, 71

309
Coming Home

heat sink applications, 22–25, 22nn48–49

HL-10 lifting body shingle structures, 151

refractory materials, 23, 71, 140

metallic superalloys

Dyna-Soar program and development of, 

175

heat shield applications, 140, 194

HL-10 lifting body shingle structures, 151

hot structure applications, 195–96

HYWARDS program, 169

metallic thermal protection systems

Apollo program, 70, 71

blunt-body reentry design heat shields, ix

Dyna-Soar program, x, 70, 71, 176–77, 178, 

186, 192, 195

heat shield concepts, x

materials for, 71, 140, 192–94

National Aero-Space Plane, x, 221, 224, 

225, 232

oxidation of materials, 168, 177, 192–94

refurbishment of, 177, 178, 186, 191

removable panels, 191

Space Shuttle program, 192–94

X-33 program, x, 221, 230–33

meteorites, 19n38

methylmethacylate (plexiglass), 196

Micro-Quartz insulation (Q-Fiber Felt), 177, 

177n128

Min-K insulation, 65

Minneapolis-Honeywell, 181

Minuteman missiles, 29, 31, 31n73

missiles

definition of, 11n24

design and shapes of, 16–17, 19n40

guidance for, types of, 11n24

nose cones, 63

range and velocity of, 16

weight considerations, 24

See also ballistic missiles and technologies

mistake avoidance, 285

Mk 1 warheads (Navy), 25, 25n61, 29

Mk 4 warheads (Air Force), 29

Mk 5 warhead (Navy), 29, 31

Mk 6 warhead (Air Force), 29, 31

Mk 11 warhead (Air Force), 31

Mk 12 warhead (Air Force), 31, 31n73

Mk 21 warhead (Air Force), 31

Mk I warhead (Air Force), 25, 25n58, 25n61, 28

Mk II warheads (Air Force), 19n40, 25, 25n58, 

28, 50–51, 52

Mk III warheads (Air Force), 25n58, 29

MOL (Manned Orbiting Laboratory) program, 47, 

67, 67n75, 179n137, 180, 180n139

molybdenum alloys

Dyna-Soar program, 177, 181

hot structures and characteristics of, 168

titanium-molybdenum alloy, 71

TZM alloy, 140, 177

Monel and Monel K, 22, 193

Moon

Apollo program Moon landings, 77, 117, 185

funding for missions, 266

Kennedy Moon-landing-challenge speech, 

69, 186

Orion CEV missions, 255, 257, 260

support for missions to, 266

Vision for Space Exploration initiative, 

254–55, 275–76

MSC-001 (DC-3) concept, 184, 187–88

Mueller, George E., 111–12

mullite and mullite shingles, 194–95, 200, 

200n20, 201

Multhopp, Hans, 135

MX (Peacekeeper) missiles, 31

MX-2276 project (Bomber-Missile [BoMi] skipglider), 160–64, 162, 163n93, 165–66, 170

N

National Aeronautics and Space Act, 50

310
Index

National Aeronautics and Space Administration 

(NASA)/National Advisory Committee for 

Aeronautics (NACA)

Aeronautics Research Mission Directorate, 

vii–viii

competence of, 236

costs of research and technological 

developments, 227–230

faster-better-cheaper strategy, 228–29

human space flight program, 48, 50

hypersonic flight research, 21

Hypersonic Research Steering Committee, 

170

number of space flights, 81, 81n1

programs and initiatives, review of, 264–68, 

270

R&D roots of, 268

rocket-powered hypersonic aircraft, interest 

in, 9

warhead research, 29

National Aero-Space Plane (NASP, X-30)

active cooling concepts, 224–25

cancellation of program, 223

design of, 220

development of, 221–23, 254

funding for, 223

hypersonic L/D research, 152

metallic thermal protection system, x, 221, 

224, 225, 232

National Air and Space Museum, Smithsonian 

Institution, 56, 259

National Reconnaissance Office, 32, 179n137

Navaho missile, 171n114

Naval Surface Warfare Center, 202

Navy, U.S.

capsule recovery operations, 56, 59, 65, 66, 

80, 82, 116, 123, 284

Mk series warheads, 25, 25n61, 29, 31

Polaris submarine-launched missiles, 25, 29

reentry vehicles, 25, 25n61

X-17 research, 21

NB-52 carrier aircraft, 144, 145, 146, 149, 156

Neice, Stanford E., 127, 128n8

Neptune missions, 282

Newton, Isaac, 131

Nextel 440 fabric insulation, 209

Nextel AB312 alumina-borosilicate fibers, 212, 

216

nickel and nickel alloys

heat shield applications, 71, 193, 194

hot structure applications, 41

Inconel X, 23n52, 41n16

nickel-chromium-iron stainless steel heat 

sinks, 23

thoria dispersed nickel-chromium (TD NiCr), 

193

Nininger, H.H., 19n38

Nitronic, 193

Nixon, Richard M., 198

Nomex felt blankets, 204, 213–15

Nonweiler, Terence R.F., 152n67

Norman, Lee, 101

North American Aviation

Apollo program heat shield development, 

70, 74

Apollo program parasail system, 122

conical piloted reentry vehicle with rotors, 

182–83

Dyna-Soar program, 171

Gemini program paraglider concept, 93, 

94–95, 96, 105, 107, 109, 112

hypersonic bomber concepts, 164

Test Tow Vehicle, 94–95

North American Rockwell

Apollo program parachute landing system, 

118

ILRV design, 186

name change, 201n25

Space Shuttle program, 153n70, 190, 198, 

200–202, 205–7

311
Coming Home

North American X-15 aircraft. See X-15 research 

airplane

North American XB-70/XB-70A/B-70 bomber, 

135, 152n55, 152n66

Northrop

Dyna-Soar program, 171

HL-10 lifting body, 144, 149

M2 lifting body program, 144, 147

Mercury landing and recovery system 

development, 86

NorAir Division, 144

Northrop Ventura, 122

nose caps

boost-gliders, 140

Dyna-Soar program, 173n119, 177, 181

PRIME vehicle, 141

Space Shuttle program, 173n119, 199, 

201–2, 208

X-38 program, 242–43

nose cone designs, 63

nylon, 29, 191

O

Obama, Barack, 254, 265–66, 267, 268, 276

Oberth, Hermann, 2, 5n8, 7, 9, 11

ocean landings. See water landings and recovery

operating experience, 285

orbital lifting bodies, 159–160

orbital maneuvering system (OMS) module/pod, 

199, 202, 211, 212, 214, 215, 217

Orbital Space Plane program, 273

Orion Crew Exploration Vehicle (CEV)

ablative thermal protection system, x, 221, 

252, 257–260, 262–63, 281, 282, 

283–84

development and design of, x, 221, 255

mission capabilities, 255, 257, 260, 264, 

268, 275–76

parachute reentry and recovery, x, 221, 261, 

263–64, 271, 284–85

schedule for, 264

Orion Multi-Purpose Crew Vehicle, 258

oscillations

Apollo command module, 118–19

pilot-induced oscillations, 102, 147, 147n55

overshoot boundary, 125

P

paint, temperature-sensitive, 175

Painter, Wen, 124, 135n24, 152n64, 284

parachute reentry and recovery

Apollo program, x, 79, 83, 113, 117–123, 

119, 121, 123, 264, 286

capsule reentries and landings, ix, 82–83, 

129

challenges related to, 83

Chinese space program, 82

CORONA program film-return capsules, ix, 

32, 33, 35, 82, 152

development of, 82, 284

drogue chutes, Apollo program, 117, 

118–120

drogue chutes, Gemini program, 91, 92

drogue chutes, Genesis program, 245, 246, 

249

drogue chutes, HSTV, 104

drogue chutes, Mercury program, 83, 85, 

86, 88

drogue chutes, Orion CEV, 261, 263, 264

Gemini program, x, 68, 82, 91–92, 91, 99, 

109, 111, 113

Genesis program, 246–47, 247, 249–250

limitations of, 116, 126

Mercury program, ix, 51, 59, 80, 82, 83–90, 

84, 92, 99, 113

midair recovery techniques, ix

number of parachutes, 82

Orion CEV, x, 221, 261, 263–64, 271, 

284–85

PRIME vehicle, 141, 142

312
Index

return to use of, 284–85

ring-sail parachutes, 84–85, 86, 88

riser connectors, 79, 122

Russian/Soviet Union space program, 82

SCORE project, ix, 82

squidding/breathing/rebound problems, 

84–85

success of, 82

paraglider/parawing landing systems

advantages of, 96, 116

cancellation of program, 105, 107, 109, 

111, 112

crew escape system, backup, 97, 98

deployment of, 93, 96–97, 98, 105

failure of, reasons for, 112–13

FIRST program, 94

flight testing, 108, 111

full-scale test vehicle (FSTV) testing, 109, 

111

full-scale trainers, 107

funding for development, 112

Gemini program, 94–114, 97, 100, 110, 

116, 238n37

half-scale test vehicle (HSTV) testing, 104–7, 

106, 111

Inflatable Micrometeoroid Paraglider, 94

limitations of and problems associated with, 

98–99, 105, 126

Parasev (Paraglider Research Vehicle), 94, 

101–4, 103

reentry and landing phase, 98

return to use of, 123, 238n37

Rogallo wing paraglider recovery system, 

ix–x, 62–63, 82, 91, 92–94, 93, 95, 111

success of, 111, 112

support for, 96, 98, 107, 113–14, 116, 

116n66

Test Tow Vehicles (TTVs), 94–95, 107–9, 

111

X-38 parafoil landing concept, 236, 

238–240, 238n37, 239, 241

parasail concepts

Apollo program, 101, 122

development of, viii

Gemini program, 99–101, 111–12, 115

return to use of, 123

Parasev (Paraglider Research Vehicle), 94, 

101–4, 103

passive thermal protection systems, 26, 207–8. 

See also heat sinks; hot structures

Pathfinder program, 249, 250, 280, 280

Paulson, John W., Jr., 134

Peacekeeper (MX) missiles, 31

Peterson, Bruce, 103, 146–47, 146n53, 149

phenolic-impregnated carbonaceous ablator 

(PICA)

applications for, 248, 250, 252, 257–59, 

262–63, 271, 280, 280, 281, 282, 

283–84

characteristics and capabilities, 252, 257, 

282

development and testing of, 250, 252, 

257–59, 260, 262

phenolic resin

carbon-phenolic composites, 141, 191, 248, 

262, 280, 282

heat shield applications, 55, 64, 65, 72

pig drop tests, 85

Piland, Robert O., 29, 69n80

piloted lifting bodies, 142–159, 145, 146, 148, 

154, 157

pilot-induced oscillations, 102, 147, 147n55

Pilotless Aircraft Research Station, Wallops 

Island

blunt-body reentry design testing, 21

free-flight model testing, 17, 46

laminar flow research, 46

Mercury capsule testing, 56, 58

313
Coming Home

Mercury landing and recovery system 

testing, 84, 86, 87, 89

warhead design and shape testing, 17

PILOT project, 140, 153

Pioneer Parachute Company, 99–100, 122

Pioneer Venus mission, 246, 280

planning and prevention actions, 285

Plan of Space Exploration (Tsiolkovskiy), 4

platinum-rhodium wire, 177

plexiglass (methylmethacylate), 196

Polaris submarine-launched missiles, 25, 29, 

230

Precision Reentry Including Maneuvering reEntry 

(PRIME) vehicles. See PRIME (Precision 

Reentry Including Maneuvering reEntry) 

vehicles

Preliminary Design of an Experimental WorldCircling Spaceship (Project RAND), 14–15

PRIME (Precision Reentry Including Maneuvering 

reEntry) vehicles

ablative thermal protection systems, 

140–41, 158

cross-range capabilities, 140, 142

goals of program, x, 140

Space Shuttle design and, 43, 142, 201

START program, 140, 153

SV-5D (X-23A) lifting body, x, 140–42, 153, 

238

SV-5 lifting bodies, 237–38

private-sector spaceflight initiatives, xi, 81, 229, 

267, 268, 270. See also Commercial Crew 

Development (CCDev) program

propulsion

for hypersonic flight, 38, 39

research and technological developments, 

vii–viii

X-33 VentureStar program, 233, 234

See also engines

propulsive lift, sustained, 126

Pugel, Elizabeth “Betsy,” 259

Purser, Paul E., 29

Putt, Donald L., 48

Q

Q-ball, 42

Q-Fiber Felt (Micro-Quartz insulation), 177, 

177n128

R

R4D (C-47/DC3) aircraft, 144

radiative cooling

boost-glider designs, 128–29

development of, 9

Dyna-Soar program, 129n12, 176, 179, 192

effectiveness of, 129n12

hot structures and, 9, 26

shingle insulation, 58, 65

skip-gliding (dynamic soaring), 8, 9, 129

spaceplanes and, 15

See also metallic thermal protection systems

radome, 196

Rainey, Robert W., 134

RAND, Project, 14–15

Rasky, Dan, 271

Raymes, Frederick, 135

RCA, 181

reaction control rockets, 42, 44

reaction control system (RCS) and RCS module/

pod, 120, 122, 199, 211, 212, 214, 217

reaction-cured glass (RCG) coating, 211, 212, 

213

Reaction Motors XLR11 rocket engines, 145, 

150, 153

Reagan, Ronald, 222–23

reconnaissance programs and missions

ablative thermal protection applications, ix

CORONA program, ix, 31–33, 33, 35, 82, 

135

film-return capsules and vehicles, ix, 32, 33, 

35, 82, 132, 135, 152

314
Index

parachute reentry and recovery, ix

recovery

challenges related to, vii–viii

research and technological developments, 

284–85

research and technological developments to 

support, vii–viii, ix

spaceplanes and, 3

See also parachute reentry and recovery; 

runway/land landings and recovery; water 

landings and recovery

Redstone missiles, 12–13, 16, 59, 132, 185

Reed, Donna, 135, 139, 143

Reed, Robert Dale

lifting body concepts, 135, 139, 236

M2 lifting body program, 136, 142–43, 144

Reef, Project, 88

reentry and return from space

accidents during, vii, 81, 209–10

ballistic coefficient and, 20, 20n41, 25, 26, 

29, 31

ballistic technologies and, viii, 1–2, 3, 

16–26, 19n40, 45–46, 51–52, 70, 279

challenges related to, vii–viii, xi, 279

concept of reentry, 8n16

concerns about, discussions of, 35–36

costs of, 122–23

difficulty and risks of, 1, 81

direct reentry, 9

human requirements, 1–2

lessons from, 285

research and technological developments, 

284–85

research and technological developments to 

support, vii–xi, 1, 14–15

semiballistic capsules, 46–47

shadowgraphs of reentry designs, 63

simplicity of, 36

spaceplanes and, 3

speed of return, 1

V-2/WAC Corporal rocket vehicle testing, 12

X-plane research goals, 39

See also heat of reentry

reentry vehicles. See warheads (reentry vehicles)

refractory thermal protection

ASSET program, 140

Dyna-Soar heat shields, x, 71, 176–77, 178, 

186

HL-10 lifting body shingle structures, 151

materials for, 23, 71, 140, 192–94

oxidation of materials, 71, 177, 192–94

refurbishment of, 177, 178, 186, 191

See also metallic thermal protection systems

reinforced carbon-carbon (RCC), x, 155, 195, 

199, 200, 201–2, 203, 208–10, 210

René 41

capsule afterbody shingle insulation, 58, 65, 

67, 69, 70

composition of, 58

Dyna-Soar construction, 176, 178, 181

heat shield applications, 71, 193

Republic Aviation

Apollo program subscale model contract, 74

Dyna-Soar program, 171

hypersonic bomber concepts, 164

Resnikoff, Meyer M., 131–32

reusable launch vehicles (RLVs), 221–22, 

222n1, 225, 235

Reusable Small Booster Program (X-34), 

158n83, 221, 234–35, 235

reusable surface insulation (RSI) blankets, 204, 

208, 214–17, 216n56, 234, 235, 240

reusable surface insulation (RSI) tiles

Boeing Replacement Insulation (BRI)-18 

tiles, 216

ceramic RSI, 195–96, 198–200, 200n20, 

201–2, 208, 210–17

densification and strengthening process, 

205, 206, 214

density of, 210

315
Coming Home

development of, 195–96, 280

diced tiles for deflection, 202

elastomeric reusable surface insulation 

(ERSI), 202, 204

fibrous refractory composite insulation 

(FRCI), 211–12, 216

gap fillers, 216

high-temperature reusable surface insulation 

(HRSI), 202–4, 203, 211–12, 213, 216, 

216n56

installation of, 196, 205–7

LI-900 tile insulation, 196, 204, 210, 216, 

216n56, 262

LI-1500 tile insulation, 196, 200

LI-2200 tile insulation, 196, 202–4, 203, 

210, 216n56, 262

LI-3000 tile insulation, 204

low-temperature reusable surface insulation 

(LRSI), 212–13, 215

maintenance and replacement of, 212, 213

materials for, 196, 201–2, 274–75

number of types on orbiter, 208, 210, 216

part numbers, 213

proof testing, 206

reaction-cured glass (RCG) coating, 211, 

212, 213

RTV adhesives, 196, 206, 214, 215

silicone-impregnated reusable ceramic 

ablator (SIRCA) tiles, 235, 250, 280

Space Shuttle RFP requirement, 198–200, 

200n20

strain-isolation pads (SIP), 204–5, 206, 

213–14

strength of, 204

strength of, problems with, 204–6

toughened unipiece fibrous insulation (TUFI), 

213, 216, 262, 274–75

weight of, 210, 215

Reuther, James, 258, 259

Richards, Charles, 102

Rickey, John, 135

ring-sail parachutes, 84–85, 86, 88, 92

Rocket Bomber (RoBo, Air Force SR-126), 

164–66, 170, 171

rocket engines

acceleration requirement for vertical liftoff, 

14

for aircraft, 5

for automobiles, 5

German research and designs, 5, 9, 11

getting into space, force and power for, 

13–14

hypersonic aircraft, rocket-powered, 9, 

21–22, 39

HYWARDS program, 167

landing rockets, 122–23

lifting bodies, rocket-powered, 144, 164

liquid-fueled, 4, 7

reaction control rockets, 42, 44

replacement of internal combustion engines 

with, 5

turbine rocket ship design, 4

XLR11 rocket engines, 145, 150, 153

XLR99 rocket engines, 167

Rocket into Interplanetary Space, The (Die 

Rakete zu den Planetenräumen, Oberth), 5n8, 

7

rockets, 4–5, 11n24. See also rocket engines

Rockwell International/North American Rockwell

name change, 201n25

Space Shuttle design, 190, 198, 200

Space Shuttle orbiter contract, 153n70, 

200–202, 205–7

See also North American Rockwell

Rogallo, Francis M., viii, 93, 95–96, 95

Rogallo, Gertrude, 95

Rogallo wing paraglider recovery system, ix–x, 

62–63, 82, 91, 92–94, 93, 95, 111, 126

Rogers, Russell L., 178n132

Rohr Industries, 232

316
Index

role mode, 147n55

room temperature vulcanizing (RTV) adhesives, 

196, 206, 214, 215, 216

Roton rotary rocket, x

rotors, 126

RSI. See reusable surface insulation (RSI) 

blankets; reusable surface insulation (RSI) tiles

RTV (room temperature vulcanizing) adhesives, 

196, 206, 214, 215, 216

runway/land landings and recovery

Chinese space program, 82

cross-range capabilities and, 132, 187, 198

delta-wing concepts, x

Dyna-Soar program, x, 178–79, 181–82

first successful, 101

Gemini program and Rogallo wing paraglider 

recovery system, ix–x, 62–63, 82, 91, 

92–94, 93

launch site, landing at, 132, 187, 198

Russian/Soviet Union space program, 82

side vision, landing with, 149, 177–78

spaceplanes, 126

Space Shuttle program, 158, 219

support for airplane-like designs, 92–94, 

113–14, 116, 124, 125, 129, 284

unpowered landings, x, 156–57

Rushworth, Robert A., 156–57

Russia/Soviet Union

ICBM program, 170

number of space flights, 81, 81n1

parachute reentry and recovery, 82

space flight accident and loss of crew, 81

space race against U.S., 82, 169–170

Sputnik, 36, 48, 169, 170

Sputnik II, 170

Rutan, Burt, 240

S

S-3 silicon elastomeric ablator, 72

safety and preventive actions, 285

Sam and Mercury recovery system testing, 86

Sänger, Eugen

antipodal bomber development, ix, 7n15, 

8n17, 81–82

dynamic soaring (skip-gliding), 7–8, 9

equilibrium altitude, 127–28

hypersonic airflow, analysis of, 6

hypersonic flight and research of, 9

lifting body concepts, 131

research and technological developments 

by, viii

Silverbird, 7–9, 7n15, 8, 128, 131, 160, 161

Techniques of Rocket Flight, 7, 7n13

satellites, viii

Saturn IB launch vehicles, 74, 75, 159

Saturn missions, 282

Saturn V launch vehicles, 76

Scaled Composites, Inc, 236, 240

schedules, 285

Schirra, Wally, 60, 62

Schweickart, Russell L., 121

scientific missions

ablative thermal protection system 

applications, ix, 245–46, 247–48, 250, 

252–53

recovery of scientific samples, x–xi, 221, 

244–253, 245, 247, 251, 253

SCORE (Signal Communication by Orbiting Relay 

Equipment) project, ix, 82

Scott, David R., 121

scramjet engines, 222, 234

Seamans, Robert C., 111–12

sea-marker dye packets, 88, 90, 118

semiballistic/semilifting capsules, 46–47, 70, 

130

semipassive thermal protection systems, 27. See 

also ablative thermal protection systems

Sergeant Battlefield missile, 137

shadowgraphs, 63

Shepard, Alan, 52

317
Coming Home

shingle insulation

capsule afterbody shingle insulation, 58, 65, 

67, 69, 70

HL-10 lifting body shingle structures, 151

radiative cooling, 58, 65

Space Shuttle program thermal protection, 

195–96

shock layer, 6, 52, 279

shock waves

blunt-body reentry designs, reentry heat, 

and, 17, 19, 63

hypersonic flight and, 6, 41, 43

shadowgraphs of reentry designs, 63

X-15 model testing, 41

Sierra Nevada Corporation, xi, 270, 272

Signal Communication by Orbiting Relay 

Equipment (SCORE) project, ix, 82

silane products, 205, 205n35

silica/silica fiber composites

ablative heat shields, 140–41, 191

Apollo program heat shield, 72, 258

Apollo program radome, 196

ceramic RSI, 195–96, 198–200, 200n20, 

201–2, 208, 210–17

HCF development, 194–95

Ludox, 205

tile densification and strengthening process, 

205, 206, 214

silicide coating, 177, 181

silicon-based ablators, 64, 72, 141, 191, 202

silicon carbide, 71, 165, 177, 208, 209

silicon dioxide, 71

silicone-impregnated reusable ceramic ablator 

(SIRCA) tiles, 235, 250, 280

siloxane resin, 141

silver, 22

Silverbird, 7–9, 7n15, 8, 128, 131, 160, 161

single-stage-to-orbit (SSTO) vehicles, 225, 227, 

233–34

680A program (Spacecraft Technology and 

Advanced Reentry Tests, START), 140, 153, 

237–38

Six Million Dollar Man series, 146n53

skip-gliding (dynamic soaring), 7–8, 9, 44, 127, 

128, 129–130, 163–64. See also BomberMissile [BoMi] skip-glider [MX-2276 project]

SLA-561/SLA-561S/SLA-561V ablator, 200, 

248, 249, 262, 280

Smithsonian Institution, National Air and Space 

Museum, 56, 259

Snyder, George M., 175

social construction of technology, 114, 116n66

Solar, 140

Soulé, Hartley A., 39–40

sound barrier, 37

Soyuz spacecraft program

accidents and loss of crew during, 81

ISS crew transportation, 244

number of space flights, 81, 81n1

safety record of, 2n4

spacecraft

commercial piloted spacecraft, 221, 266

cooperative agreements and coordination of 

development, 227–29, 281

costs of research and technological 

developments, 227–230

reusability of, 142, 159–160, 186, 193, 

221–22, 222n1, 225, 227, 235

slender, streamlined designs, 16–17, 19n40, 

20, 31, 63

Spacecraft Technology and Advanced Reentry 

Tests, Program 680A (START), 140, 153, 

237–38

Space Exploration Initiative, 254

Space Exploration Technologies (SpaceX), xi, 

270–71

space flight

accidents during launch, 81

ballistic missiles and technologies and, 12

318
Index

getting into space, force and power for, 1, 

13–14

number of flights, 81, 81n1

paths and orbits between planets, 14, 14n28

programs and initiatives, review of, 264–68, 

270

speed needed to escape gravity, 3, 14

von Braun’s vision of, 11

See also human space flight; reentry and 

return from space

Space Maneuver Vehicle program, 273

spaceplanes

abandonment of idea, 3

advantages of, 3, 125–27, 128

concept of, 125, 125n1, 222

design and shapes of, 3, 126

engines for, 222

first-generation spaceplane, 185

future designs and use of, 3–4

hypersonic flight technology and, 5–6

landing systems, ix

launch vehicles, 185–86

limitations of, 126

orbital velocities of, 3

preference for winged vehicles over 

capsules, 92–94, 113–14, 124, 129, 130, 

185

runway landings, 126

smooth aerodynamic surfaces, 126

support for designs, 124, 125, 126–27, 284

thermal protection systems for, 126

transonic flight research and development 

of, 38

vision of human spaceflight and, 2–6, 15, 

81–82, 185–86, 222

von Braun’s applications for, viii, ix, 34

wings and reentry technique, 2, 3, 15

X-37/X-37B vehicle, 272–75, 274, 276, 277

See also National Aero-Space Plane (NASP, 

X-30)

space race

capsule designs and, 2, 3, 47, 69–70

Navy ships for capsule recovery during, 123

parachute reentry and, 82, 92

Sputnik, Sputnik II, and, 169–170

SpaceShipOne, 81, 81n1

Space Shuttle program

aluminum orbiter construction, 202, 207, 

209

approval of, 198

ASSET and design of, 43, 142, 201

blunt-body reentry designs, x, 187–88

contractors, RFPs, award of contract, 186, 

190, 198–201

costs of research and technological 

developments, 227–28

cross-range capabilities, 132, 186–87, 188, 

189, 190, 198, 201

delta-wing concepts, x, 188–89, 190, 198, 

199–200

design and shape of orbiter, 187, 193, 197, 

198

development and construction of orbiter, 

40n14, 186–201

Dyna-Soar and design of, 182, 186, 201

entry, approach, and landing procedures, 

217–19

first-generation spaceplane, 185

glider landing technique, 151–52

graphite (carbon) construction, 24

grounding of Shuttle fleet, vii

guidance and control, 217–19

heat of reentry, 207–8, 217–19

HL-10 lifting body and design of, 150, 

151–52

jet engines, x

landing engines, 156, 198

launch sites, 180n139

L/D ratios, 187

lessons from, 285

319
Coming Home

lifting body design, 187

managerial, organizational, and cultural 

issues associated with, vii

maneuverability of, 132

MSC-040C orbiter, 198–200, 201

nose caps, 173n119, 199, 201–2

number of flights, 81, 81n1

orbital maneuvering system (OMS) module/

pod, 199, 202, 211, 212, 214, 215, 217

orbiter contract, 153n70

painting flags and letters, 217

payload bay and mission goals, 187

Phase A design and development, 186–89, 

191

Phase B design and development, 186, 

189–191, 190, 195

Phase C design and development, 186, 

198–200, 201–2

Phase D design and development, 186, 

198–200

pilot-induced oscillations, 102

PRIME and design of, 43, 142, 201

reaction control system (RCS) and RCS 

module/pod, 120, 122, 199, 211, 212, 

214, 217

research and technological developments to 

support, 43–44

retirement of and replacement for, 253–54, 

275–77, 281, 284

reusability of vehicles, 186, 193, 198, 

222n1, 227–28

runway landings, x, 158, 219

spaceplane concepts and, 3

straight-wing design, 187–89, 190

structural tests, 205

success of, 185

support for program, 284

technical problems with, vii

unpowered landing concept, x, 156

wings and reentry technique, 2

X-15 and design of, 43–44, 142, 201

X-24A (SV-5P) and design of, 154–55

See also specific orbiters

Space Shuttle program thermal protection

ablative thermal protection systems, 199

ceramic tiles, x, 2, 155, 159, 199

damage to, vii, x, 207, 209–10, 219

debris strikes, 209–10, 283

development and design of, 190–96, 280

effectiveness of, 221

failure of, x, 2

HCF materials, 194–95, 200, 200n20, 201

HYWARDS program and, 169

insulating materials, 195–96, 208

leeward-side heating, 202, 208

OMS/RCS pods, 202, 211, 212, 214, 215

passive system, 207–8

refurbishment of, 207

reinforced carbon-carbon (RCC), x, 155, 195, 

199, 200, 201–2, 203, 208–10, 210

removable, 191

RSI tiles (see reusable surface insulation 

(RSI) tiles)

strain-isolation pads (SIP), 204–5, 206, 

213–14

testing of, 158

wing leading edges, 199, 200, 201–2, 203, 

208–10, 210

space station, 187. See also International Space 

Station (ISS)

Space Task Group, 54

Space Travel Symposium, 36

spaceworthy orbital Bomber-Missile (S-BoMi), 

165–66

Spearex paint, 213

speed

gravity, speed needed to escape, 3, 14

reentry and return, speed of, 1

X-plane speed records, 37, 42, 42n19

spiral mode, 147n55

320
Index

splashdowns. See water landings and recovery

Sputnik, 36, 48, 169, 170

Sputnik II, 170

SSTO (single-stage-to-orbit) vehicles, 225, 227, 

233–34

stability research and technological 

developments, vii–viii

Stack, John, 9

stainless steel, 72, 140

Stamer, Friedrich, 5

Stardust program, xi, 221, 246, 247, 249, 

250–53, 251, 253, 257, 258, 280, 280, 282

START (Spacecraft Technology and Advanced 

Reentry Tests, Program 680A), 140, 153, 

237–38

Stearman sport biplane, 103

Stewart, David, 275

Storms, Harrison A. “Stormy,” Jr., 39–40, 42

strain-isolation pads (SIP), 204–5, 206, 213–14

Strong, 56, 59, 80

supersonic flight, 144, 149

Supersonic Pressure Tunnel (Langley), 41

SV-5 lifting bodies

orbital lifting body research and, 160

SV-5/A3-4 lifting body, 135, 137, 140, 153

SV-5D (X-23A) lifting body, x, 140–42, 153, 

238

SV-5J lifting body, 155–56

SV-5P (X-24A) lifting body, 133, 153–55, 

154, 156, 158, 238, 240

SV-5P-2 (X-24B) lifting body, 152, 155–59, 

157

X-38 Crew Return Vehicle (CRV), 237–38

Swigert, John L., Jr., 123

synar-silicon carbide, 177

Syvertson, Clarence A., 133–34

T

TAEM (terminal area energy management) 

interface, 218–19

TD NiCr (thoria dispersed nickel-chromium) alloy, 

193, 194

technical inquisitiveness, 285

Techniques of Rocket Flight (Sänger), 7, 7n13

technology, social construction of, 114, 116n66

Teflon, 29

Teller, Edward, 48

temperature, 22, 35. See also heat of reentry

temperature-sensitive paint, 175

terminal area energy management (TAEM) 

interface, 218–19

Test Tow Vehicles (TTVs), 94–95, 107–9, 111

Textron, 28. See also Avco Corporation/Avco 

Manufacturing Corporation

thermal protection systems (TPS)

aerodynamics and, 126, 169, 283

ASSET program testing, 140

capsule afterbody shingle insulation, 58, 65, 

67, 69, 70

challenges related to, 279

Constellation program, 221

coordination of development, 281

damage to, vii, x, 207, 209–10, 219, 283

delta-wing concepts, 126

environments, natural and induced, 190, 

282–83

function of, 282–83

HL-10 lifting body, 150, 151

HYWARDS program, 168–69

importance of, 279

liquid cooling systems, 19

Mars Science Laboratory (MSL), 278, 281, 

284

materials for, ix (see also specific materials)

number of types to have for projects, 281

refurbishment of, 191, 207

removable panels, 191, 192

research and technological developments, 

vii–viii, 15, 221, 280–84, 285

for spaceplanes, 126

321
Coming Home

Space Shuttle program, x, 2, 155, 158, 169, 

190–96, 198

success of, 279

types of thermal protection, 26–28

for warheads (reentry vehicles), 16

X-34 program, 234–35

X-37/X-37B vehicle, 273–75

X-38 program, 238, 242–43

See also ablative thermal protection systems; 

ceramic thermal protection systems; 

metallic thermal protection systems; 

reusable surface insulation (RSI) blankets; 

reusable surface insulation (RSI) tiles

Thermal Structures Tunnel (Langley), 72

Thermoflex/Thermoflex RF insulation, 58, 64, 

65, 67, 169

Thompson, Milton O.

Dyna-Soar program, 178n132

lifting body concepts, 139

M2 lifting body program, 143, 144, 146, 146

Parasev flight testing, 101

Thor ballistic missiles, 25, 28, 29, 138, 138n31

thoria dispersed nickel-chromium (TD NiCr) alloy, 

193, 194

thoria-tungsten, 140

Titan intercontinental ballistic missiles (ICBMs)

ablative thermal protection applications, 52

development of, 173, 173n120, 174

Dyna-Soar launch vehicles, 173, 174, 

180n139, 182

Gemini-B and MOL program, 67

Gemini program launch vehicle, 66, 91, 

173n120

launch site, 180n139

M2 lifting body program, 144, 147

Mk series warheads, 28, 29–31, 30

orbital lifting body launch vehicle, 159–160

piloted spacecraft launch vehicle, 45, 47, 

185–86

Titan I, 29, 174

Titan II, 29–31, 30, 182

Titan III, 174, 182

Titan IV, 180n139

titanium and titanium alloys

BoMi skip-glider hot structures, 161

ERSI, 202, 204

heat shield applications, 193, 194

hot structure applications, 41

hot structures, 161

HYWARDS program, 169

NASP program, 224, 225

PRIME vehicle, 140

titanium aluminide and composites, 224

titanium-molybdenum alloy, 71

TZM alloy, 140, 177

toughened unipiece fibrous insulation (TUFI), 

213, 216, 262, 274–75

toughened unipiece fibrous-reinforced oxidationresistant composite (TUFROC), 248, 275

transonic flight

aircraft testing of, 37–38

concerns about, 37–38

rocket-powered lifting bodies, 144

spaceplanes and research about, 38

wind tunnel testing of, 37

transpiration cooling systems, 27–28, 165

Trident missiles, 31

Tsiolkovskiy, Konstantin Eduardovich, 2, 4

turbine rocket ship, 4–5

turbojet engines, 150–51, 164

2001: A Space Odyssey, 185

TZM alloy, 140, 177

U

undershoot boundary, 125

Unitary Plan Wind Tunnel (Ames), 18

Unitary Plan Wind Tunnel (Langley), 72

Utah Test and Training Range, 246–47, 247, 

249–250

322
Index

V

V-2 (A-4) missiles, 8–9, 10, 11–12, 16

V-2/WAC Corporal rocket vehicle, 12

Valier, Max, 5, 5n8, 7n13, 9, 11

Vandenberg Air Force Base (AFB), 180n139, 198

variable-geometry wings, 126, 152

VentureStar. See X-33 VentureStar program

Venus missions, 246, 280, 282, 284

Verein für Raumschiffahrt (VfR, German Society 

for Space Travel), 9, 11, 14, 36

Viking Lander program, 246, 249, 280

Virgin Galactic WhiteKnightTwo carrier aircraft, 

xi, 272

Vision for Space Exploration initiative, 254–55, 

264–65, 275–76

Volkswagen Beetle paraglider concept testing, 

100

von Braun, Wernher

ballistic missile development, role in, 11

fluid cooling systems, viii, 28

human space flight, advocacy for, 36

hypersonic airflow, analysis of, 6

public relations skills of, 11n23

Redstone missile development, 12

rocket flight, interest in, 11

space flight vision of, 11

spaceplane applications, viii, ix, 34, 81–82

VfR, work with, 11

von Kármán, Theodore, 15, 38

von Opel, Fritz A.H., 5, 5n9

Vought

boost-glider program, 140

Dyna-Soar program, 171, 173, 173n119, 

177

W

WAC Corporal rocket, 12

Walker, Joseph A., 42, 42n19

Wallops Island. See Pilotless Aircraft Research 

Station, Wallops Island

warheads (reentry vehicles)

ablative thermal protection applications, ix, 

28–31, 30, 52, 279

ballistic coefficient of, 25, 26, 29, 31, 52, 

71, 132

better warheads, development of, 28–31, 30

design and shapes of, 16–17, 19n40, 26, 31

heat sink applications, 25, 26, 28, 35, 52, 

59

high cross range capabilities, 132

interception of, 25

limitations of, 25–26

liquid cooling systems, 19

maneuverable, 132

reentry vehicle name for, 16

thermal protection for, 16

weight considerations, 24

water landings and recovery

Apollo program, 47, 77, 78, 79, 117–18, 

119, 122–23, 123

capsules and human-in-space goal, 3

Constellation program, 3

Gemini program, ix–x, 47, 63–64, 66, 82, 

91, 92

Mercury program, ix, 51, 59, 82, 83, 90

Navy capsule recovery operations, 56, 59, 

65, 66, 80, 82, 116, 123, 284

preference for runway landings over, 92–94, 

124, 129, 284

research and technological developments, ix

sea-marker dye packets, 88, 90, 118

U.S. space program, 82

water-wall (double-wall) cooling systems, 28, 40, 

71, 161, 165

Western Electric, 171

Whitcomb, Richard, 37

White, Bob, 42n19

White Sands Proving Ground, 12

Williams, Bob, 222

Williams, Walter C., 39–40, 42

323
Coming Home

wind tunnels

full-scale 40-by-80 foot wind tunnel (Ames), 

143, 145, 148, 149, 153–54

High Speed Tunnel (Langley), 145, 149

High Temperature Tunnel (Langley), 72

hypersonic tunnel (Ames), 165

Hypersonic Tunnel (Langley), 21, 38, 39

limitations of testing in, 37

slotted-throat wind tunnels, 37

Supersonic Pressure Tunnel (Langley), 41

Thermal Structures Tunnel (Langley), 72

Unitary Plan Wind Tunnel (Ames), 18, 72

Unitary Plan Wind Tunnel (Langley), 72

X-15 program and verification of data from, 

43

winged vehicles and reentry technique

flying, advantages of over falling, 116

HYWARDS program, 166–67

limitations of, 45, 126

preference for over capsules, 92–94, 

113–14, 124, 129, 130, 185, 284

research and writings about, 14, 15

support for designs, 44–45, 92–94, 

113–14, 116, 124, 125, 129, 130, 132, 

186, 284

See also delta-wing concepts; spaceplanes

wing loading, 129, 130, 143, 143n43, 166, 167

wings

deployable, 126

variable-geometry, 126, 152

WLB-1 shape, 138

Wong, Thomas J., 132

Wood, James, 178n132

World War II

hypersonic flight research and, 8–9

jet-assisted (rocket-power-assisted) 

technologies, 5

V-2 missile use, 11, 12

Wright-Patterson Air Force Base (AFB), 133, 

133n21, 158. See also Air Force Flight 

Dynamics Laboratory (AFFDL)

WS-199D research program, 137

X

X-1/XS-1 vehicle, 17, 37, 37n5, 144

X-2 aircraft, 22, 37

X-15 research airplane

ablative thermal protection systems, 158, 

191–92, 196

acceleration requirement, 14

blunt-body reentry design, 35

capabilities of, ix

construction materials, 22, 23n52, 24n53, 

40–41

contributions of to piloted space flight, 42, 

43–44, 43

design of, ix

development and construction of, 39–42, 

40n15

g-loads and human tolerances, 35

heating projections, 40

hot structures, 22, 23n53, 24n53, 26, 

40–42, 168, 191–92, 195–96, 230

hypersonic flight research, 43–44

landing of, ix, 82, 156–57, 179, 182

NB-52 carrier aircraft launch, 144, 146

number of flights flown, 42

reentry and recovery research and, ix

shock waves and X-15 model testing, 41

simulation training and mission planning, 42

Space Shuttle design and, 43–44, 142, 201

speed and altitude records, 42, 42n19

unpowered landing concept, x

X-20 Dyna-Soar. See Dyna-Soar (X-20) program

X-23A (SV-5D) lifting body, x, 140–42, 153, 

238. See also PRIME (Precision Reentry 

Including Maneuvering reEntry) vehicles

324
Index

X-24A (SV-5P) lifting body, 133, 153–55, 154, 

156, 158, 238, 240

X-24B (SV-5P-2) lifting body, 152, 155–59, 157

X-30 program. See National Aero-Space Plane 

(NASP, X-30)

X-33 VentureStar program

cancellation of program, 158n83

costs of research and technological 

developments, 227–230

development and design of, 225–230, 226, 

228, 233–34, 254

funding for, 229

metallic thermal protection system, x, 221, 

230–33

technical problems with, 233–34

X-34 Reusable Small Booster Program, 158n83, 

221, 234–35, 235, 254

X-37/X-37B advanced-flight demonstrator, 

272–75, 274, 276, 277

X-38 Crew Return Vehicle (CRV)

development and design of, 236–241, 237, 

239, 241

funding for, 236, 243–44

nose caps, 242–43

parafoil landing concept, 236, 238–240, 

238n37, 239, 241

purpose of, 236, 237, 238, 243–44

thermal protection system, 238, 242–43

X-40 vehicle, 273

X-43A scramjet vehicles, 158n83

XB-70XB-70A/B-70 bomber, 135, 152n55, 

152n66

XLR11 rocket engines, 145, 150, 153

XLR99 rocket engines, 167

X-planes

naming of, 37

research goals, 37, 37n4, 39, 170

Round One aircraft, 37

Round Three aircraft, 44, 170–71

Round Two aircraft, 39–44

speed records set by, 37

transonic flight research, 37–38, 37n4

Y

Yeager, Charles E., 37

Young, John W., 67, 119

Z

Z-6070 methyltrimethoxysilane, 205

Zima, Bill, 155

zirconia, 177, 194–95

zirconium, 140, 177

zirconium oxide, 71

Zuckert, Eugene M., 179–180, 181

325
F as ro el t yb eh S epu 

I tn re en :t tskoob ro e g.opg. 

aF :x ( 215 )202 

ISB N 978-0-16-091064-7

9 780160 910647

90000

National Aeronautics and Space Administration

Washington, DC

NASA SP-2011-593