Editors: 

Steven C. Fisher

Shamim A. Rahman

John C. Stennis Space Center

REMEMBERING THE GIANTS

APOLLO ROCKET PROPULSION DEVELOPMENT

The NASA History Series

National Aeronautics and Space Administration

NASA History Division

Office of External Relations

Washington, DC

December 2009

NASA SP-2009-4545
Library of Congress Cataloging-in-Publication Data

Remembering the Giants: Apollo Rocket Propulsion Development / editors, Steven C. 

Fisher, Shamim A. Rahman.

 p. cm. -- (The NASA history series)

 Papers from a lecture series held April 25, 2006 at the John C. Stennis Space Center.

 Includes bibliographical references.

1. Saturn Project (U.S.)--Congresses. 2. Saturn launch vehicles--Congresses. 3. Project 

Apollo (U.S.)--Congresses. 4. Rocketry--Research--United States--History--20th century--

Congresses. I. Fisher, Steven C., 1949- II. Rahman, Shamim A., 1963- 

 TL781.5.S3R46 2009

 629.47’52--dc22

 2009054178
Apollo Rocket Propulsion Development 3

Table of Contents

Foreword...............................................................................................................................7

Acknowledgments .................................................................................................................9

Welcome Remarks

Richard Gilbrech ..........................................................................................................11

Steve Fisher...................................................................................................................13

Chapter One - Robert Biggs, Rocketdyne - F-1 Saturn V First Stage Engine.......................15

Chapter Two - Paul Coffman, Rocketdyne - J-2 Saturn V 2nd & 3rd Stage Engine.............27

Chapter Three - Gerald R. Pfeifer, Aerojet - Attitude Control Engines ................................39

Chapter Four - Tim Harmon, Rocketdyne - SE-7 & SE-8 Engines .....................................51

Chapter Five - Clay Boyce, Aerojet - AJ10-137 Apollo Service Module Engine...................59

Chapter Six - Gerard Elverum, TRW - Lunar Descent Engine.............................................73

Chapter Seven - Tim Harmon, Rocketdyne - Lunar Ascent Engine ....................................87

Appendix A - Event Program...............................................................................................97

Appendix B - Speakers’ Group Photograph .......................................................................101

Appendix C - Robert Biggs’ Presentation Viewgraphs........................................................103

Appendix D - Paul Coffman’s Presentation Viewgraphs.....................................................113

Appendix E - G. R. “Jerry” Pfeifer’s Presentation Viewgraphs............................................123

Appendix F - Tim Harmon’s SE-7 & SE-8 Presentation Viewgraphs .................................133

Appendix G - Clay Boyce’s Presentation Viewgraphs.........................................................143

Appendix H - Jerry Elverum’s Presentation Viewgraphs.....................................................151

Appendix I - Tim Harmon’s Lunar Ascent Engine Presentation Viewgraphs......................171

Appendix J - Event Photos and Highlights ........................................................................179

Further Reading ................................................................................................................191

Glossary ............................................................................................................................193

The NASA History Series..................................................................................................197

On April 25, 2006, NASA’s John C. Stennis Space Center hosted a series of lectures on Apollo 

Propulsion development. This monograph is a transcript of the event, held as part of the celebration to mark the 40th anniversary of the first rocket engine test conducted at the site then 

known as the Mississippi Test Facility. On April 23, 1966, engineers tested a cluster of five J-2 

engines that powered the second stage of the Saturn V moon rocket.

This transcript has been edited for readability and clarity. The opinions expressed are solely 

those of the individuals presented. The report does not in any way promulgate policies or 

state the official opinions of the National Aeronautics and Space Administration or the U.S. 

government.

John C. Stennis Space Center

History Office

B-1100

Stennis Space Center, MS 39529

228.688.2646

2006 Event Title “On the Shoulders of Giants,” Apollo Propulsion Development Seminars.

Moderator: Steve Fisher, Rocketdyne

Speakers: Robert Biggs, Paul Coffman, Gerald Pfeifer, Clay Boyce, Gerard Elverum, 

and Tim Harmon

Initial Transcription: Michele Beisler, Transcription/Technical Writing: B. Nicole Wells 

(Jacobs Technology Facility Operating Services Contract)

Editors: Mr. Steve Fisher, P&W Rocketdyne, and Dr. Shamim Rahman, NASA 

Event Coordinator: Rebecca Strecker, NASA

About the Cover –

At left, the Apollo Saturn launch vehicle; middle, a J-2 engine on a static test stand 

operating on oxygen and hydrogen; and at right, lunar surface photographed from 

approaching spacecraft.
6 Remembering the Giants

Dr. Shamim Rahman began his professional career at 

The Aerospace Corporation in El Segundo, CA, in 1985 as an 

aerospace engineer on launch vehicle and spacecraft flight 

programs for the Air Force Space and Missile Command. He 

later joined the TRW Propulsion Research Center to work on 

innovative rocket propulsion devices for civil and military 

space applications. Dr. Rahman joined NASA in 1998. At 

Stennis Space Center (SSC), Dr. Rahman has provided 

technical leadership and oversight on national rocket propulsion test facilities, and over 

a variety of research and development test activities for advanced rocket propulsion 

development. Dr. Rahman assumed his current position in November, 2007, as the 

Deputy Director for the Engineering & Test Directorate at NASA SSC.

Mr. Steve Fisher has spent 37 years at Rocketdyne in 

the design, development, testing, and evaluation of liquid 

rocket engines, and recently retired as a Technical Fellow 

from Pratt & Whitney Rocketdyne. He initially joined North 

American Aviation of El Segundo to work on the original B-1 

Bomber, and later transferred to the original Rocketdyne 

Division of the company. During his tenure, he was integrally 

involved in the development of the Space Shuttle Main Engine 

components at the Coca 1 and Coca 4 test facilities of Rocketdyne, and has further 

contributed to a wide variety of propulsion activities including advanced research and 

development projects with storable and cryogenic propellants, as well as flight engine 

projects such as the RS-68 engine design and development. Mr. Fisher served as a 

Boeing Technical Fellow prior to the Rocketdyne division being sold to Pratt & Whitney.
Apollo Rocket Propulsion Development 7

It gives us great pleasure to provide this historical compendium of what will likely be 

remembered as one of the most remarkable achievements in the evolution of rocket propulsion. 

This achievement was the simultaneous development, testing, and flight use of a series of firstever propulsive devices that delivered Apollo 11 astronauts safely to the surface of the moon 

and back to Earth. These devices helped assure three individuals, Armstrong, Aldrin, and 

Collins a place in the history of humankind.

From the F-1 booster engine to the lunar module ascent engine of the Apollo vehicle stack – 

all built and delivered by the new United States space industrial base – these individual rocket 

propulsion development stories provide a glimpse of how technical ingenuity rose to meet the 

challenge of the race to the moon.

The development histories and lessons learned about the various engines are told by the 

engineers and project managers, and were recorded on DVD so that the lecture series held 

at NASA’s John C. Stennis Space Center near Bay St. Louis, Mississippi, could be replayed 

again and thus live on. Remarkably, to those who attended, it was apparent that these speakers 

recalled their Apollo challenges as if they had happened “just yesterday.” It was clear in their 

voices that the engines carried not just the hardware but also the hope of the nation that this 

“moon shot” could even be done at all.

Although this monograph comes some years after the actual date of the lectures, and describes 

work from decades ago, the lessons will continue to carry space exploration forward. The 

story told within is not how one particular engine was built, but rather how ordinary people 

persisted and were driven to do extraordinary work. The country owes these resourceful and 

dedicated engineers a debt of gratitude for giving us the technical precedents upon which 

today’s space programs rest in a continuing story of human exploration.

It would not have been possible without the sanction and enthusiastic support of NASA 

Stennis Space Center’s “front office” (center director, deputy director, and associate director), 

and the excellent support and facilitation of the local NASA public affairs staff. The 2006 

event was officially designated at NASA SSC as “On the Shoulders of Giants”, and in this 

monograph is more aptly designated by the title, “Remembering the Giants.”

Shamim A. Rahman, PhD,

NASA Stennis Space Center 

Steven C. Fisher,

Technical Fellow, Boeing and 

Pratt & Whitney

Retired

 

 

Foreword
8 Remembering the Giants
Apollo Rocket Propulsion Development 9

The initial transcription of these seminars was done by Michele Beisler and Shelia Reed, based 

upon the excellent videography of Rex Cooksey of CSC who provides Stennis’ media services.

This document then was ably constructed by Nicole Wells of Jacobs Technology in draft form 

to be further edited and taken into publishable form by an excellent media services team. This 

team included Gene Coleman, Shannon Simpson, Courtney Thomas, and Lacy Thompson.

We wish to recognize Angela Lane and the members of the Media Services Department at the 

Stennis Space Center for their assistance with graphics and layouts.

This could not have been published without the full support of the NASA History Office at 

Stennis Space Center, and also NASA’s History Division at Headquarters, and in particular 

Steve Garber who has steered many such documents.

Special thanks also go to Rebecca Strecker, Paul Foerman and Tessa Quave of the NASA Stennis 

Space Center External Affairs Office, who deployed their excellent support team whenever 

needed.

Also at NASA Headquarters, thanks go to the professionals in the Communications Support 

Services Center (CSSC). Printing specialists Tun Hla and Hanta Ralay managed this critical 

final step. CSSC managers Gail Carter-Kane and Cynthia Miller, as well as Michael Crnkovic 

and Thomas Powers in the Information Technology and Communications Division, oversaw 

this project as a whole and made it possible.

Our gratitude to all for their professional and expert contributions.

Acknowledgments
10 Remembering the Giants

Panel of speakers seated from left to right: Boyce, Elverum, Harmon, Pfeifer, Coffman and Biggs.
Apollo Rocket Propulsion Development 11

Welcome Remarks

Dr. Richard Gilbrech

We at NASA Stennis Space Center are honored to host On the Shoulders of Giants. We have 

a big mission ahead of us, as did those who worked at the start of the Mercury, Gemini, and 

Apollo days. We really appreciate the participants coming to Stennis and sharing their wealth 

of experience with us. Thank you to everyone else attending the event: our friends from Marshall Space Flight Center, Johnson Space Center, and Kennedy Space Center. I have always 

been in awe of the Mercury, Gemini and Apollo programs. That is really what got me hooked 

into NASA: the accomplishments that were made in that era, given the state of modeling and 

computers compared to what we have today. It has always fascinated me. I always felt like I 

had missed probably one of the most exciting times in engineering history with that feat. My 

generation got the Space Shuttle Program, which has been exciting, but I always had the “bug” 

to be part of a “moon shot.” I’m very excited that we are getting that chance now. I had the 

opportunity to accompany Gen. Tom Stafford on a tour of one of the Stennis Space Center 

test stands. Of course, he had flown two Gemini missions and a Saturn V mission on Apollo 

10, and then, later, the Apollo/Soyuz Test Program. I said to him, “I really felt like I missed 

out.” He said, “You are in as exciting a time as I had, and your time will come also.” He also 

commented on the Gemini ride versus the Apollo ride. He said, “The Apollo rides were like a 

Cadillac compared to the roll maneuvers they did on the Gemini.” It is our turn now. I’m very 

excited.

We do have a big challenge ahead of us. There are tremendous technical challenges, and one 

thing we don’t want to do is repeat mistakes of the past. That is part of what the On the Shoulders of Giants program is about: to listen to the wealth of knowledge from these panelists. They 

have forgotten more about stage testing than I’ll ever know. I’m happy to have them participate 

in this event. Stennis Space Center is poised to play its role in this. We hope to see those clusters of engines hanging on stages and upper stage tests, and then, eventually, to see that cargo 

lifter we hope will be plugged into the B-2 Stand. We can maybe break some windows again.

One of my mentors, Roy Estess1

, always used to pick on the hardware guys. He’d say, “Stennis 

has a long history of testing rocket engines and we’ve even got a longer history of waiting on 

delivery of rocket engine hardware.” I know there is difficulty on the development side, and we 

appreciate all the challenges that the hardware developers are up against. One thing we want 

to do is make sure we are poised to get the kind of test data needed to make sure all the test 

lessons we learned in the Apollo Program are clear and present with our engineers today, so the 

developers get the data they need to make the decisions they are going to be facing. 

1

 Roy Estess served as SSC Center Director from 1989 to 2003.
12 Remembering the Giants

S-1C Stage Installation into Hot Fire Stand at Mississippi Test Facility (MTF) - 1967 

(NASA image number: GPN-2000-000559)
Apollo Rocket Propulsion Development 13

I am really happy to participate in the On the Shoulders of Giants series, especially among the 

six Apollo experts who will speak during the event. It kind of makes me feel like a kid, which 

is neat. These guys were doing stuff back when I was watching the result on TV. So, this is 

really exciting for me. It was interesting to pull the lecture series together. It was done before 

at a Joint Propulsion Conference2

, and we wanted to do it again during summer or fall 2005, 

but different things prevented that from happening. If you ever had to coordinate with four 

different companies, four different management groups, and expert people from those four 

companies, you know what kind of interests that raises. Well, you ought to try doing that with 

six retired folks. These people are not really retired; these people are active. They are all out 

doing stuff, and they often tell me, “Hey, we have things to do. We just can’t drop what we 

are doing and show up for a conference.” I really want to thank them for participating in this 

event. It really was a challenge to coordinate and get them all to Stennis Space Center together, 

and I appreciate them all showing up. 

Going to the moon was a monumental undertaking. There’s no doubt about that. It required 

seven different rocket propulsion systems to get the vehicle there and back again. The participants in this lecture series discussed each one of those systems. These were the folks who actually made that dream come true. The series of presentations was very similar to what was shown 

and what was presented at the 2004 Joint Propulsion Conference during the 35th anniversary 

of the Apollo 11 first landing on the moon. What was quite interesting to me was that while 

a lot of these veterans had heard each other’s names mentioned and knew of them, some of 

these folks had never met before that conference. I thought it was quite interesting to see the 

interaction of these folks as they finally got together after all these years.

Steve Fisher

2

 2004 Joint Propulsion Conference held at Fort Lauderdale, Florida, sponsored by the American Institute of 

Aeronautics and Astronautics (AIAA).
14 Remembering the Giants

F-1 Engine (NASA image number: MSFC-6413912)
Apollo Rocket Propulsion Development 15

Robert “Bob” Biggs worked forty-seven years at 

Rocketdyne, and spent nine years as lead development 

engineer and development project engineer on the F-1 

Engine Program. He spent several months on the Navaho 

cruise missile project, three years as lead engineer in the 

Jupiter Program performance analysis group, and a year as 

manager of the Dynamic Analysis Laboratory. Biggs also 

worked thirty-four years on the space shuttle main engine, serving as development manager 

and chief project engineer.

Before I go into the history of F-1, I want to discuss the F-1 engine’s role in putting man on 

the moon. The F-1 engine was used in a cluster of five on the first stage, and that was the only 

power during the first stage. It took the Apollo launch vehicle, which was 363 feet tall and 

weighed six million pounds, and threw it downrange fifty miles, threw it up to forty miles of 

altitude, at Mach 7. It took two and one-half minutes to do that and, in the process, burned four 

and one-half million pounds of propellant, a pretty sizable task. (See Slide 2, Appendix C)

My history goes back to the same year I started working at Rocketdyne. That’s where the F-1 

had its beginning, back early in 1957. In 1957, there was no space program. Rocketdyne was 

busy working overtime and extra days designing, developing, and producing rocket engines for 

weapons of mass destruction, not for scientific reasons. The Air Force contracted Rocketdyne 

to study how to make a rocket engine that had a million pounds of thrust. The highest thing 

going at the time had 150,000 pounds of thrust. Rocketdyne’s thought was the new engine 

might be needed for a ballistic missile, not that it was going to go on a moon shot.

Chapter One

Robert Biggs

Rocketdyne - F-1 Saturn V First Stage Engine
Chapter One

16 Remembering the Giants

In July 1958, we began the International Geophysical Year, wherein the scientists of the 

world were going to plan and execute experiments to learn more about the earth. It was an 

eighteen-month program. As part of that, the United States planned to launch an artificial 

satellite to circle the earth, and they were going to do it within the geophysical year. President 

Dwight Eisenhower intervened with one portion of that. He had a fear of what he called 

the “military industrial complex” becoming too powerful. He wanted the space program, the 

satellite program, to be done without benefit of any weapons system or, in fact, weapons 

system personnel; so, all of the people who knew best how to make rockets go up weren’t 

allowed to work on the project. They had to get other people, “civilians” as they referred to 

them. That’s what happened to bring the Vanguard Program into being outside of what we 

had; the Jupiter, the Thor, the Atlas, the Redstone, the Navaho, all of these were already flying 

and almost operational. This was happening at a time of decreasing national prestige due to a 

number of events.

The first one was in October 1957. The world and the United States were shocked by the Soviets 

announcing they had put up a satellite, Sputnik I. It was a satellite weighing approximately 

183 pounds. It was shocking to us that they could do it before the United States did. Before 

we recovered from that shock – it was within a month – they launched Sputnik II; it had a 

dog as a passenger. Sputnik II also gave the United States information on how well the Soviets 

were doing. The third stage did not separate from the orbiting package, and the third stage 

alone weighed 16,000 pounds. The Soviets had managed to put 16,000 pounds into orbit! Our 

Vanguard satellite was going to be eighteen pounds, by the way. From what was put into orbit, 

the United States was able to calculate that the Soviet booster must be three times as powerful 

as the most powerful booster we had, which was the Atlas cluster of three engines. That gave 

emphasis for starting to look at a much higher thrust engine. During this time period, an 

organization called the Advanced Research Projects Agency was formed within the Department 

of Defense to manage the new technology for all of the services. The Air Force, Army, Marines, 

and Navy were all developing the same technology. The new agency was designed to prevent 

that; it was going to develop the new technology for space. (See Slide 3, Appendix C)

The first attempt at launching the Vanguard satellite was December 6, 1957. It went up four 

feet; it blew up four feet off the launch pad. By this time, Eisenhower had relented and allowed 

the Army Ballistic Missile Agency (ABMA) to step in and put up a satellite. They actually used 

a Redstone vehicle with a Redstone engine, and they called it a Jupiter-C. Dr. Wernher Von 

Braun, who headed up the ABMA, let it be known that he could put one up in two months, 

and they finally turned him loose.1

 Very shortly after that, Explorer I was in orbit. NASA 

1

 For more details on this, see page 111, of Dr. Space, The Life of Wernher von Braun, by Bob Ward, Naval Inst. 

Press, Md, 2005.
Robert Biggs, Rocketdyne - F-1 Saturn V First Stage Engine

Apollo Rocket Propulsion Development 17

absorbed ABMA in March 1960 and took over all of the non-military space effort. In May 

1960, the U-2 spy plane, which was supposed to be able to fly above Russia without harm 

coming to it, couldn’t be reached. The Union of Soviet Socialist Republics had shot the plane 

down. The next shock came when Yuri Gagarin was put into space; a human was in space for 

the first time on April 12, 1961. That was followed five days later by the Bay of Pigs fiasco. 

John Kennedy had been president for just a few months and had allowed himself to agree to 

a CIA-coordinated invasion of Cuba under the assumption that all Cubans would rise up and 

fight against the dictatorship of Fidel Castro, and that didn’t happen.

In May 1961, Alan Shepard completed a fifteen-minute suborbital flight test as the first 

American in space. Based on the results of that, John Kennedy made his famous speech about 

sending man to the moon and returning them safely. He waited until we had success in space 

before making the announcement. The success was this small fifteen-minute suborbital flight. 

On May 25, 1961, that speech was made, the same day the first test was performed on the F-1 

engine. (See Slide 4, Appendix C)

The F-1 mission was to provide seven and one-half million pounds of thrust for the Saturn V 

first stage. It was based on, except for size, pretty much the same technology being used for 

the ballistic missiles. The difference was that it was a lot bigger, and the biggest challenge was 

just that – its bigness. It was ten times the thrust of the biggest current production engine, 

which was 150,000 lbf (pounds-force). This was going to a million and one-half. Upratings 

(upgrades in power, size, etc.), traditionally, even in other things besides rocket engines, are 

reaching if they go 10 percent. We had several 10 percent upratings before, but nothing like 

this. The engine was eighteen feet tall, twelve feet wide. It had a thrust level that started out at 

1.5 million pounds force, plus or minus 3 percent. They made a change to it and called it an 

uprating by going to 1,522,000, plus or minus 1.5 percent. If you do the math, all that did 

was move the million up to the middle or the high part of the tolerance. The engine essentially 

had to be designed for the same pressure anyway. The specific impulse (a measure of engine 

performance efficiency) was 265 seconds, which was good for those days. It looks not so good 

today. (See Slide 5, Appendix C)

Combustion chamber pressure was another challenge. This was designed at 1,100 pounds per 

square inch (psi). The highest chamber pressure of the day was 520 psi, so this was a doubling 

of the chamber pressure and significantly more thrust. The engine burned liquid oxygen and 

Rocket Propellant-12

 (RP-1, a highly refined form of kerosene) at a mixture ratio3 of 2.27, and 

2

 Rocket Propellant-1 (RP-1), highly refined form of kerosene.

3

 Mixture Ratio - Ratio of mass flowrate of oxidizer (liquid oxygen) to mass flowrate of fuel (RP-1).

The F-1 mission was 

to provide seven 

and one-half million 

pounds of thrust for 

the Saturn V first 

stage. It was based 

on, except for size, 

pretty much the 

same technology 

being used for the 

ballistic missiles.
Chapter One

18 Remembering the Giants

it had a mission duration of 165 seconds. Well, it started out at 150 seconds, and the first two 

Mercury flights were 150 seconds, but they were unmanned. They had the whole stack, but 

they didn’t have the astronauts. They made flight changes for the first manned one, and from 

that point on, the burn duration was 165. It was designed for a qualification life of twenty 

starts and 2,250 seconds with a weight of 18,000 pounds. We later ran this engine at 1.8 

million pounds of thrust4

, which gave it a thrust-to-weight ratio of 100, but it was pretty close 

to 100 anyway, and if you stripped off the stuff that was not providing thrust, the remainder 

of the engine would have been a 100:1 thrust-to-weight ratio. The major features are that it 

had a single turbopump, and it had a shaft that was parallel to the axis of the thrust chamber. 

It had a liquid oxygen (LOX) pump on top of the shaft, then a fuel pump, then a turbine. 

It took the single turbine to run the two pumps. The nozzle had an area ratio of 16:1. It was 

tube-wall down to the 10:1 expansion, which is where the turbine exhaust was put into the 

nozzle through a skirt extension. The tubes went down to 10:1, and from then on, it was a 

double-wall, hot, gas-cooled nozzle extension. In the turbopumps, both the fuel pump and 

the liquid oxygen (LOX) pump had a double discharge, a discharge on each side of the pump. 

The reason for doing that was to prevent the entire head rise across the pump from becoming 

a delta pressure radially for a fixed load. It divided the total load by two, having two different 

directions.

How was the RP-1 fuel used for purposes other than just burning? These were unique things 

at the time. For instance, we used the fuel for lubricating turbopump bearing. Prior engines 

and, in fact, the original F-1 design had a separate lubrication pumping system to lubricate 

bearings. It also was used as hydraulic power instead of using hydraulic fluid. It was used to 

power the engine valves and the thrust vector control actuators, and in addition, it did not 

require any auxiliary starting power. (See Slides 6 and 7, Appendix C)

The F-1 turned out to be a very simple engine, but it started out with its initial design as a very 

complex engine. The first design, which went through a design review and was approved, had 

three turbopumps on it. It had a hydrazine pump, which required a hydrazine gas generator 

to run the turbine on the hydrazine pump. The hydrazine gas generator would run the LOX 

pump and turbine and the fuel pump and turbine. It had a separate pump for the LOX pump, 

and a separate one for the fuel, and a separate one for hydrazine. There were three pumps, each 

with a shaft and each with a turbine, and that was replaced during development. Actually, it 

never made it to the first engine test. It was replaced with a single turbopump. It also had the 

lubrication oil system for the bearing that was eliminated. 

4

 F-1A was rated at 1.8 million pounds force of thrust.
Robert Biggs, Rocketdyne - F-1 Saturn V First Stage Engine

Apollo Rocket Propulsion Development 19

There was one design that was really troublesome, and we did not get rid of it before the first 

test. We had to live with it for some period of time, and it was called a triple manifold thrust 

chamber. Picture three doughnuts stacked on top of each other; those were fuel manifolds 

at the top end of the thrust chamber. The fuel for the thrust chamber went into the middle 

manifold, then went down the tubes to cool the tubes. It went down half the tubes, then back 

through the other half of the tubes to another manifold. There were four trombone tubes that 

ran from the bottom manifold to the top manifold in order to get the fuel from that end of 

the injector. They had tangential inlets and outlets. It set up a racetrack in both the bottom 

manifold and the top manifold, and the pressure loss was horrendous, but we managed to 

struggle by for a while until we could get a better design. (See Slide 8, Appendix C)

On the F-1 engine schematic, the turbopump was in the upper left, with green for oxidizer and 

red for fuel. Since the turbopumps had two discharges, they required two fuel lines and two 

LOX lines going to a valve. There were two LOX valves, one on each side of the engine, and 

two fuel valves, one on each side of the engine as well. The hypergol cartridge in the middle of 

the control system held a fluid that would automatically ignite as it got into the chamber. It 

didn’t need a pyrotechnic device to ignite it. (See Slide 9, Appendix C)

Many people have wondered how the start sequence looked. It took about five seconds for 

the engine to start under what is known as a tank head start5

. It had no help from outside. It 

started by opening the engine control valve, which was a four-way hydraulic valve that ported 

hydraulic fluid to open the main LOX valve. The hydraulic system on this engine was quite 

robust. All of the valve actuators were way overpowered. There was no concern of hydraulic 

contamination with the engine. By opening the LOX valves first, the LOX started flowing 

through the pump in such a way that it treated the LOX impeller as a turbine, and it started 

to turn the pump. It would turn the pump up to 700 to 1,000 revolutions per minute, which 

got rid of the breakaway torque concerns and started the engine going up in power. Right after 

that, the gas generator valves opened. The gas generator had mechanically-linked valves, one for 

the fuel and one for the LOX, so they both opened at the same time. The gas generator would 

ignite at a very low pressure and then start building up. It would take two or three seconds for 

the pressure to get up to the point where the igniter fuel valve would open, allowing the fuel 

to go into the chamber and start the thrust chamber ignition. As soon as the thrust chamber 

ignition was sensed by another valve, by just measuring the fluid pressure, the main fuel valves 

opened, and the thrust was allowed to come on up. At about 80 percent, the thrust laid over 

5

 Tank head refers to the pressure at the pump inlet due to the weight of propellant in the tanks located above the 

engine.
Chapter One

20 Remembering the Giants

for a little while and then went on up at a different ramp (i.e. rate of increase). The reason was 

that the thrust chamber was primed with ethylene glycol, so the initial fuel burning during this 

engine start transient period was an ethylene glycol and water mixture. It didn’t provide as high 

a chamber pressure as the full RP-1 fuel so, when that ran out, the chamber pressure kept going 

on up and reached the 100 percent value. (See Slide 10, Appendix C)

There was a qualification test requirement to take two engines off the flight line. Supposedly, 

the concept was to grab two engines at random and put them through a qualification test 

series. One of those series was life demonstration, which demonstrated the twenty tests and 

the 2,250 seconds. The other one did that and also demonstrated a lot of simulated engine 

malfunctions and different environmental tests. These two series were quite successful and 

completed the contractual requirements for qualifications testing, but they were not the end 

of running a lot of seconds on engines. This was a supposedly expendable throwaway engine 

with a mission duration of 165 seconds, but we actually tested six engines to an excess of 5,000 

seconds. The only reason this was a non-reusable engine was it was non-recoverable. If we 

could have recovered it after a flight, it would have reusable. (See Slide 11, Appendix C)

We adopted the idea of acceptance testing, using the concept of an “all up” test at every 

opportunity – test the whole thing. The concept was to do that with the full stage, so, when 

we had a series of production engines, they first would have to be acceptance tested on the 

engine test stand. That required two tests of duration – a forty-second calibration test and a 

165-second mission duration test on every engine. After that, the engines were grouped in 

clusters of five and put into an S-IC stage, the actual flight stage. They then were tested for 125 

seconds, which was the maximum duration they could get without having some acceleration. 

The first three stages were tested at NASA’s Marshall Space Flight Center in Huntsville, 

Alabama. The rest of them were tested at NASA’s John C. Stennis Space Center near Bay St. 

Louis, Mississippi, on the B-2 Test Stand. The total time required for the acceptance test was 

495 seconds – for a mission duration of 165 seconds. All of that added up, but the highest time 

any engine achieved was 800 seconds. (See Slide 12, Appendix C)

There were a couple of development problems with the F-1. The first one was the most famous 

one. I think many people who know nothing much about the F-1 know that it had a combustion 

instability problem. It was thought before the first engine test – a couple of years before – that 

just by making it this much bigger, it would not be stable. It would have to be unstable. We 

ran some tests on a device called “King Kong,” which was a solid-wall, steel chamber, and the 

tests were run for a very short duration. The test stand could only go through a start, then 
Robert Biggs, Rocketdyne - F-1 Saturn V First Stage Engine

Apollo Rocket Propulsion Development 21

maybe half a second of main stage firing and then a shutdown before it ran out of propellant. 

Of course, this was all pressure-fed, a pretty difficult thing to arrange. The first few tests had 

some instability, which never was understood properly, because with the demonstration tests, 

everything was stable. That allowed us to make the conclusion that it probably was stable, and 

to go on with the plan. (See Slide 13, Appendix C)

In the first year of testing – because of other problems – we were unable to operate at rated 

thrust (the performance level an engine is designed to achieve). That’s not unusual. In fact, 

it’s more usual than not that the first engine in a new design is unable to achieve rated thrust 

for some reason. The maximum thrust the F-1 could hit for the first engine was 1 million 

pounds; so, for the first year, testing was limited to 1 million pounds. During that year, we got 

some spontaneous combustion instability strikes, but only seven times. That’s a rate of about 

10 percent. The tests we ran6

 would not have been able to predict whether it would have a 

random instability with some probability; it would only detect if it was always unstable. Well, 

this was a low probability, but not nearly low enough to be used. During the year, we would 

have instability detected by accelerometers on the thrust chamber; then, the engine would be 

shut down by a thing they called the rough combustion cutoff device. It was a nuisance simply 

because we were losing tests. It wasn’t doing much damage. I guess it wasn’t doing any damage 

because we were able to go test again. 

Then, in the second year, we were able to run rated thrust for the first time. The thrust level went 

up to 1.5 million, and the testing continued. Then, we got a spontaneous instability that was 

so severe, it destroyed the engine. The initial shock was bad enough to cause both of the dual 

fuel lines to rip off the engine, so the engine was running with LOX and no fuel. It was burned 

up. It was a total loss. We had eleven such episodes in that time period. All of them did some 

damage. Only three of them blew the ducts off and totally destroyed the engine, but the rest 

of them did a considerable amount of damage. As soon as that first one hit that destroyed the 

engine, the problem of instability got more attention than it had previously. “Project Go” was 

formed, headed up by Jerry Thompson at Marshall Space Flight Center and Paul Castenholz at 

Rocketdyne with Bob Levine, Dan Klute, and Bob Fontaine. They made some changes to the 

injector. The injector had baffles sticking down, dividing it into thirteen compartments. We 

had fifteen different arrangements of baffles that were tested. There were also fourteen different 

injector configurations behind that. All of those were tested, and, eventually, the instability 

went away. We never again had a spontaneous instability. After a while, in the third year, we 

were not getting any spontaneous instabilities. However, we were getting forced instabilities 

6

 Tests with King Kong, the solid-walled steel chamber.

. . . in order to 

prove the injector 

was a stable 

design, we had 

to put a bomb

in it, detonate 

the bomb, and, 

then, damp it out 

within forty-five 

milliseconds. 
Chapter One

22 Remembering the Giants

because, in order to prove the injector was a stable design, we had to put a bomb7 in it, detonate 

the bomb, and, then, damp it out within forty-five milliseconds. Even though the injector was 

not spontaneously going unstable, it was still a good requirement to make sure that if you ever 

did that, it would damp out immediately. In that third year, we got nine cases where we were 

bombing it to prove that it was stable, and it went unstable. However, a lot of the other tests 

were successful, and at the end of the testing, the configuration was smooth running. It then 

went through a complete qualification test8

 and a lot of bomb tests, and never again was there 

any instability. That problem was solved considerably before the first flight.

We also had another problem that was kind of interesting because it had a catastrophic effect. 

The LOX pump had six vanes in it. The vanes were weak. I don’t want to say under-designed, 

but it was weak for that environment. I think the basic problem was they were scaled up from 

the smaller ones, and you never really had to worry much about the forces across the vane as 

much as this size did. It made it worse. We had failures of that vane four different times. These 

were four engines that were destroyed by vane failure in the LOX pump. It occurred at 110 

seconds, 110.5 seconds, 107.7 seconds and, then, at 109 seconds. Statistically, this looked 

extremely significant. There was something going on that was a function of time, and at 110 

seconds we had a problem. We did an exhaustive investigation twice and could not find any 

justification for that and concluded it was just a freak coincidence. We made some changes to 

eliminate fretting (wear and corrosion damage). What we really did was to set a limit at 3,500 

seconds for the impeller for ground testing. All the flight engines never got up above 800 

seconds. The last engine that went up in 1965 had 5,000 seconds on the impeller. From that, 

we concluded you could set the limit at 3,500 seconds and be okay. One thing about this was 

that on every flight, the people who were familiar with the F-1 program would notice when 

110 seconds went by and, then, breathe again. (See Slide 14, Appendix C)

One significant milestone in the F-1 program was moving from the first test to the first rated 

duration and thrust test. It took one year to run a test that was rated thrust and full duration, 

and I think that’s better than a lot of programs were, including some recent ones. The first 

flight was November 1967. I was in Firing Room 2 at the window for that flight. I considered 

that the highpoint of my career. I’ve never had such a feeling of pride, just watching the thing 

go up and imagining that the earth was trembling. The glass window in front of me was 

moving inches. It was quite an experience. (See Slide 15, Appendix C)

7

 A small charge to provide a pressure impulse in the combustion chamber to test for combustion stability.

8

 Qualification testing for rocket engines.
Robert Biggs, Rocketdyne - F-1 Saturn V First Stage Engine

Apollo Rocket Propulsion Development 23

We got over 280,000 seconds of total burn time throughout the entire program. There were 

twelve Apollo flights that used the F-1 engines, then the Skylab used the last one that flew. 

That added up to thirteen flights or sixty-five total engines in flight. For those flights, those 

sixty-five engines were 100 percent reliable. In addition to that, we had a formal reliability 

demonstration program where we pre-declared the firing of the ground test engine would be 

one that counts and, then, went 336 total equivalent tests without a failure, all pre-declared. 

The engine was quite successful in that regard. (See Slide 16, Appendix C)

At times, I’ve been asked, “What was it like to work on such a program?” Well, it was great. It 

was the most important thing going on in the country, and we were in the middle of it. A lot 

can be said about it. 

I want to share one other aspect of it. At the time we were doing ballistic missile development 

in 1956, the total employment at Rocketdyne was 5,000. At the peak of production of these 

missiles, the employment was 14,000. When Kennedy’s “Man on the Moon” speech was made, 

there were 11,000 employees. That got a fast buildup to 20,000 in 1965. At the time of the first 

unmanned Saturn V flight, early in the program, employment was down to 14,500. When the 

first moon landing took place, we were down to 9,000, and when the last moon landing took 

place, we were down 2,500; 17,500 people were laid off in a few years. 

The total work force on Apollo peaked at about 400,000, and I imagine the whole program 

went down about the same amount. (See Slide 17, Appendix C)

---------------------------------------

Editor’s Note: The following information reflects a question-and-answer session held after Biggs’ 

presentation. 

STEVE FISHER9

: Just one comment I wanted to make while Mr. Biggs was talking about work 

force population and stability in the F-1 Program. Several years ago, I had the opportunity or 

the task to create a briefing on F-1 stability. While I was doing that, somebody lent me an old 

Rocketdyne phone book. Within the Combustion Devices Group - there were several different 

groups: main injector, gas generator, gas generator injector, main chamber, etc. - there was a 

group called Main Injector Stability. All they did was stability, nothing about performance, 

9

 Steve Fisher served as facilitator during the On the Shoulders of Giants seminar series.

It was the most 

important thing 

going on in the 

country, 

and we were in 

the middle of it.
Chapter One

24 Remembering the Giants

fabrication, whatever. Under Main Injector Stability, there were thirty-five names in the phone 

book that year. 

QUESTION: Back in 1961 to 1965 when you were working with RP fuels, what was the state 

of the liquid hydrogen technology?

BIGGS: The J-2 (engine project) was developing hydrogen technology.

QUESTION: In 1963, the plan was changed to do “all-up” engine testing or “all-up” stage 

testing. Did that have any effects on the F-1 Program? Are you aware of any changes to the 

qualifications/certification program?

BIGGS: Well, I can say the “all-up” testing requirement probably saved us from a flight disaster 

because of a stage test that found a problem. The acceptance test of S-IC 11 had a large fuel 

leak caused by an incorrect installation. The hydraulic fluid was being supplied from an engine 

to the hydraulic actuators through a line that was installed on the engine by Boeing at the 

stage level. It went through Rocketdyne with a test plate, a good test plate cover. At that time, 

the line was installed for the stage test, and that would be the same as it would be for a flight, 

supposedly. Well, the mechanic installed a line over a cover that was left on the port. It caused a 

fuel leak during the test, which caught on fire. It led to other things on that test, which resulted 

in scrapping two engines. It would have been bad in flight. I can’t say how bad because there 

were other things wrong on the test, but certainly, that was a condition only an “all-up” test 

would have found.

QUESTION: I’m Pete Rodriguez from Marshall Space Flight Center, and we’re doing an 

assessment of what kind of testing stands and capabilities we need to have available. What kind 

of damage happened to the actual test stands during the instability failures, and how long did 

it take to fix and to get back up and running?

BIGGS: Engine test stands, in general, are quite robust because if they fail and come apart, 

that energy is lost in destroying the engine. The test stand can be damaged by fire, but I think 

most of the damage cases were wiring and similar items. In some cases, a lot of damage can be 

done. But in the instability of the F-1, the worst it did was to blow off the propellant supply 

lines. That was pretty bad, but it kept the additional damage local to the engine.

FISHER: I think typically, there is a lot less damage in an engine stand than in a highpressure stand, as you can imagine. Obviously a lot of repairs would need to be done, but most 

infrastructure remains. 
Robert Biggs, Rocketdyne - F-1 Saturn V First Stage Engine

Apollo Rocket Propulsion Development 25
26 Remembering the Giants

One of the J-2 engines that power the upper stages of Saturn V liquid 

hydrogen, on its way from the fuel turbopump, is used to cool the walls of the 

thrust chamber regeneratively. (Source: http://history.nasa.gov/SP-350 - Apollo 

Expeditions to the Moon”, published 1975)
Apollo Rocket Propulsion Development 27

Paul Coffman earned experience during many Apollo-era 

assignments. He served as lead engineer on the J-2 thrust 

chamber assembly development, supervisor of engineering 

test for J-2 components and engines, and manager of 

J-2 engine development and flight support. Following the 

Apollo era, some of Coffman’s program management and 

business development assignments included serving as 

director of controls and engineering, program manager of the compact steam generator, 

engineer coordination for the space shuttle main engine, and project engineer on the gas 

dynamic laser. He also led successful proposal efforts on an advanced storable propellant 

spacecraft engine and a system test bed for potential space station propulsion.

The J-2 engine was unique in many respects. Technology was not nearly as well-developed in 

oxygen/hydrogen engines at the start of the J-2 project. As a result, it experienced a number of 

“teething” problems. It was used in two stages on the Saturn V vehicle in the Apollo Program, 

as well as on the later Skylab and Apollo/Soyuz programs. In the Apollo Program, it was used on 

the S-II stage, which was the second stage of the Saturn V vehicle. There were five J-2 engines at 

the back end of the S-II Stage. In the S-IV-B stage, it was a single engine, but that single engine 

had to restart. The Apollo mission called for the entire vehicle to reach orbital velocity in low 

Earth orbit after the first firing of the Saturn-IV-B stage and, subsequently, to fire a second 

time to go on to the moon. The engine had to be man-rated (worthy of transporting humans). 

It had to have a high thrust rate and performance associated with oxygen/hydrogen engines, 

although there were some compromises there. It had to gimbal for thrust vector control. It was 

an open-cycle gas generator engine delivering up to 230,000 pounds of thrust. 

Chapter Two

Paul Coffman

Rocketdyne - J-2 Saturn V 2nd & 3rd Stage Engine
Chapter Two

28 Remembering the Giants

We delivered 152 production engines. The specific impulse of 425 seconds did represent a 

small compromise. Chamber pressure was relatively modest, a little over 700 pounds per square 

inch absolute (psia), and the mixture ratio was 5.5:1 at that chamber pressure. It also had the 

capability of operating at a mixture ratio 4.5:1. Basic engine weight was relatively light, and if 

you make the numbers, you will come out a little bit over 80 percent for thrust-to-weight ratio. 

Propellants were liquid oxygen and liquid hydrogen. Nozzle area ratio was 27.5:1. Because of 

the difference in densities, there was a requirement for two different turbopumps. They were 

separate and located, as conventional wisdom would have it, on opposite sides of the engines 

just for safety sake. The bearing lubrications were liquid oxygen and liquid hydrogen, so there 

was no separate lubrication system. The gas generator burned main propellants. (See Slides 2 

and 3, Appendix D)

As far as the engine operating in the main stage, there was a gas generator that was fed with 

fuel and oxygen off the main propellants ducts. The gas generator drove the turbomachinery, 

in a series turbine arrangement, through the fuel turbopump and then over into the oxidizer 

turbopump, with a bypass for calibration and a heat exchanger to heat up oxygen for tank 

pressurization (or helium in some instances), before dumping into the thrust chamber at the 

nozzle midpoint. It was a fully tubular thrust chamber with a fuel inlet at the reduced epsilon1

portion of the nozzle. It was one tube down for every two tubes up. There was a 2:1 split, 

and we used the opening at the 2:1 split to dump the hot gas into the nozzle of the thrust 

chamber. The oxidizer turbopump was a fairly conventional centrifugal device. There was a 

rather strange shape for the fuel turbopump. The fuel turbopump was an axial machine. It 

was derived from machines that had been used in the nuclear rocket programs. While that was 

deemed a sound idea at the start, it resulted in probably the single major development problem 

– a start sensitivity for the engine. (See Slide 4, Appendix D)

Again, the engine had to restart. It started at altitude on the S-II and S-IV-B first burn and, 

then, on the S-IV-B restart after one and one-half to two and one-half orbits. The energy for 

start was supplied by cold hydrogen that had been loaded either on the ground or reloaded 

from the main fuel injection manifold during the initial burn operation, and it was this start 

tank that would discharge cold hydrogen through the two turbines to get the engine started. 

Inside that start tank was a small helium tank because the engine carried its entire helium 

supply with it in a cold helium tank.

1

 Epsilon refers to nozzle area ratio, exit area to throat area.
Paul Coffman, Rocketdyne - J-2 Saturn V 2nd & 3rd Stage Engine

 Apollo Rocket Propulsion Development 29

Altogether, 152 engines were produced. The engine development, while it may have ceased 

with the qualification efforts, continued with the F-1 and persisted throughout the entire 

program. It also spawned a J-2S engine and the Linear Aerospike Engine test beds that were 

operated at Rocketdyne as well. From the start in 1960, the first 250-second duration test 

occurred a couple of years later, not a year as the F-1 managed. The 500-second full duration 

test happened another year after that. We then began a series of tests to demonstrate formally 

the engine readiness through preliminary flight rating tests (PFRTs). Those were followed by 

flight readiness tests (FRTs), followed by qualification tests I (qual). Those tests were on the 

225,000-pound version of the engine, fifty-nine of which were delivered. Qual II tests came 

in 1966. Those overlaid with the first production engine delivery, the first flight of the S-I-B, 

and the first Saturn V flight in 1967, with the moon landing coming in 1969. At the time 

we landed on the moon, we were very close to delivering the last of the J-2 engines. Perhaps 

interestingly enough, from a contract standpoint, the engine deliveries were incentivized in the 

initial contract, so there was a bounty on delivering them on time, even though the capsulerelated delays that were experienced by the Apollo Program caused them to be delivered to 

warehouses in many cases. (See Slides 5 and 6, Appendix D)

We had thirty-eight development engines. This was not a hardware-poor program. This was 

a schedule-driven program; cost was no object. We had 1,700 tests through qualification. 

That was the qual II series. We continued testing and had over 3,000 tests by the time the 

program ended. We had multiple test facilities and a lot of component facilities as well. We 

had five engine test stands at Rocketdyne Santa Susana in California, including one that was 

a simulated altitude facility. We had test facilities at NASA’s Marshall Space Flight Center in 

Alabama. We had test facilities at NASA’s Stennis Space Center in Mississippi, and one at 

Sacramento, California, for the S-IV-B. We also did some very significant testing at the J-4 cell 

at the Arnold Engineering Development Center (AEDC) in Tennessee. That was significant 

because that was where we finally resolved the start problems that were the major issue with 

the engine. (See Slide 7, Appendix D)

The engine was required to complete thirty tests and 3,750 seconds of operation for the formal 

demonstrations. If you added up all of these numbers, it came to something over forty tests 

because there were more than one at a time. Indeed, we managed to do that to the point that 

some of them were just start-stop tests as we actually finished the program. The most significant 

ones were what we blandly called “the safety limits tests.” Those were the extremes of the boxes 

for start energy and build up energy. The actual demonstration engine was Engine 2072, which 

did all but the last altitude simulation test. The program office was very, very conscious of, and 

was not comfortable with, the fact that we ran 57.4 seconds over the requirement.

This was a 

schedule-driven 

program; 

cost was no 

object.
Chapter Two

30 Remembering the Giants

There were four key development issues. (See Slide 8, Appendix D) The first was thrust 

chamber ignition detection. It was a good idea to have the thrust chamber lit before the main 

propellants were added. The thrust chamber was ignited by an augmented spark igniter in the 

middle with two spark plugs that used fuel and hydrogen and oxygen. It also had a spot for an 

ignition detection device. The development of ignition detection devices was quite significant 

in that the simplest device was a fusible link. Burning the link showed there was a fire. This 

process was used for most ground tests, except at AEDC, an Air Force-owned test site, where 

we wanted to run multiple tests within a given air-on period, or vacuum period. At this point, 

it seemed advisable to have an ignition detection device for all tests subsequent to the first one. 

There was also an idea that we would like to detect ignition on flight engines as well, since a 

reusable probe underwent a lot of development. That was used for the AEDC engine testing, 

but it never demonstrated enough reliability for flight. The net result was that I don’t believe 

there was ever a ground test conducted without an ignition detection device, but none of the 

engines were flown with one. That’s a point to consider as we move forward to a new crew launch 

vehicle2

. (See Slide 9, Appendix D)

The engine start was a significant development issue, and one we didn’t do tremendously 

well. I think Dr. Richard Gilbrech3

 mentioned that, in this era, the analytical capabilities and 

modeling were somewhat limited. That was probably a very positive statement, but a sensitive 

one as well. There was a very strong potential for gas generator burnout if you didn’t tiptoe 

through the start transient very carefully. There were several components that contributed 

heavily to that sensitivity. The first was the axial fuel turbopump. Fuel turbopumps of that 

configuration have a tendency to stall if there’s too much downstream resistance. That required 

a great deal of thrust chamber assembly thermal conditioning on the ground beforehand, and 

a variation in fuel lead time, depending on how long it had been since it had been conditioned 

on the ground, in order to make sure that it was not too cold or too warm. 

A complex two-position main oxidizer valve was utilized. It featured a pneumatically actuated 

butterfly valve. To operate, the valve had to be moved from the closed position to a fourteendegree position and the flow had to start to begin the engine. It had to dwell there for about 

one-half second, then go to the full-open position required during one and one-half seconds. 

It is characteristic of butterfly valves, however, that if the flow force increased too rapidly on 

them, they would just stop moving, unless the actuator force was sufficient to move them. 

Now, if the actuator force was pneumatic, a force balance issue occurred.

2

 Italicized text represents “lessons learned” by the conference presenters.

3

 Dr. Richard Gilbrech served as director of Stennis Space Center from January 2006 through August 2007. He 

recently returned to Stennis Space Center as associate director.

The engine 

start was a 

significant 

development 

issue, and one 

we didn’t do 

tremendously 

well. 
Paul Coffman, Rocketdyne - J-2 Saturn V 2nd & 3rd Stage Engine

 Apollo Rocket Propulsion Development 31

The third issue was the turbopump start energy from the high-pressure cold hydrogen. The 

dump of that energy had to be precisely accommodated, and it had to be both the requisite 

temperature and pressure in order to get the satisfactory start energy. That process was controlled 

by the start tank vent relief valve. People have wondered why we used two valves. Well, we 

knew there was going to be a warm-up of the cold hydrogen during the boost phase, and we 

wanted to compensate for that. We had a big valve that was going to dump immense quantities 

of cold hydrogen, and we had a leak rate that had to be calibrated as well. With all of these 

issues combined, extensive thermal conditioning was needed. We were conditioning a thrust 

chamber with cold helium on the launch pad. (See Slide 10, Appendix D)

The initial problems were encountered at Santa Susana and at Marshall. The simulated altitude 

testing at the Santa Susana Vertical Test Stand 3, 3A position, was a small capsule. The engine 

was mounted horizontally, and it had a steam ejector to pull the initial altitude. It really didn’t 

do a very good job of simulating the start transient with the spit-back. When we started flying 

the S-I-Bs, which shows the wisdom of flying early if you possibly can, we indicated the problem 

was not very well understood, and certainly was not resolved. Bottom line: we started putting 

immense activity into AEDC’s J-4 cell. For those who are not familiar with it, there was a hole 

in the ground about 100 feet in diameter and 300 feet deep. We could attach a very strong 

vacuum to it. There was a huge bell jar configuration metal container, inside of which very 

easily fit an S-IV-B stage with a J-2 attached. We did a year and one-half worth of testing 

there and, concurrent with that, fruitfully advanced our analytical modeling capability, both 

at NASA and at Rocketdyne. The net result was that we were able to achieve satisfactory test 

starts on all the ground and flight applications.

The engine started relatively slowly. It could go from 10 percent to 80 percent in about one and 

one-half seconds. Although it was relatively slow, it was quite repeatable. The bandwidth was 

fairly repeatable for 152 engines, which was pretty tight, especially considering that the inlet 

conditions to the engine depended to some extent on the stage application. Again, we were 

able to replicate this and advance the cause by advancing our analytical modeling. (See Slides 

11 and 12, Appendix D)

Another interesting little problem we encountered was high transient sideloads. We realized 

that to get at a twenty-seven and one-half epsilon nozzle and still facilitate testing at sea 

level conditions, it was going to take some doing. That was a goal. We used a conventional 

nozzle, and at a 27.5:1 ratio, it was below the three-tenth level. We assumed we were going 

to have full flow in the nozzle. We wanted to accomplish that, so we developed an adverse 

pressure gradient, or APG, nozzle that would be back up over that range as a result of overexpansion in the nozzle. There was about a two- to three-second specific impulse performance 
Chapter Two

32 Remembering the Giants

loss attributed to that. We proved the concept with both analysis and model hot fire testing. 

We had tenth-scale, solid-wall nozzles of several configurations that were evaluated before the 

APG configuration was chosen. However, once we started testing, we noticed some fairly high 

sideloads during start transients. Similar sideloads could occur also when thrust was reduced 

by changing the propellant utilization to run at the 4:5 mixture ratio and dropping the thrust 

level about 20 percent. The thrust chamber damage that we incurred was rather dramatic. The 

attach points for the actuators on the J-2 were at the upper end of the engine, and the clevis 

pins started elongating the holes. That may have been due to the fact that we were putting 

more than 100,000 pounds of pressure through them, and they weren’t designed for that. (See 

Slide 13, Appendix D)

The second thing was spotted by an associate of mine in combustion devices at the time. He 

stopped to tie his shoelaces after he looked at the thrust chamber to make sure the injector was 

okay. When he happened to glance back over his shoulder, he noticed the exit of the thrust 

chamber had a unique configuration. It was no longer straight; it was concave, and he said to 

himself – being a MIT grad – “There’s something wrong there.” There’s just no discounting a 

good education. The net result was a panic because the conventional diffuser to attach to the 

chamber was going to be approximately twenty feet long, and that wasn’t going to fit real well 

into the test stands of the day. The original configuration had a very clever turnout at the tail 

end of the tubes. We began with the subscale model that had been used to define the contour 

in the beginning, and we started testing, first with wood, then eventually with metal. We 

finally settled on something that was about a quarter of an inch long on the subscale model 

– six inches long on the actual full-scale nozzle. That eventually evolved into a water-cooled 

version. We reinforced the hat-bands by cutting the top out of them, and welding in tubing. 

We also attached clevis pins to the area where the fuel injection and hot gas manifolds, which 

allowed us to tie into a sideload attachment mechanism, or SLAM, system. (See Slide 14, 

Appendix D)

There was a strange device hanging on the backend of the chamber. We put it on to allow us 

to get through start transient. We “belt-and–suspendered” it for both single and multi-engine 

ground tests, which solved the problem with no further difficulty. We needed to have the 

capability for this to release during main stage, so we could demonstrate gimbal. If we didn’t 

have to demonstrate gimbal, we would usually just leave it attached at all times. As an aside, the 

initial version of this cheap metal monstrosity was welded to a combustion devices chamber on 

a test stand in the area, a pressure-fed test stand. The first test had been quite satisfactory. We 

planned on coming back the next day and testing it more. However, when we got back, there 

was nothing but a jagged edge. The engine systems people had been over and cut it off and it now 

was welded on the engine next door.
Paul Coffman, Rocketdyne - J-2 Saturn V 2nd & 3rd Stage Engine

 Apollo Rocket Propulsion Development 33

Now, the final major issue of the J-2 came when the Apollo 502 vehicle was launched in 1968. 

It was the second launch of the Saturn V vehicle. It was unmanned and was planned to be 

the last unmanned launch. One of the five J-2 engines cut off prematurely, and the S-IV-B 

engine operated apparently satisfactorily on the first burn, then failed to restart. We really 

went after that one. Flight data indicated there was probably a similar failure mode that caused 

both malfunctions. Hard as it was to believe, there were certain similarities in the engine area 

temperatures and in the engine operating characteristics that led us to believe that it was very 

likely the case. What we wanted to do was verify the anomalies with ground tests, define 

the failure mechanism analytically at the same time, verify it, and then please, please, please 

incorporate corrective action, as soon as possible. (See Slides 15 and 16, Appendix D)

A good example of this was the S-IV-B stage timeline. The engine experienced five engine tests 

and two stage acceptance tests prior to flight, and it had been pretty nominal. We saw engine 

chilling beginning about sixty-five seconds into flight. Then, there was a heating that began 

about forty seconds after that, and we got into chilling again during the restart. We saw very 

small performance shifts of about 4 psi and 12 psi. We also saw a little bit of change in the 

required thrust vector control. There was a slight shift in actuators that indicated something 

changed in the nozzle. Everything else was dead nominal. As a matter of fact, we noticed that 

we turned loose the start tank discharge valve and waited a number of seconds before there was 

a second burn cutoff initiative because it didn’t go.

We started looking at the augmented spark igniter. The fuel line to the augmented spark 

igniter had a single-ply bellows, and we said, “Feed until that bellows fails.” The idea was if 

that bellows failed, it would account for burnout of the augmented spark igniter, and that 

would certainly mirror, or have the possibility of mirroring, the chilling and heating that we 

observed. We set up an engine at Santa Susana to investigate that. We planned a sequence of 

events where we would have sixty-five seconds of normal operation, then we would incur a 

small augmented spark igniter fuel leak. We would increase the leak subsequently to complete 

the failure and allow backflow of the augmented spark igniter combustion products for the 

final thirty seconds of the operation. I had never been so happy to see copper in a flame of 

a rocket engine in my whole life. It went right down the line. There was substantial erosion 

of the device. The cavity burned out. The performance loss was similar and so was the thrust 

vector change. We figured we had it nailed. Within a day at Marshall Space Flight Center, the 

S-2 engine that had been set up and tested showed the same characteristics, so we figured we 

had done a good job in duplicating the failure on the ground. The analysis did indicate that 

the single-ply bellows could fail and that probably liquid air, since there was hydrogen passing 

through it on the ground, could damp the oscillations. That hypothesis was tested on similar 

One of the five J-2 

engines cut off 

prematurely, and 

the S-IV-B engine 

operated apparently 

satisfactorily on the 

first burn, then failed 

to restart. We really 

went after that one.
Chapter Two

34 Remembering the Giants

bellows both with inert fluids and liquid hydrogen, and showed the failure. Some people have 

asked why the bellows were there in the first place. Well, our analysis initially indicated the 

bellows was a wise idea because of thermal changes that were going to occur during engine 

operation. So, the bellows were put in to compensate for that. In hindsight, a good idea turned 

out to be a bad idea. We got rid of the design modification. Solid-line Class 1 welds, and the 

revised configuration was qualified and incorporated on all the engines in the field. Eight 

months later, the first manned Apollo 503 flew – a pretty good response to the issue. (See Slides 

17 and 18, Appendix D)

We began this project in 1960, and it was schedule-driven. It was what we would characterize 

today as a low technology readiness level. The thermodynamic characteristics of para-hydrogen 

were under investigation by the National Bureau of Standards4

. Given the criteria, the 

development program was quite satisfactory: two years to the first 250-second test, formal 

development completed in six years. The J-2S and linear spawned off from that, to investigate 

and keep people sharp. There were over eighty modes of unplanned cutoffs. The engine and 

the analytical tools utilized to develop it enabled us to be satisfactory on the space shuttle main 

engine, the RS-68 engine, the X-33 engine, and what is now being characterized as the J-2X 

engine. (See Slide 19, Appendix D)

4

 Currently known as National Institute of Standards and Technology.

The engine and the 

analytical tools 

utilized to develop 

it enabled us to 

be satisfactory on 

the space shuttle 

main engine, the 

RS-68 engine, the 

X-33 engine, and 

what is now being 

characterized as the 

J-2X engine.
Paul Coffman, Rocketdyne - J-2 Saturn V 2nd & 3rd Stage Engine

 Apollo Rocket Propulsion Development 35

---------------------------------------

Editor’s Note: The following information reflects a question-and-answer session held after Coffman’s 

presentation. 

STEVE FISHER: I’d like to introduce Mr. Manfred “Fred” Peinemann, a guest from The 

Aerospace Corp. Please stand up, Fred, so we can pinpoint you while you ask your question.

QUESTION: Did you encounter interesting materials compatibility problems during the 

development with the J-2? Did you encounter hydrogen embrittlement?

COFFMAN: We didn’t encounter hydrogen embrittlement. The first problem we encountered 

was the first summer we started trying to do engine testing. We found that we couldn’t get 

liquid hydrogen through the engine because the facility dump system and the engine bleed 

valves weren’t big enough to allow it in a Santa Susana summer5

 to achieve liquid at the inlet 

to the engine.

QUESTION: This goes basically to both F-1 and J-2. Can you comment on the philosophy 

when you had failures and anomalies? How fast you turned those around? From what we 

can tell, it looks as though you were continually testing. You didn’t stop and have failure 

investigation teams, and factor all that back in. How did your process for getting that corrective 

action work into the design occur?

COFFMAN: It is true that we were constantly testing, but there was a formal Unexplained 

Condition Report (UCR) system. When a component failed in an engine test, or in a 

component test, the UCR was generated. The first thing we had to determine was whether the 

latest configuration had failed, or whether it was something else that had already been replaced 

on the engine and the failure wasn’t useful. There was enough hardware richness that we had 

five engine test stands, running two shifts. I remember one of the technicians looked up and 

said, “Wow, Friday. Only two more work days until Monday,” which characterized the era 

pretty well. The attempt was to incorporate any corrective actions immediately, and put out 

kits for retrofit of everything in the field. The logbooks of which kits applied to which engines 

were pretty interesting. But, certainly, we had a very strong cadre of developmental personnel, 

at both the component and engine system levels, and no problem was ignored.

5

 Meaning the hot climate in north Los Angeles desert.
Chapter Two

36 Remembering the Giants

BOB BIGGS: I think for major problems we have always - we did in the F-1 program, we do 

in space shuttle main engine program - established an investigation team with an autonomous 

team leader to run the investigation. We assigned on the order of twenty to thirty people to 

work the problem, and it is generally totally resolved within thirty days. Today, we would not 

test for a period of time. I think on F-1 that was rare. We would continue testing while the 

investigation was going on.

QUESTION: Paul, from your talk, it appeared that you did a lot of testing over at Santa Susana 

and, obviously, we all know that the facility is not in great shape right now. You mentioned you 

had as many as five test stands going at the same time. Do you have any recommendations for 

as to how to get prepared for those tests?

COFFMAN: I think you are starting from a much more advanced position with respect to 

looking at developing future engine and stages. You also are not pushing the envelope nearly as 

much. So, I think that from what I have seen of the test planning for the J-2X, for example, I 

think that the group is well-connected between Marshall Space Flight Center and Rocketdyne 

in some of the net meetings where I have been an observer to make sure they do a thorough 

job of testing. Now, the question of “thorough” will always have different meanings to different 

people. I would encourage, until we are very satisfied with the start characteristics of that engine at 

altitude, that we keep open the vacuum capabilities. I think the test planning is headed in that 

direction.

FISHER: This is my opportunity as a moderator to throw in my two cents. I think, as Paul 

mentioned, we have come a long way. Our analytical tools, especially our transient modeling 

and understanding of the processes, are much better than they were. However, I think that 

with some of our recent successes in engine programs, we have done very well. We tend to get 

a little bit optimistic on a no-failure plan. I think one of our weaknesses over the years is we 

haven’t planned for failure. If you are in a program, and you have a lot of extra hardware, and 

you have money to resolve it like we did back in those days, that is one thing. In today’s world, 

where we plan for every engine test being a success, that may rise up to bite us. So, for my two 

cents, a little bit of reserve is required for those failures that are unknown. It would be nice if we 

could carry that. I don’t know how you go get that. 

QUESTION: I’m Thomas Carroll. I work systems for J-2X testing. Were there any compromises 

that were made for any of the previous J-2 programs, other than the planning for failures? Are 

there any compromises that were made that we are looking at fixing for J-2X that we want to 

fix, that we want to change, and want to make better?
Paul Coffman, Rocketdyne - J-2 Saturn V 2nd & 3rd Stage Engine

 Apollo Rocket Propulsion Development 37

COFFMAN: I don’t think so, at least as the engine is evolving now. There seems to be a rather 

thorough study of the “S” (i.e. J-2S) and the basic J-2. You must recognize that the “S” was 

commissioned to be physically and functionally interchangeable, so it had to meet the 80-inch 

diameter, and the same length, and the same attach points, and it had to fit the same buildup 

characteristics, and so forth. There may be some real possibilities for engine definition for the 

“X” (i.e. J-2X) in coordination with the stage to be mutually advantageous, but I don’t have 

any concrete examples of compromises that would cause a problem going forward. 
38 Remembering the Giants

Apollo Reaction Control System Engines (Source: Pfeifer presentation charts, Appendix E)

Apollo Spacecraft/LM Adapter (Source: Apollo Training Manual “Apollo Spacecraft & 

Systems Familiarization” – March 13, 1968)
Apollo Rocket Propulsion Development 39

Gerald R. “Jerry” Pfeifer joined the Marquardt Company1 in 

1961, performing early work to develop the parametric rules 

used for Marquardt rocket design and teaming with a small group 

to build the company’s first hypergolic rocket test facilities and 

rockets. When the Apollo reaction control system project came 

to Marquardt, Pfeifer helped shepherd it through its qualification 

phase. The work led to involvement with many other successful 

space propulsion programs, including work for NASA’s space shuttle. Pfeifer eventually went to 

work at Rocketdyne to help develop its Terminal High Altitude Area Defense (THAAD) attitude 

control engines. He now is working on the Airborne Laser Program at The Boeing Company. 

In late 1960, I went to work for Marquardt, which was an interesting company. It was a big, 

little, big, little, big, little, gone company2

. When I went to work there in 1959, they had 5,000 

people. I went to work on a program called Bomark, which was an interceptor missile. That 

program was downsized, and we had a chief engineer named Warren Boardman, who had this 

great idea after Sputnik that there had to be a future in space somewhere, and since we were 

a propulsion company, we would probably be good at something like that. We went off to 

figure out what it might be. He convinced the company to give us some internal research and 

development, and they turned five of us loose to figure out what propellants were and to try 

and figure out what the parameters were that could be used to make one of these silly things. 

(See Slide 2, Appendix E) 

Boardman had a really novel idea that instead of what the industry was headed for at the time – 

1

 Marquardt Company of Van Nuys, Calif., is no longer in existence. Carl Stechman is also mentioned in this article 

as a collaborator of the speaker at Marquardt.

2

 Italicized text represents certain “lessons learned” emphasized by the conference presenters.

Chapter Three

Gerald R. Pfeifer

Aerojet - Attitude Control Engines
Chapter Three

40 Remembering the Giants

many different sizes of engines where you would run long durations to control attitude in space 

– if you pulse-modulated the things, you could do a whole job with thrust levels from five to 

about 100 pounds. We developed enough credibility in that early time that we managed to win 

a little program called SINCOM. In that program, we had two thrusters, with a single unlike 

doublet injector bolted assembly on both of them. Both the twenty-five-pound and five-pound 

single unlike impinging injectors3

 were what we had been working with in the backyard. The 

backyard was really what it was. This little group that we put together built our own test 

facility, and we made one mistake. Basically, the facility consisted of a six-foot high concrete 

wall with propellant tanks on one side, and two guys on the other side looking through the 

window to see what happened. The mistake was we didn’t remember that the wind blew from 

the west, and we were on the east side of the test stand. The engines weren’t very efficient, so, 

frequently when we evacuated, we went through this red cloud. 

However, that program gave us enough credibility that we were able to convince North 

American Aviation Inc. and NASA to give us the contract to build what they called a common 

engine. It had three attitude control applications: the Apollo service module, lunar module, 

and the Apollo command module, which was to be a varied engine. We got about a year into 

the program, almost a year, before everyone came to the conclusion that this little company, 

which had gone from 5,000 to 500 employees really couldn’t handle a large program of that 

magnitude; so, the command module part of that went to the Rocketdyne Company. (See 

Slide 3, Appendix E)

My colleague Carl Stechan started at Rocketdyne and came to Marquardt. I started at Marquardt 

and ended up at Rocketdyne. We did all of our development work with single doublets,4

 and 

one of us was convinced that a 100-pound single doublet could work at the 100-pound level 

very satisfactorily. It really wasn’t too bad an idea, except the chief engineer said we had to 

have a specific impulse of 300; 270 didn’t cut the mustard. The thing evolved then into what 

it is now and started out from the original configuration, the R-4A. The R-4A was an eightdoublet with eight film cooling holes and two valves. The fuel valve canted off to the side to a 

circumferential manifold, and the oxidizer valve fed right through the center of the injector to 

eight outward pointed doublets. It featured a molybdenum nozzle, coated with molybdenum 

disilicide or Durac-B. (See Slide 4, Appendix E)

Basically, we had three groups working. We had a guy doing the injector and combustor design. 

3

 Injector with a single element (i.e. unit) consisting of one fuel orifice and one oxidizer orifice that impinge the 

resulting fuel and oxidizer streams together just downstream of the injector face.

4

 A type of injector where two fluid jets impinge to create a spray.
Gerald R. Pfeifer, Aerojet - Attitude Control Engines

Apollo Rocket Propulsion Development 41

Carl Stechman was doing heat transfer work. We had another group doing materials studies. In 

actuality, the molybdenum system they came up with wasn’t too bad except for two features. 

One feature was the material, which, if not coated, vaporized when it got hot. It was very, very 

sensitive to chipping. The way we tested the nozzles to find out if they were okay was to put 

them in the oven at 2,500 degrees; if they smoked, they were gone because you could see them 

while they smoked. They were also very brittle. It’s a very brittle material. The ductile brittle 

temperature was about 60 degrees Fahrenheit, so once they got cold, they were like just tapping 

a piece of glass and, with our shocks, that was not a good thing. In order to keep an engine in 

one piece – because the nozzles kept splitting – we got the R-4B configuration in which we put 

on a ribbed L-605 nozzle attached to the molybdenum combustion chamber. That continued 

to grow, and they all looked the same. From the outside, they were all pretty much the same. 

(See Slide 5, Appendix E)

The original eight-on-eight injector basically stayed the same from the first injector we put 

together in 1962. The valves improved. Scads of things changed. However, the Preliminary 

Flight Rating Test (PFRT) injector differed only from the qualification injector in flight units 

in the valve seat materials. We were having difficulty with the high temperature soak back after 

long runs. We decided it would be a good thing to strengthen the Teflon valve seat. In order to 

do that, we put fiberglass into the Teflon valve seat, and it really stayed together, except when 

the engine heat soaked back; the Teflon grew and grew and grew, and the oxidizer didn’t get 

through the oxidizer side, if it got too hot. We flew two successful missions around the moon 

with that PFRT engine. Boeing made that particular device. We held our breath all the way 

because it was used as the brake engine and an apogee engine both; we were really worried 

about the valve seats getting too big. In the middle of the chamber of the basic R-4D engine 

as it evolved into the engine that flew on all the missions, there was a little tube sticking down, 

and that was a magic thing called the pre-igniter. It was both part of our problem, and a savior 

of part of our problems. The thing was a cooled molybdenum chamber with an L-605 nozzle. 

The reason we stuck the L-605 on was that it was a high temperature material. The ribs had 

to be on there to keep it round. It also had one other nice feature on the service module, they 

stuck out in all directions. The flight engines all ended up with a no-step design on the nozzles 

because they turned out to be convenient. (See Slides 3-5, Appendix E)

On the lunar orbiter, the 100-pound engines faced down. They used cold gas for attitude 

control. That was a pre-qualification test engine. We did have some technical difficulties. The 

first one was high heat transfer burning. Today, I can confess it was first-tangential combustion 

instability mode, but at the time this happened, it was really an interesting thing because in 

. . . a magic 

thing called 

the pre-igniter.

It was both 

part of our 

problem, and 

a savior of 

part of our 

problems.
Chapter Three

42 Remembering the Giants

the Apollo mission duty cycle,5 it had to perform a ten-second run followed by a very specific 

off time, then a three-second run followed by another thirty seconds, and then ten seconds of 

continuous burn. The characteristic was they would go through the first two very nicely, and 

on the third burn, in about one and one-quarter seconds, the temperature on the chamber wall 

would pass 3,100 degrees Fahrenheit. Shortly before that happened, the fire would come in a 

thrust vector mode, which wasn’t planned. Actually, we compromised with NASA on that one. 

We just decided that was probably not a good duty cycle because nobody quite knew what it 

was or how to cure it. We compromised and went through a different duty cycle to get around 

that particular problem, which was the only place that it really occurred on that engine. (See 

Slide 6, Appendix E)

We had a series of difficulties with spikes. That engine had some neat things to get through. It 

was generally operated in ten milliseconds on-time duty cycle. Thermal control was something 

we thought about, but the combustor in that mission had to sit there and, characteristically, it 

would go down to minus-100 degrees Fahrenheit. When the engine shut off in short pulse duty 

cycle in that kind of a radiative environment, the oxidizer entered the manifold first because it 

had gotten a very high vapor pressure relative to the fuel, and it went to the closest cold spot, 

which was at the chamber wall, and the fuel would come dribbling out. It also looked for a 

place to stop. It would freeze right on top of the oxidizer. We would build multiple layers of 

fuel oxidizer, and then, we would get a run long enough up to really do something. We’d get a 

really exciting event. I deemed it “pre-unplanned disassembly” at the time. That problem really 

went away when we learned a little bit about fuel control and kept injectors and propellant 

temperatures a little warmer. Aerozine-50 (A-50) propellant was the original propellant we 

selected because the big engines that had to be used to get us to the moon all used A-50, so that 

had to be the propellant of choice for us. (See Slides 7 and 8, Appendix E)

Carl Stechman didn’t think the pre-igniter did something, but someone else did, and the 

NASA folks kind of agreed. The pre-igniter really was a good thing, but the serendipitous thing 

it did was the hot phase or first tangential instability went away because we had changed the 

interior configuration of the chamber a little bit. It took years for us to figure out what a good 

thing we had done without planning it.

The other major technical issue was the inner manifold explosions resulting from monomethyl 

hydrazine (MMH) evaporating and depositing like dew in the cool spot in the engine. That 

5

 A parameter referring to time-varying, on-off cycling of any rocket thruster or engine (also to intermediate states 

of thrust, if needed).

It took years 

for us to figure 

out what a good 

thing we had 

done without 

planning it.
Gerald R. Pfeifer, Aerojet - Attitude Control Engines

Apollo Rocket Propulsion Development 43

was a recurrent problem, and it was the reason the space shuttle has a limit of 70,000 feet 

operation today. In space, it can’t happen because there is no way to get the phenomena to 

occur because the pressure is inside the chamber and injector after engine firing, and it is less 

than the thermodynamic triple point of the fuel. We did a lot of premature disassembly (i.e. 

failures). The chamber would disintegrate during pulse mode firings in ten-millisecond pulses. 

The thing that really convinced us to drive to the pre-igniter was the program manager’s data 

plot. After a night of testing, he had drawn a plot that showed chamber wall thickness across 

the bottom. The ordinate was the peak value of the measured pressure. He had infinity drawn 

both directions. It seemed like that might be a wise thing to change. Later, the pre-igniter 

really did alleviate the spike problem to a high degree. When we changed to warmer operating 

conditions, like a 70-degree Fahrenheit propellant, the problem went away and never came 

back.

In later developments, we put in a columbium chamber, which was a lot more tolerant. A-50 

going away was also a good deal because it was more sensitive to that chamber. There was a lot 

of resistance going to the niobium chamber (niobium and columbium are the same material) 

because we had a molybdenum chamber that worked. If it was good enough and if it ain’t 

broke, don’t fix it. Stechman wrote an article about this feature in the Journal of Spacecraft and 

Rockets. It was not a real detailed discussion, but there is some discussion there. (See Slide 8, 

Appendix E)

Another problem we ran into was a thing called a ZOT. There used to be cartoon called BC.

A couple of the key characters were an anteater, who was always getting after ants in the ant 

hill, and a snake, who always got struck by lightning when he would do something inane. The 

ZOT was something like that snake. We didn’t really understand it too well, but it sure was 

something. It wasn’t called the ZOT originally. It was called, “What the hell caused that?” (See 

Slides 9 and 10, Appendix E)

If you could picture the oxidizer valve off in one of the early Apollo engines after it disassembled 

itself from the engine, there was a little squirrel thing running around in there; that was the 

seal between the injector and the valve. We thought something happened; after it happened 

a couple of times, we began calling it a ZOT, because, like that cartoon snake’s observation, 

something was really interesting. Its characteristics were pretty definable. Ordinarily, when you 

opened a valve, you got a smooth pressure decay. Instead, what would happen was the pressure 

at the instant of the valve opening would disappear off the chart. We finally associated that 

with high pressure. That was the clue that we were looking for, and we finally figured it out 

after somebody asked us to run some vertical, facing-up engine tests. We discovered that it 

really was kind of an interesting phenomena confined to a relatively small operation area. We 
Chapter Three

44 Remembering the Giants

made a ZOT plot written by Stanley Mitten, who was one of the better engineers that I dealt 

with in the science part of the world over the years. Basically, it said that at very low pressure, 

there were no ZOTs because the fuel emptied the manifold and evaporated really quickly and 

was gone from the chamber almost instantly. (See Slide 10, Appendix E)

At very low pressures, it wasn’t such a good deal – particularly in the pulse duty cycle, which 

was pretty hardware dependent. If there was a place in the engine between the valve seat and 

the injector face, that could be cold. Guess how it got cold. When the oxidizer evacuated the 

injector, all these small engines had the valves separated from the injector face by a thermal 

standoff. If the valves seats could be kept cool, while the injector face was running very hot, 

the little thin standoffs in between the valve seat and the injector could get very cold when 

the oxidizer was evaporating and emptying the manifold. The fuel came out after, and if the 

ambient pressure was just right and the ambient pressure was above the triple point pressure, 

dew would develop inside the oxidizer manifold. If enough dew collected, on the next pulse, 

liquid fuel would develop inside the oxidizer manifold, and it would generate some horrendous 

pressures. We’d get premature disassemblies6

 again. It wasn’t a problem for Apollo because 

that particular engine never ran in real life in its missions until it was well out in space. This 

wasn’t a space problem; it was purely a ground problem. It was a problem for the space shuttle 

because the shuttle engines operated for attitude control during the early phases of reentry. 

There was a big concern because through stage down to 70,000 feet, there is the risk of having 

a ZOT occur and a valve leak or too much condensed fuel in the oxidizer manifold harming 

the hardware pretty severely. The ZOT story, today, is probably still a problem that people 

need to think about for the upcoming vehicles, because the shuttle engine is going to be, with 

any probability, one of the attitude control engines for the new large vehicles. The 870-pound 

thruster is the ideal size for intermediate control. This is one of those problems now that we 

recognize, we know what it is, and we know how to deal with it. In the current shuttle engine, 

that is an easy thing to fix. (See Slides 11 and 12, Appendix E)

People often wonder about schedules. Marquardt had no schedule problems at all. Our task 

was really pretty easy. We just had to do a little development, which happened in the 1962-

1963 time frame, and then, we did a little more development. Then, we got the R-4D, and we 

did some more development. By the time we got out to the 1965 time frame – and that is a 

September 1965 schedule as a matter of fact – we had managed to make a little bit of hardware 

and break quite a bit of it. We were just getting into the qualification program, and we had this 

small task to make 700 engines. Interestingly enough, once we decided we had the problems 

6

 Meaning that the thruster exploded.

We were just 

getting into the 

qualification 

program, and we 

had this small 

task to make 

700 engines.
Gerald R. Pfeifer, Aerojet - Attitude Control Engines

Apollo Rocket Propulsion Development 45

behind us and were pretty well along the path of making hardware that we believed was okay 

to support manned flight, we geared up and produced about ten engines a week out of that 

little factory, which had grown from 500 to about 800 people during the Apollo Program. 

There wasn’t a straight-forward acceptance test that consisted of four pulse-firing test series. 

There were two ten-second runs to get good performance data. Then, we had to run sixtysecond durations, so we could demonstrate engine life margin. On average, we ran two of those 

engines a day for two years to get all of that hardware out of there. It was an interesting time. 

(See Slide 13, Appendix E)

The cost was always an issue. The hardware evolved up to be an Aerojet product today. That same 

R-4D heritage resides up at Aerojet Redmond now. Carl Stechman is their corporate knowledge 

of what the history of that engine is and really knowledgeable about what makes those engines 

work. In 1960, the price of an engine was about $30,000 each. If you take the Consumer Price 

Index and the ratio of that – just taking cost of living adjustments – it would be up to about 

a factor of six. I’m absolutely sure they would have liked to produce 2,000 more. I’m pretty 

well convinced they would have been just as happy as clams to make another couple thousand 

of these things. We made about 650 engines. That was 650 production units. We actually 

flew 469 of them during the Apollo Program, which is an astounding number of little rocket 

engines that actually fly in space all on one program. In all that time, and the millions of cycles 

that were put on during that whole program, there was not one R-4D valve or engine failure. 

We were really kind of tickled that we may have done something good. (See Slides 14 and 15, 

Appendix E)

What happened to the other 200 engines? NASA had this requirement that you had to keep 

both the hardware and the documentation around for at least ten years after the conclusion 

of the program. We had a storage spot. We tested the original lunar module qualification unit 

with the four engines, which on both the service module and the lunar module were stored 

structures. That unit sat in the back of a storage room for about forty years before the company 

downsized; they were throwing stuff away, and Carl Stechman managed to recover one. I know 

where one other one is for sure – in my office. A bunch of them were sold to Rocketdyne for 

some classified program. They sold eleven of them for another program. NASA’s Jet Propulsion 

Laboratory got a bunch of them, primarily to be taken apart and used for valve testing. There 

are some in the Smithsonian, a few in the space museum down in Alamogordo, New Mexico. 

There are some in NASA’s Kennedy Space Center museum. There are some in other museums 

somewhere in the South. There is some facility that has quite a number still sitting in the 

original containers. Last I heard, there were still a number of them in storage down at Johnson 

Space Center, waiting to find a home. (See Slide 16, Appendix E)

In all that time, 

and the millions of 

cycles that were 

put on during that 

whole program, 

there was not one 

R-4D valve 

or engine failure.
Chapter Three

46 Remembering the Giants

I was running a 200-pound thrust development program at Marquardt in about 1992. I actually 

took one of those engines off a particular module because we were having a data problem in 

our test stand. We couldn’t seem to get data to repeat, or we couldn’t get anyone to believe that 

we could repeat data anyway. We actually pulled a logbook for one of those engines out of its 

original dead storage, took one of those engines off of that module, put it into the test stand, 

and re-fired the thing. It repeated its original acceptance test within two seconds of specific 

impulse and about one-half pound of thrust. They seemed to have stood the test of time fairly 

well. There was some documentation still around up until a couple of years ago; I had a copy 

of the original PFRT report and the qualification report I wrote. Carl Stechman is a good 

source because he kept a lot of the original stuff from a historical standpoint. The thing has 

really gone a long way from where it started. It started as an injector. The basic injector was the 

R-4D that is still sold today. It was the same one that we developed just before the pre-igniter 

was incorporated because we got some thermal control. We got a little smarter. We turned 

that one with a large area nozzle made out of niobium into one with a 311-second engine. I 

left Marquardt at about that time. Then, they decided, “Gee, if a little is good, more must be 

better.” It had more nozzles and some reductions in film cooling. (See Slide 17, Appendix E)

Today, the engine marketed by Aerojet is still that basic eight-on-eight configuration with the 

changes that they have made into it. It has a current specific impulse of about 323 seconds. 

From where it started out in the good old days at 292 seconds to where it is today, it’s come a 

long way. (See Slide 18, Appendix E)
Gerald R. Pfeifer, Aerojet - Attitude Control Engines

Apollo Rocket Propulsion Development 47

--------------------------------------

Editor’s Note: The following information reflects a question-and-answer session held after Pfeifer’s 

presentation. 

QUESTION: Give us a sense, if you would, of where all these different articles were tested – 

the ones that were to fly as well as the ones that were part of the development activity. Were 

there multiple facilities up and down the West Coast or were they all contractor? Were they 

altitude as well as ambient?

PFEIFER: All the engines were both qualification and acceptance tested at Marquardt’s 

facilities. After we won the Apollo Program contract, we went off and built two vacuum test 

facilities, which simulated altitude continuous firing for as long as we wanted to run an engine. 

They would run days and days with the same capability we had on steam ejection. We did all 

of the testing in both for the qualification and the acceptance test. One of them was a large 

ball, which was an eighteen-foot diameter sphere, evacuated again with a big steam ejector 

system that could be used for system testing; that’s where we did the Lunar Excursion Module 

testing. We put the whole cluster in there and tested the entire cluster at the simulated altitude 

conditions. The lowest altitude we tested at – typically an acceptance test – was 105,000 feet 

simulated altitude. The big ball – because people were interested in what they called goop 

formation, which is an unburned hydrazine product migrating to cold surfaces on different 

parts of spacecraft – was built to address those kinds of issues. We ran long-life tests in a 

simulated space environment with the entire inside of the test cell around the test article, 

liquid nitrogen cooled, so it could act as getter for any of the exhaust products. That particular 

facility could pull down to about 350,000 feet (atmosphere) equivalent altitude, which was 

pushing pretty close to the thermodynamic triple point of the MMH. It was a good test 

facility. Those facilities are no longer there. When the guys at Marquardt sold the company to 

what eventually became part of Aerojet, all those test facilities were cut off at the roots. I think 

they have a movie studio there at this point. That part of it is truly not recoverable, but it did 

some excellent high-altitude, space-equivalent testing at the time.

STEVE FISHER: Regarding your facilities, and probably more so towards the end of 

Marquardt, how did you guys test in the San Fernando Valley with all the leaks and the stuff 

you had? How did you guys manage the Environmental Protection Agency (EPA)?
Chapter Three

48 Remembering the Giants

PFEIFER: Surprisingly, we had very few problems while testing in the San Fernando Valley. 

In the early 1960s, nobody had ever seen dinitrogen tetroxide (N2

O4), so that wasn’t too big 

a deal. We really did only make small, red clouds. In all the hundreds of thousands of tests 

and probably well over one million firings that I was around that place for, in all that thirtysomething years, we had a total of one serious injury associated with rocket engine testing and 

propellants. Because we were trying to figure out what propellants would really be good, we 

tried all of the fun stuff like the carbon tetrafluoride, chlorine pentafluoride, and pure fluorine. 

The materials knowledge wasn’t all that great at the time. On one test, the fluorine we had 

didn’t react well with the copper they were using for tubing, and it managed to cause another 

unscheduled disassembly of the facility. It was very serious. It’s like one of those Korean War 

stories. The technician happened to be walking past the test facility when it decided to blow 

itself up. A piece of copper tubing pierced one cheek and came out the other. That was the only 

serious accident in all of the engines handled in all those years.

Now, we did have a problem with the EPA later because they figured out what the brown 

clouds were about. We built a whole bunch of exhaust mitigation scrubbers to take care of 

engine testing in the daytime. In general, we operated the big shuttle (RCS) engine, the 870-

pounder at nominal conditions; they scrubbed the effluents pretty well. If you operated that 

same 870-pound force engine at a level where you get a lot of excess oxidizer, yeah, there’s a 

brown cloud. But, you know, it doesn’t show up well in the dark. They did do some of that. 

But, that’s gone; it was addressed one way or another.
Gerald R. Pfeifer, Aerojet - Attitude Control Engines

Apollo Rocket Propulsion Development 49
50 Remembering the Giants

Saturn S IV B Stage Ullage & Apollo Command 

Module Reaction Control Engines (Source: Tim 

Harmon’s presentation charts, Appendix F)
 Apollo Rocket Propulsion Development 51

Tim Harmon has more than forty-one years of experience 

with The Boeing Company and its Rocketdyne Division. He 

retired as the chief systems engineer on the MB-XX Program, 

a cryogenic upper stage propulsion system. Harmon also 

was involved in ten Rocketdyne engine development 

campaigns, ranging from small attitude control engines to 

a large Aerospike engine. In the Apollo Program, he worked 

on development of the lunar module ascent and command module attitude control engines. 

Harmon earned a bachelor of science degree in aeronautical engineering from Purdue 

University, a master of science in mechanical engineering from the University of Southern 

California, and a master of business administration from Pepperdine University.

I hired in at The Boeing Company’s Rocketdyne Division in 1963. There are two themes I 

would like to emphasize about my experience there: the environment at Rocketdyne, and 

using lessons learned. 

It started out in 1957 with the first Sputnik launched into space. The United States’ response 

was a vehicle that went all of four feet off the ground before exploding. Then, in 1961, the 

Soviets put an astronaut in orbit. What was our response? We got Alan Shepard up in space 

for fifteen minutes. When I started work at Rocketdyne, there was a space race, and we were 

behind. We didn’t look that good. That kind of environment was intense. Rocketdyne’s response 

was to hire more staff. We had employee serial numbers in those days. I hired on in April, and 

my serial number was 318083. By the end of the summer, there were many new hires. The 

staff I talked to that hired in June, July, and August had serial numbers that were a thousand 

points higher than mine. It was a good environment from the standpoint that it was growing. 

Chapter Four

Tim Harmon

Rocketdyne - SE-7 & SE-8 Engines
Chapter Four

52 Remembering the Giants

When something is growing, there is opportunity, and it’s a fun environment. On the other 

hand, in my first six months at Rocketdyne, I had three managers because they were getting 

promoted, moved, or changed. That was a little disconcerting. You have pluses and minuses 

with a growing environment. 

Now, let’s talk about off-the-shelf technology and lessons learned with the Apollo Program. 

There were a dozen SE-8 engines in the Apollo capsule; in the S-IV-B Stage, there were two 

SE-7 engines. These engines were similar in design and function. They were based on the 

Gemini Program vehicle attitude control engines. The Gemini Program used small engines that 

Rocketdyne built. These were buried engines. One of the technical issues with buried engines 

is keeping engine heat from soaking into the Gemini capsule. The solution was ablative-cooled 

engines that insulate the vehicle from engine-firing heat. The Gemini Program success was the 

basis that Rocketdyne used to win the contract to develop the engines for the Apollo Program’s 

Saturn vehicle. (See Slide 2, Appendix F)

On the Gemini vehicle, we had orbital maneuvering engines (OMEs), and we had reaction 

control system (RCS) engines. The OMEs were 100-pounders (generating 100 pounds-force of 

thrust). The RCS engines were twenty-five-pound-force (lbf), buried systems, and depending 

on the angle, the nozzles were scarfed (truncated at an angle), so they wouldn’t impinge on the 

vehicle. Rocketdyne said, “Well, let’s go with an off-the-shelf design where we and you (NASA) 

can save money.” It was a good idea, and I think it proved itself. (See Slide 3, Appendix F)

Gemini’s 100-pound force Orbital Attitude Maneuvering System (OAMS) illustrates some of 

the problems we had. We had two propellant valves. We had an ablative thrust chamber with 

different wraps of ablative material. There were a silicon carbide liner and a silicon carbide 

throat. One of the problems with silicon carbide material, while very sturdy, was that it would 

crack due to thermal shock. The thermal shock issue was never a problem on the Gemini 

vehicles. The throats would crack and the sleeves would crack, but they would all stay in 

place. That was the design, and NASA went with it. It was man-rated (deemed safe to carry 

humans) and successfully flew. The Gemini thruster performance was adequate. As a pressurefed system, it was fairly simple. The S-IV-B engine has the same look as the Gemini engine. 

They did de-rate the thrust level. NASA didn’t feel they needed the thrust. It also was pressure 

fed (no propellant pumps). It had a 274-second specific impulse using nitrogen tetroxide and 

hydrazine propellants. The design year was 1965. This engine application, for the S-IV-B 

Stage, was to settle propellants into the bottom of the stage, so they could start and restart the 

J-2 engine. It was a continuous burn, with no engine pulsing. This fairly simple engine was a 

“lessons learned” direct application, and there were no development problems. (See Slides 4 

and 5, Appendix F)
Tim Harmon, Rocketdyne - SE-7 & SE-8 Engines

 Apollo Rocket Propulsion Development 53

The 70-pound SE-7 engine is very similar with its two valves, ablative material, a silicon carbide 

liner, a silicon carbide throat, and overall configuration. There were different wraps. One had 

a ninety-degree ablative material orientation. That is important because it caused problems 

with the SE-8, but not for this application. It was not overly stressed. It was a validation of 

the off-the-shelf application approach. There were two SE-7 engines located on the stage near 

the bottom. They had their own propellant tanks. That was the application. All it did was give 

a little bit of gravity by firing to push the propellants to the bottom of the tanks for start or 

restart. It was not a particularly complicated setup. (See Slides 6 and 7, Appendix F)

What had we learned? This was a proven engine in a space environment. There weren’t any 

development issues. Off-the-shelf seemed to work. There were no operational issues, which 

made the SE-7 very cost-effective. Besides NASA, the customer for this application was the 

Douglas Aircraft Company. Douglas decided the off-the-shelf idea was cost-effective. With the 

Gemini Program, the company was McDonnell Aircraft Corporation, which was part of the 

reason the off-the-shelf idea was applied to the Apollo. (See Slide 8, Appendix F)

However, here are some differences between Apollo and Gemini vehicles. For one thing, the 

Apollo vehicle was really moving at high speed when it re-entered the atmosphere. Instead of 

a mere 17,000 miles per hour, it was going 24,000 miles per hour. That meant the heat load 

was four times as high on the Apollo vehicle as on the Gemini craft. Things were vibrating a 

little more. We had two redundant systems. Apollo was redundant where it could be as much 

as possible. That was really a keystone or maybe an anchor point for Apollo. We decided to 

pursue the off-the-shelf approach. However, the prime contractor was a different entity – the 

North American Space Division. They thought they ought to tune up this off-the-shelf setup. 

It was a similar off-the-shelf application, but at a higher speed. They wanted to improve it. 

What they wanted to improve was the material performance of silicon carbide. They were 

uncomfortable with the cracks they were seeing. They were uncomfortable with the cracks 

in the throat, and feeling that the environment was a little tougher, that maybe it was going 

to rattle, perhaps something would fall out, and they would have a problem. They wanted to 

eliminate the ceramic liner, and they wanted a different throat material. (See Slides 9 and 10, 

Appendix F) 

The Rocketdyne solutions were to replace silicon carbide material with a more forgiving ceramic 

material. Also, due to the multiple locations within the vehicle, the shape of the nozzles varied. 

Some nozzles were long, and some nozzles were short. We came up with a single engine design 

with variable nozzle extensions and configurations to fit particular vehicle locations. (See Slides 

10 and 11, Appendix F)

Instead of a mere 

17,000 miles per 

hour, it was going 

24,000 miles per 

hour. That meant 

the heat load 

was four times as 

high on the Apollo 

vehicle as on the 

Gemini craft.
Chapter Four

54 Remembering the Giants

The engines cost about $35,000 in 1963, which was how much you paid for a Rolls-Royce 

automobile. So, there were twelve Rolls-Royces on every Apollo vehicle. That’s important because 

it meant we really had a fair amount of investment by the time we got to the end product, and 

we didn’t want some problem to show up that would cause us to discard the engine.

Another kind of issue was the work environment. There was a lot of competition at Rocketdyne 

for test site slots. We had a lot of simultaneous programs in the works of large engines and small 

engines. Test slots were an issue or a competition between various groups. There was another 

competition going on as well. It was organizational, between the quality and manufacturing 

departments. The environment was such that quality wanted to catch engineering oversights; 

or manufacturing wanted to catch engineering oversights. It was friendly, because we were all 

trying to win the race to the moon, but it was still a different environment.

I was the test engineer during engine development testing, and had to cope with the space 

race and its schedule pressure. I was assigned to assemble an SE-8 engine with a nozzle to 

get it ready to test. We accidentally installed an O-ring made of the wrong material. I tested 

the engine, and the seal leaked post-test. My name was Mudd. My boss, Cliff Hauenstein, 

who was the materials engineer, had to explain this to the vice president at Rocketdyne. This 

Rocketdyne vice president was a tough, old-school manager. He looked like Yul Brynner1

, and 

his management style was like Attila the Hun. It was not going to be a comfortable meeting. 

The vice president had jelly beans on his desk. He always ate jelly beans. He was a nervous guy. 

We walked into the office, and a jelly bean caught in his throat. He ran out, turning purple 

before the meeting started. Finally, he came back in, having coughed and dumped the jelly 

bean. But, we felt he must have seen God because he was extremely mellow. Everything was 

fine. “Don’t let it happen again,” he said. After that, my manager always had jelly beans in his 

office. It’s a true story, but the lesson learned was, “You know, these things are expensive. You 

ought to have a little better control and awareness, so you don’t make a mistake like that.”

We also had an uneven throat erosion issue caused by uneven injector performance. When we 

went away from the silicon carbide throat material, the substitute material we chose was not 

as sturdy. The SE-8 program was still schedule driven. We could identify injectors that caused 

high erosion problems by hot-fire testing each injector in a specially built thrust chamber that 

used steel throat inserts. A poorly performing injector would erode the steel throats quickly. 

The plan was to devise some sort of tool for inspecting throat erosion post-test to identify 

samples with good erosion characteristics and samples with bad erosion characteristics.

1

 Accomplished American actor who played several roles requiring a serious demeanor.

We came up with a 

single engine design 

with variable nozzle 

extensions and 

configurations to fit 

particular vehicle 

locations.
Tim Harmon, Rocketdyne - SE-7 & SE-8 Engines

 Apollo Rocket Propulsion Development 55

By this time, we probably had a database of fifty or so injector tests on steel throats of various 

discolorations or erosion patterns. I made up a board that basically consisted of samples that 

were good and samples that were bad. We put them on a metal strip. To do that, and do it in 

a hurry because we were in production, I went into the shop, found a welder, and said, “This 

is a ‘G job.’” In those days, a G job was special. It took no paperwork, and you could get 

anything done you wanted. So, we would weld a throat on a strip of metal. Now, it was hot. 

You had to wait for the first weld to cool down before you could weld the next one. Again, we 

had injectors at the Rocketdyne test site we were trying to inspect and show were acceptable. 

This tool was needed in a hurry. We would weld one on, then wait for the thing to cool. It was 

time-consuming, and I needed to solve the problem. There was a men’s bathroom right next to 

welding station. The welder would weld one strip, and I would run and flush it in the urinal 

and cool the metal weld immediately. We would then complete the next weld, flush it in the 

urinal, cool it right down, and repeat the process. (See Slides 12 and 13, Appendix F)

I had this nice board, all these welded pieces, delivered to the test site. I brought the tool out to 

inspection, and they were really happy with it. We fired the next injector. An inspector looked 

at it and said, “It feels all right to me. It looks just like this one on the inspection board.” 

However, the quality inspectors were unaware that the samples had been in the bathroom 

being flushed. I thought that was kind of payback for all the grief I got from the inspectors.

Another issue we had was the appearance of the thrust chamber post-test. This ablative material 

looked like a piece of polished wood before hot-fire testing. After we hot-fired it, it had ablated, 

so it looked a little rough, with something like the appearance of burned wood or charcoal. 

After one hot-fire test, the Air Force inspector said, “Well, it doesn’t look like this nice piece 

of wood, and we don’t know whether to accept it or not because it looks worn.” The solution 

to that was elegant, but silly. We changed the chamber inner diameter appearance to one that 

looked like it had been fired. We took this piece of the inner sleeve, and we charred it in an 

oven, so that it looked already used. Then, when it was hot-fired, it didn’t change in appearance 

and the inspectors thought that was neat. Now, that kind of fix for an issue was a waste of 

money. Everybody was happy, but it was an added step. Eventually, we convinced them that it 

was a stupid idea, and this fix was eliminated. That was sort of humorous, but it also was a sign 

that people were not working together or were not understanding the processes. Who was at 

fault? Nearly everyone – engineering, inspection, NASA, and the customer.

Those were two big issues and their solutions. 
Chapter Four

56 Remembering the Giants

Our test facility, Santa Susana (California), had an interesting history. A lot of B Western 

movies were made there before it became a test facility. If you see a 1920s or 1930s Western, 

there was a good chance it had been filmed at Santa Susana. Rocketdyne purchased the area 

for testing rocket engines and built many facilities. This was the facility that tested most small 

engines. It had four or five small engine test stands. It was a very active facility. Engineers were 

pushing for their test slots and trying to get things done. We had to test our engine at altitude 

conditions. We had a very limited capability in terms of altitude simulation testing. It could 

run approximately fifteen minutes before needing to be reloaded with propellants. To hot-fire 

calibrate an engine or run development tests, we conducted a test, analyzed our data, and 

determined our next test condition. If we wanted to change the mixture ratio and the various 

tank pressures to get another test, we tried to do that as fast as we could with our handy, dandy 

slide rule calculators. Times were different for testing in the 1960s than they are now. It was 

fairly frantic because we tried to get as many test conditions as possible in the shortest time 

span. What we did have were Datelink Independent Gateway Retrofit (DIGR) data recorders 

- we called them diggers - for instrumentation, which was interesting. The DIGR was a circular 

plot with an ink pen that showed the test data. A development engineer would run a test, go 

over and look at the DIGR, and it would tell the value of pressure or thrust or whatever. But, 

it wouldn’t tell right away because there was a hysteresis (or lag time) in the pen. We would 

have to go over and tap the DIGR to get our values. As soon as we ran a test – we had probably 

ten DIGRs we had to pay attention to – we were going along tapping every DIGR to get our 

value. It looked like we were doing some sort of comedy because we would tap all the DIGR, 

use our slide rules, do all the calculations, and then yell new test values to the test conductor. 

We would yell, “Change the tank pressures to this.” Then, we would do the next test, and start 

all over again. It was very frantic and very challenging. (See Slides 14 and 15, Appendix F)

When I was training to be a development engineer, I was mentored by a very seasoned engineer 

who had hired in six months earlier than I had been. We worked together for about two weeks, 

then I was on my own for running tests. The very first test we ran, the engine split in half. I 

could not believe it. I had a guilt trip that I had just ruined the space program. I got over it, 

but it was still traumatic. However, in those days, failures were somewhat acceptable because 

we were trying to develop engines, and it was exciting to be part of the space race. 

Development programs generated a huge amount of data. We had a lot of people doing tests. If 

one person wrote the test report on an engine, that was fine. That was not a problem. However, 

when you had twenty engineers conducting tests and writing reports, there was no consistency. 

What one engineer thought of the engine, another engineer would have a different slant. 

That turned out to be a major issue. The solution was to develop a simplified reporting process, a 

template system for reporting processes. (See Slide 16, Appendix F)

Times were different 

for testing in 

the 1960s than 

they are now. It 

was fairly frantic 

because we tried 

to get as many 

test conditions 

as possible in the 

shortest time span. 
Tim Harmon, Rocketdyne - SE-7 & SE-8 Engines

 Apollo Rocket Propulsion Development 57

Engine development issues were generally resolved through an extensive test program. That’s 

a lesson learned. For the next Apollo-type program, I think you have to be able to anticipate 

issues. Every application is slightly different, and the customer is slightly different. The customer 

wants to improve the engine. Many of those improvements cost development activities that are 

not necessary. It’s hard to say, but we certainly had an engine that was developed and worked. 

However, you have to anticipate there will be problems. I think that is the same situation you 

will find with the J-2X upper stage engine development program. That application is slightly 

different. Nothing goes quite off-the-shelf, but it’s better than starting from scratch. 

In summary, Apollo in the early days was certainly a challenge. The challenge was the fact that 

we were behind. As a country, we didn’t look nearly as smart technically as the Soviets. That 

was a challenge, and that was a driver. Using the off-the-shelf approach saved time. It leveraged 

previous technology. However, we also needed to work together. The goal was the same for 

engineering, for quality, for inspection, for the customer, and for NASA – work together. 

---------------------------------------

Editor’s Note: There were no questions taken at this time. The question-and-answer session was 

held after Mr. Harmon’s second presentation, the transcript of which appears in Chapter 7. 

Engine 

development 

issues were 

generally resolved 

through an 

extensive test 

program. 

That’s a lesson 

learned. 
58 Remembering the Giants 

AJ10 Service Propulsion Engine (Source: Clay Boyce’s Presentation Viewgraphs, 

Appendix G)
Apollo Rocket Propulsion Development 59

Clay Boyce left the U.S. Air Force in 1955, and joined Aerojet 

at its Azusa, California, facility. After the launch of Russia’s 

Sputnik Program, he joined the Thor/Able Second Stage 

Program, an adaptation of the Vanguard second stage. 

Subsequently, he participated in the conceptual design 

and development of the AbleStar upper stage. Variations of 

this stage are still flying. In 1960, he was assigned to the 

Win Apollo team and, then, became engineering manager of the Apollo Service Propulsion 

System (SPS) engine development. In 1969, he was assigned to NASA’s Johnson Space 

Center in Texas for technical support of the SPS engine. Later assignments included the 

Space Shuttle Orbital Maneuvering System engine, the Japanese N-2 Upper Stage, and the 

National Aerospace Plane. Boyce retired from Aerojet in 1991, and has provided consulting 

services to the company since that time.

Some of the most tense times I ever had were on the Apollo mission. The question I always 

got asked the most when somebody talked about it was, “What if the service module engine 

doesn’t start, especially when the astronauts are ready to come home?” I will tell you why it was 

tense. I knew it would start, but until that first mid-course firing on the way to the moon, I 

didn’t know if they had connected it right. Once I got a firing signal, I knew they had mated 

the connectors properly in place.

Our Apollo proposal effort started for me in 1960. The Thor AbleStar vehicle was being developed then. (See Slide 3, Appendix G) We used technology like the titanium nozzle for the 

Chapter Five

Clay Boyce

Aerojet - AJ10-137 Apollo Service Module Engine
Chapter Five

60 Remembering the Giants

SPS engine. I was partying in a swimming pool down in Florida just after the second AbleStar launch when they came and got me about 3 a.m. and said, “We booked you a flight out 

of Orlando, 7 a.m. Be in Philadelphia, go to General Electric, find the Apollo group, and tell 

them our engine is the best.” That was my first introduction to the word “Apollo” in the space 

program.

We had written twenty-nine proposals to twelve different primes (contractors competing for 

Apollo), and all but one of them selected our engine. We had a pretty good chance of winning 

the job. The engine was the same height as a space shuttle main engine, and the exit diameter 

was slightly larger than the space shuttle main engine. (See Slide 5, Appendix G) The size was 

set before Apollo adopted the Lunar Excursion Module (LEM) mission approach. Prior to the 

LEM, the whole Command Module was to be landed on the moon, and this engine would 

have been the engine to lift it off the moon and return it to Earth. I would have had to wait a 

long time for that first firing, if they had followed the original concept. Aerojet was awarded 

the Apollo SPS engine development contract in April 1962, and the final decision on LEM 

was made about November or December of that year. We had proceeded far enough that 

even though they went to the LEM concept, they decided it wasn’t worth starting over with a 

smaller, lower-thrust engine, so they kept the large size. The large size resulted because it was 

specified to be a pressure-fed engine. They didn’t want to have to worry about the reliability 

of pumps and everything on the moon at that time. It was a 20,000-pound-force (lbf) thrust 

engine, but it was as big as a space shuttle main engine (SSME). Chamber pressure was only 

100 pounds per square inch (psi). The major change that it made from the mockup was at the 

head-end, gimbaled configuration was changed to a throat-gimbaled design, which reduced 

the overall height of the Saturn vehicle when they put the LEM in behind the service module. 

The throat gimbaling of the engine saved about three and one-half to four feet of overall vehicle 

length and weight.

The technology we had at that time was the Saint/Apollo subscale engine. (See Slide 4, Appendix 

G) It was for a satellite interceptor system for the Air Force. It used what we called earth-storable propellants, N2

O4 (dinitrogen tetroxide, a powerful oxidizer) and some of the hydrazine 

families, because the Saint Program actually was an in-orbit satellite that was supposed to be 

able to detect an orbiting enemy satellite and go after it. That program was started in 1957-

1958. It never flew because the computer capabilities and the control system it required just 

weren’t mature enough to do it. But, we did build the ablative engine, the biggest at that time 

with 2,000 pounds of thrust. It had a nice titanium nozzle, an aluminum injector, and an 

ablative thrust chamber–all the things we needed for the SPS because we were looking for the 

simplest, lightest-weight engine we could get. We later used that engine for subscale testing of 

ablative materials and nozzle extensions for the full-scale engine. 
Clay Boyce, Aerojet - AJ10-137 Apollo Service Module Engine

Apollo Rocket Propulsion Development 61

The general configuration of the SPS engine was 20,000 pounds of thrust, with a chamber 

pressure of 100 psi and specific impulse (Isp) of 314.5. The very large nozzle had an area ratio 

of 62.5:1 (exit area to throat area). The propellants were nitrogen tetroxide (also known as 

N2

O4 and nitrous oxide) and A-50. A-50 was a hydrazine family fuel. Aerojet developed it for 

the Titan Missile Program when they went with Titan II, to store it in the launch silos. They 

wanted the highest performance they could get. N2

H4 was just pure hydrazine, which doesn’t 

take low temperature very well. In fact, it freezes about like water. We started adding unsymmetricaldimethylhydrazine (UDMH) to the hydrazine until such time as it would meet the 

environmental specifications the Air Force needed for Titan II. It turned out it’s roughly a 

fifty-fifty mix. We still had to be careful with that fuel because the two fluids didn’t mix very 

well chemically. We had to spray the two fluids through some special nozzles to get them to 

emulsify with each other into a single fluid. If we ever got it too cold or froze it, the hydrazine 

separated back out. Then, if we tried to run the engine, things could go boom in the night.

The inlet pressure was only 165 pounds per square inch absolute (psia), but we needed at 

least forty psi pressure drop across the injector just to get some kind of stable flow. It was a 

whole new game for some of us. We didn’t have much supply pressure to work with. It had the 

aluminum injector to keep the weight down. That was a couple feet in diameter, and we didn’t 

have a lot of propellant to cool it. In fact, we had to use both propellants to keep the injector 

cool. There were twenty-two ring channels in the injector. Specification required 750 seconds 

duration, or fifty engine restarts during a flight. 

There were several first flight things we accomplished with the engine. It was the first ablative 

thrust chamber of any size to fly. (See Slide 6, Appendix G) There were no liners in it. It was 

just straight ablative material. It took us a while to figure that out. It was a throat-gimbaled 

engine, and it was the first engine to fly with columbium (also known as niobium, used as an 

alloying element in steels and superalloys) in the nozzle. 

The first time we fired that assembly, we got a couple of surprises. (See Slide 7, Appendix G) 

The chamber pressure was supposed to be 100 psi. We fired it up, and it ran just fine for about 

a second, then the chamber pressure (Pc) dipped down to forty psi and gradually came back 

up to 100. The engine went on running, but we were holding our breath a little. One of our 

requirements was to use the fuel to actuate the propellant valve. With those low pressures, it 

pulled so much fuel from the inlet of the engine to actuate the valve that there wasn’t enough 

to run the engine for a second or two. The good news was it stayed stable through all of that. 

During these very first firings, we hadn’t received the requirement for dynamic stability, so the 

engine contained an unbaffled injector, and it recovered. It never really ever went unstable. It 

just turns out it was fairly throttleable. 

It was a throatgimbaled engine, 

and it was the first 

engine to fly with 

columbium in the 

nozzle.
Chapter Five

62 Remembering the Giants

There was another unexpected surprise that involved the interface with the actual Service 

Module. The decision for our approach evolved during proposal time. Half of the contractors 

looked at a single engine with redundant controls; others had groups of four and five engines 

for reliability. (See Slide 8, Appendix G) The final contractual decision, from a weight and an 

efficiency standpoint, was to make a single thrust chamber with dual redundant valving to 

assure that we could always get a start, and always get a shutdown. The gimbal actuators, which 

were the other moving parts on the engine, had semi-redundancy. The propellant valve was 

about two feet square and weighed 100 pounds, the biggest valve on which I ever worked. It 

was kind of a mirror image. One side of it was the oxidizer; the other side was the fuel. There 

were four ball valves in each propellant circuit in this arrangement: a fuel ball, a common 

shaft, and an oxidizer ball. The actuator to rotate the shaft was located between the balls. 

There were four of these individual assemblies utilized into a series-parallel configuration. On 

an engine start, any one of those actuators could fail, and the engine would still start. If, on 

shutdown, one of the actuators failed, it would still shutdown. We could run the engine, which 

we normally did with all four sets of valves open, but it would run equally well on one pair in 

each circuit, which we did a lot of in qualification testing.

One of the development challenges we faced didn’t have to do so much with the valve operation as it did with manufacturing. It was a complex casting with lots of machining. We were 

having them machined outside our plant by a subcontractor, and it was taking him about a 

week to make one casting. Somewhere in the middle of trying to get more castings, Aerojet 

purchased what, at that time, was one of the newest numerically controlled machines with 

multi-axis operations. Management decided they needed something to do with it, so I got 

directed to bring the machining of the valve in-house. We did that, and it took a couple of 

months to get the things all set up. We started machining valves. On the fourth day of the 

six-day cycle, two shifts a day, the machine tool would jam a tool through a side of the casting 

and ruin it. At that time, the machine control was not digital. There was a punched paper tape, 

about one inch wide and over 100 feet long. Every time they would ruin one of those castings, 

they would have to go through that whole tape, make a few changes, and punch out an entire 

new tape. It was three and one-half months before we started getting good castings. After that, 

it worked fine, but it did set us back a little on our schedule.

The other major valve problem was the actuation. We were required to use engine fuel as the 

hydraulic actuator fluid. They wanted to minimize the number of connections that crossed the 

interface between the Service Module and the engine. We were only allowed the main propellant lines and one redundant, electrical cable. That was it. We switched to a pneumatic valve 

actuation because we had to stay in the same envelope, and we couldn’t cross the interface. We 

removed the fuel hydraulic system from the cylinders and put a big spring in there that was 

used to close the valve and an engine-located, high-pressure nitrogen subsystem to pneumati-
Clay Boyce, Aerojet - AJ10-137 Apollo Service Module Engine

Apollo Rocket Propulsion Development 63

cally open the valve. We were required fifty starts of the engine. The amount of nitrogen we 

had would restart the engine about eighty times. (See Slide 8, Appendix G)

Those were the two major events during valve development. Once we got them going, we had 

the usual problems of seal material in the valve. We started out with Teflon, and it would coldflow too much. We finally used some fiberglass-impregnated Teflon materials that would work. 

We had a three-micron finish on those valve balls, and they were each about two and one-half 

inches in diameter. It was a job manufacturing them, but not intolerable. The balls were made 

out of stainless steel. During one of our weight reduction exercises, we got the bright idea to 

make the balls from some lighter-weight material. We tried beryllium, which was about the 

only candidate that was significantly lighter. They worked, except that beryllium is pretty soft, 

and it didn’t have wear life that would meet the mission requirement.

Our injector assembly was about two feet in diameter. It was all aluminum. (See Slide 9, 

Appendix G) It had five baffles for the dynamic stability, and I think we did 240 or 250 bomb1

tests in developing this baffle configuration. The original injector had twenty-two rings. There 

was one main rule in the fabrication of the injector: we had to keep parent material between 

the propellants. There couldn’t be any weld that both propellants could touch. With the low 

pressures and the big surface area to cool, we didn’t have much room left for lands between 

the rings. That was a problem. When we went to the baffles, we had to reduce the number of 

channels down to fifteen. When we started injector development, because of the narrow land 

area, we thought we would have a welding problem, so we decided we would braze the injector. 

Well, it turns out there were about 200 linear inches of area that had to be brazed. If we didn’t 

have it perfectly cleaned and perfectly aligned,we didn’t get a perfect braze. We never got a 

brazed injector that was 100 percent leak-free. 

About that time, NASA was still looking at the future Nova launch vehicle, and Aerojet had 

the contract for the M-1 engine for the Nova’s second stage. It was 1 million-plus-one-pound 

thrust hydrogen/oxygen engine. Aerojet had bought an electron beam welder for use on the 

program. When the M-1 got canceled, it sat in a corner, doing nothing. We decided we would 

give it a shot at welding the SPS injector, and it worked quite well. After we started firing the 

injectors for long periods of time, we encountered a problem that became a phenomenon in 

very low-pressure, large engines with the particular propellant combination we had. We named 

it “popping.” We would start the engine up and we never knew when, but if we were watching 

the chamber pressure (Pc) trace on the oscillograph, all of a sudden, the trace would expand 

1

 Here the ‘bomb’ refers to an impulse charge to induce instability in the combustion chamber for assessing the 

engine’s susceptibility to combustion instability.
Chapter Five

64 Remembering the Giants

into a little football-like bubble. Then, it would go away. It would just last a few milliseconds. 

But, every time it would make a bubble, to us in the test bay, it sounded like a “pop;” hence, 

the name “popping.” Eventually, it became a large-enough concern that NASA convened 

a meeting at North American Aviation, Inc., and brought all the stability experts from the 

universities, and we spent two days talking about that. I thought, “Boy, I’m finally going to get 

some help.” The next day, I got the results. They told me they didn’t like the name “popping.” 

We eventually solved it. At the low pressures and the low pressure drop across the injector, the 

length of the drilled orifice became critical, and the propellant stream sometimes would come 

out clean, and sometimes would attach to the edge of the orifice, causing some propellant to 

flow onto the injector face. When the propellant stream ran along the injector face, it changed 

the combustion characteristics of the element, and it would cause the pop. It never hurt the 

engine, but the first few times really worried us. We counter-drilled those and got a shorter, 

fixed-length orifice that solved the problem.

We had to do the bomb tests to check the dynamic stability. That turned out to be relatively 

easy to do. We had to worry more about damage to the injector from bomb fragments than 

from anything else. The ablative chamber had no kind of liners or anything inside; it consisted 

only of laid-up ablative material with some fluted aluminum flanges at each end. (See Slide 10, 

Appendix G) At the time we started the program, and even back in the old Saint Program, we 

didn’t have nice impregnated tape like they do today. The coated glass fiber string material was 

made in whatever length rolls you wanted. The continuous string was a little bit bigger around 

than a piece of heavy kite string. Whatever thickness chamber we wanted, we’d cut that length 

of strings and lay them on a piece of sticky tape. All of them were hand laid, side-by-side-byside. The tape with the strings was then wound around the chamber mandrel (a spindle to 

support the piece during machining) and impregnated. We did that until the ablative material industry progressed and developed the flat tape. The flat tape made the chamber lighter 

because the ratio of the glass to resin was a lot greater, and provided more heat-dissipating 

capacity. This allowed the chamber wall to be thinner. 

The next major component on the engine was the gimbal actuator. This particular actuator 

is a semi-redundant device. I say “semi” because the main actuator movement device was a 

single ball screw, much like that used in present-day car steering mechanisms. The ball screw 

was operated by two electric motors transmitting power through magnetic particle clutches. 

One motor/clutch assembly extended the actuator, and one retracted the actuator. There was a 

completely redundant set of motors and clutches in each actuator for mission reliability. During 

engine operation, all four motor/clutch assemblies were running, and control signals from the 

spacecraft activated a specific magnetic clutch for the desired engine position. A couple of our 

problems with the actuator didn’t have anything to do with technical accomplishment. At 

the time we got the contract for the SPS engine, we planned to subcontract the actuator to a 

We had to worry 

more about 

damage to the 

injector from bomb 

fragments than 

from anything else.
Clay Boyce, Aerojet - AJ10-137 Apollo Service Module Engine

Apollo Rocket Propulsion Development 65

company called Lear Inc. in Grand Rapids, Michigan. Lear was the biggest actuator company 

in the U.S. at that time; all Boeing and McDonnell Douglas Aircraft Company airplanes used 

Lear actuators for moving controls surfaces and flaps. They had the best reputation for that. 

We gave them a contract, and they went to work. 

The actuator assembly had to be enclosed in a pressurized, hermetically sealed can. (See Slide 

11, Appendix G) The electric motors and the magnetic clutches wouldn’t work in the vacuum 

of space. After a year or so, we received a pair of actuators to start testing. Both units were 

overweight and larger than the required envelope. Right about that time, another company by 

the name of Siegler bought Lear. The Lear group kept working the size and weight condition, 

along with a stiffness problem. Then, Siegler bought another company called Jack and Heintz 

Motor in Cleveland. Well, the Lear plant was overloaded with work; they didn’t want to build 

a new building, and the Jack and Heintz plant in Cleveland had lots of empty space. Lear/

Siegler moved the actuator program to Cleveland. This was a problem because the actuator 

engineers wouldn’t move to Cleveland. All of a sudden, we had this contract with about half 

the money spent, but we didn’t have a part, and we didn’t have an engineering staff anymore. 

About that time, North American Aviation and NASA were getting nervous enough that they 

decided we’d better find another actuator company. Meanwhile, a couple of people who had 

worked at Lear had gone to work for an actuator company down in Costa Mesa, California, 

called Cadillac Gage. They were building actuators for the nozzles on the Polaris missile under 

contract from Aerojet, and had produced good hardware. The ex-Lear people talked to the 

Cadillac Gage management, and they gave us a bid to build the actuator. At that point, NASA 

and North American interceded and said, “We have already spent half the money. We can’t 

afford that again. You are directed to give a fixed-price contract.” We gave them a fixed-price 

contract, and they started to work. 

Then, the second problem we found on the Lear actuators resurfaced. The stiffness requirement for the actuator was to be a spring rate of 300,000 inch-pounds to meet the vehicle 

control system loop frequency. The stiffness on the two actuators we’d received from the old Lear 

Company and the first two from Cadillac Gage were well below requirements. One of them 

was about 220,000 inch-pounds; the other one was 160,000. By then, the second company had 

used up all their fixed-price money and a little bit more. They said, “Sorry, but we have to stop 

work.” It became a contractual battle for a while. It eventually was resolved. About that time, 

back at NASA Headquarters in Washington, D.C., Joe Shea, then the chief engineer of the 

Apollo Program, was reviewing his overall Program Evaluation and Review Technique (PERT) 

chart, and the actuator popped up as the umbrella for the whole Apollo Program, because if we 

didn’t have that, the vehicles didn’t go. Shea came to visit us one day. We went through everything with him. The North American guys were there and everything. After the review, they all 

went back to North American and said they would call us with their recommendations. 
Chapter Five

66 Remembering the Giants

The next afternoon I got a call that said, “Don’t bother to come down. The problem is solved.” 

I said, “What happened?” My 300,000-pound actuator was being tied to a 90,000 inch-pound 

spring rate bracket on the Service Module. It didn’t make any difference as long as my spring 

rate was over 100. By then, we’d worked for almost two and one-half years trying to get that 

thing stiff, so we hadn’t performed life testing on the motors and magnetic particle clutches. 

The clutches became the next challenge. There had never been a magnetic clutch that big. 

Existing technology were little things, less than three-quarter inches in diameter. This one was 

not quite three inches in diameter. As you might expect, we found we had overheating and 

other problems. We got it all done, but the problem illustrates the fact that sometimes you run 

into events that have nothing to do with your engineering capability. My final challenge out of 

the actuator development was the boss saying, “Okay, go get out of that fixed-price contract.” 

If you’ve read government contracting regulations, you know there’s no way you could do that. 

I had to figure out a whole different contracting scheme so we could get Cadillac Gage back to 

work. (See Slide 11, Appendix G)

The first nozzle for the Service Propulsion System engine was developed out of Saint and 

AbleStar technology. The nozzle was made out of titanium. We calculated that the area ratio 

at the head end was about 6:1 as the point where exit temperatures would be down to titanium capabilities for radiation cooling. The exit area ratio was 62:1. The first problem came 

with fabrication of the gores. It takes sixteen gores to make one nozzle. Our nozzle fabrication 

contractor’s plan was to hot stretch-form the titanium, make a mold of the shape we needed, 

and use Calrod® units (elongated heating elements) to make it hot. That, in turn, would heat 

the titanium stretched over it. That didn’t work too well. We were using titanium for the nozzle 

because it took lots of temperature. We had to get it very hot to stretch-form it. Titanium has 

a tremendous memory. It doesn’t want to stay in a new place. It kept wanting to go back to its 

original flat shape when it cooled. We were in a quandary about that. We had only been able 

to make four gores that were even halfway close. We gave them to our subcontractor who was 

developing the technique for welding them all together. We were scratching our heads about 

what to do when one of those fortuitous events occurred. (See Slide 12, Appendix G) 

About the same time we started scratching our heads, the then-president of our company was 

at a cocktail party in Washington, D.C., and he got to talking with the good lady senator from 

Maine. She said, “Dan, I gave you a bunch of money, supported you for that Apollo thing, 

but you’re not putting any of that work in my state.” I then got a memo that said, “Go see if 

they do anything in Maine that’ll help your engine program along.” I put all my problems in 

my briefcase and went and found a little guy in Maine who worried about the state’s economic 

development. He listened to me for about half a day and said, “We’ve got nothing like that up 

here. But, two weeks ago, there was a guy in here wanting us to help him start a small fishing 

boat business. He wants to make them out of aluminum, and they’re all nice and round, instead 

We were 

scratching our 

heads about what 

to do when one of 

those fortuitous 

events occurred.
Clay Boyce, Aerojet - AJ10-137 Apollo Service Module Engine

Apollo Rocket Propulsion Development 67

of square. He had a machine that makes stuff kind of round and long like you’re describing.” I 

got the boat maker’s name and an address, which was down in Massachusetts.

When I got to the address, it was at a road crossing with a little building about the size of an 

old outhouse sitting there and that was it. Finally, about a mile away, I found an old cotton mill 

from back in the textile days, and here was this little, old Norwegian toolmaker, and he had 

built this big machine. He was making boat parts, specifically the double-curved part of the 

bow. Well, he hadn’t been able to sell the boat business, so he had taken a couple of contracts 

to make gores for radar antennas. 

Those gores, curved two ways, weren’t too different from what was needed for the nozzle. I 

described what I needed and he said, “Yeah, I think I can do that. Send me a mold and send 

me some material.” We did. Probably a month or six weeks later, he called and said, “Come on 

back. I’ll show you what’s going on.” We went back to his shop in Massachusetts, and he had 

been working with his machine, and he had made a bunch of gores out of aluminum while he 

was working. So, I was going to be there for the first titanium pull. When we had tried the hot 

stretch-forming, we had to work with one big sheet of titanium. The boat maker’s new process 

worked so efficiently, he could cut the sheet diagonally and we could get two gores out of every 

sheet. He laid the sheet out on his machine, brushed on a lubricant from a nearby pot sitting 

on a hot plate, and started the machine. It was a cold-working process, and the machine put 

that titanium through something like old washing machine washer rollers. He could vary the 

axis of the rollers and their distance apart. The machine put that titanium through a little S 

curve about two inches high; it took away all its old form memory and put in a new one. 

The very first one that rolled out the end of that machine looked perfect to me, better than 

any I had seen. We laid it on the mold, and he got a little hammer and went all over it. “Well, 

it’s not right. There’s one little place that sounds a little different,” he said. He went over and 

tweaked his machine, put the same part back in there, ran it through again and had a perfect 

part. I said, “How many of those could you make a day?” He replied, “I can probably make 

twenty of them in a day.” I said, “How much?” He said, “Fifty bucks apiece, if you supply the 

metal.” That was the last gore I got from him for three and one-half months because his business didn’t exactly meet NASA 200-2 quality specifications. I had to create a Quality Department, a Purchasing Department, all of the “isms” and documentation to certify the parts were 

useable for nozzle fabrication. But, by then, that problem was solved, so I didn’t worry about 

it anymore. 

We finally got a nozzle together, and because of the size and the pressure, the only place that 

you could hot-fire test it was in the big Air Force altitude facility at Tullahoma, Tennessee. I 

talked recently to a gentleman who works at Arnold Engineering Development Center, and 
Chapter Five

68 Remembering the Giants

we decided the tests were conducted at the facility’s J-3 cell, a very tall structure, that was used. 

I found out later when we got the test cell operating satisfactorily, that firing the SPS engine 

caused the test stand to flex and caused misalignment of the performance measuring instrumentation. The thrust and side forces measuring cells had to be biased for the thrust stand 

movement so we could get accurate test data. Around early 1964, we got the first altitude 

engine all together and set up at Tullahoma. For the first engine firing, I was sitting in the 

control room with one of the North American Aviation guys. We heard the big countdown 

going on, and they finally said, “Fire.” We were looking at our oscilloscope and nothing seemed 

to be happening. We thought, “Man, that engine runs awfully smooth.” Then, somebody 

said, “Okay, fire that engine,” and they threw the SPS start switch. It turned out the whole 

countdown was for the rocket engine injector they were developing to pull the tunnel down to 

altitude. That was the big thing to them. We were going to fire the SPS engine for fifty seconds 

as a checkout. We noticed a little decay in thrust level after ten to fifteen seconds. But, then, 

it leveled right off and everything was fine, and we thought we had a successful run because 

everything looked very good. 

As expected, the nozzle got very hot. Things finally cooled down, and we were able to get in 

the cell. It had gotten hot all right – too hot. The whole nozzle had moved forward about 

eighteen inches. About ten inches down, the nozzle had curled right back over itself. We had a 

nice, big S curve all around and a nozzle that was a foot and one-half shorter. That’s when we 

figured out we needed columbium. The only reason we found columbium is when titanium 

wouldn’t work, we went to the periodic table of the chemistry book and said, “Okay, what’s 

the next lightest metal?” We determined it needed to handle at least 2,100 degrees Farenheit. 

We were running around 1,900 degrees at the nozzle attach point, and columbium appeared 

to be capable. Then, we tried to find some columbium. It turns out the nuclear power industry 

used columbium to make the rods in which they put the nuclear material for power plants. We 

went to the company that was making those rods.

The company was called Wah Chang, in Albany, Oregon, and was owned by a very nice 

Chinese gentleman. We wanted thin material. All the material he made was one-half inch or 

more plates. He thought he could make us some thinner stuff. We wanted it forty-thousandths 

(0.040) of an inch thick. He was able to produce the thinner sheet stock we required. The 

titanium nozzle had a machine flange on the front. We tried to follow the same design. We 

were going to make a flange out of columbium also, but nobody had ever made anything like 

that out of columbium. They had always used the flat roll stock. But we found a company that 

would try to make a casting. They had a big, spin-casting rig. That flange was around twenty 

inches in diameter. After making a casting, it turned out they never had machined columbium, 

and it is extremely tough to machine. They could get about two cuts around the flange with 

one, good silicon carbide tool, then it was shot. We later got rid of the flange and figured out 
Clay Boyce, Aerojet - AJ10-137 Apollo Service Module Engine

Apollo Rocket Propulsion Development 69

an attachment where we could just roll a lip on the end of the nozzle sheet material and clamp 

it to the end of the ablative chamber. While they were busy making sheet metal stock for the 

full-scale nozzle assemblies, we got a small piece of columbium and just rolled it up to make 

a small regular conical nozzle. We put that on the Apollo subscale engine and took it down to 

Tullahoma to see how the columbium worked. We fired a partial simulated SPS duty cycle. 

That was 200 seconds firing and one half-hour coast, then re-fire. Everything went fine until 

the re-fire, when the thrust level was not where it should have been. We opened up the tunnel, 

and the nozzle was gone about two inches down from the flange. It was just all jagged.

We determined that the temperatures to which we raised the columbium caused it to absorb 

hydrogen from what little atmosphere there was in test cell. The hydrogen crystallized the 

columbium. The inside of the nozzle, next to the fire, was clean and nice–no penetration there. 

But, on the outside, you could see a crystalline surface. The engine re-starting shock had shattered the nozzle.

 

Operating the engine out in space would have been fine, no hydrogen. But, we had to develop 

and qualify the engine here on the ground. We had to find some way to protect the columbium 

so it could pass all the tests. We went through about twenty different coating materials and 

finally found one that North American Aircraft Division was using in some of the jet engines 

exit areas. We still had the same problem. We had to get the columbium up to 1,800 degrees 

to bake the coating on. There were ovens big enough to get up to the temperature we needed, 

but none of them had vacuum capability. Finally, working with North American Aviation, we 

developed a big retort (a closed laboratory vessel) that we could set the nozzle in, weld it all 

closed, pressure-test it, put in an inert atmosphere, and, then, put that whole business in the 

big oven. Remember, that nozzle was about ten feet tall and about eight feet in diameter. Facilities with the capability to heat and bake the large assembly for several hours were not readily 

available. Final nozzle configuration used columbium down to about the 40:1 area ratio, then 

titanium the rest of the way to the end. There were some quite unique issues in developing 

the welds between columbium and titanium. We welded them together using titanium rod. 

Micrographs of weld samples looked like there was no bond between the columbium and the 

titanium. Columbium material surface appeared just as smooth as it was before welding. But 

whenever we performed pull tests of joint samples, failure occurred in the weld-heat affected 

zone of the titanium. Whatever went on in there was strong. It worked well. 

The SPS engine flew nineteen times. (See Slide 13, Appendix G) The most starts probably 

were performed on three of the engines, eight starts per engine, with 6,000 seconds of total 

time. In the overall development program, I know we tested more than 200 injectors. I think 

we had about 230 bomb tests. We went through 124 different injector patterns to get one that 

was totally compatible with the ablative material without any liners, and still approached the 

We determined that 

the temperatures 

to which we raised 

the columbium 

caused it to 

absorb hydrogen 

from what little 

atmosphere there 

was in the test 

cell. The hydrogen 

crystallized the 

columbium.
Chapter Five

70 Remembering the Giants

98 percent combustion efficiency we needed without making any grooves down the ablative 

chamber. All in all, the total testing during the development and qualification totaled about 

twenty-eight hours on the engine assembly. That’s a long time for that small engine. 

Just before the launch of the Apollo 8 mission, which took the first men out around the moon, 

Tom Paine was the acting administrator of NASA. He called George Mueller, who was the 

associate administrator for manned space flight. He said, “George, are you sure that engine is 

going to start? Find out.” Mueller sent a team off to study the engine. The results of his study 

were the reason I knew the engine would start. He said, “There have been over 3,200 starts of 

that engine in development and qualification. There were only four times it didn’t start, and 

those were all due to test stand stuff.” The engine had started every time all through the development and qualification program. When I used to go to my little hometown in north Idaho, 

which was a long way from the space program, some of guys back there would ask me what 

I was working on. I told them, and they said, “What if it doesn’t start?” I said, “It’ll start. I’ll 

lay it out for you. You probably start your car an average of three times a day. In a year, you’ve 

started it a thousand times. I’ll bet you at least once, it doesn’t go hmmmm, and, then, it’s going 

to grind a little. That’s one failure in a thousand.” The specification I was working with was one 

failure in 5 million starts. So, I knew it was going to work as long as North American Aviation 

put that electrical cable on correctly. 

---------------------------------------

Editor’s Note: The following information reflects a question-and-answer session held after Boyce’s 

presentation. 

QUESTION: You had your fuel and your oxidizer in one valve assembly; would you do that 

differently now or did that work well enough?

BOYCE: It worked very well because the two independent castings had a cavity in between 

them. We had dual seals on each shaft where they went into the cavity/actuating area, and that 

area was vented. There never was a problem there. The biggest difficulty they had with respect 

to propellants was decontamination between the series ball valves after engine testing. There is 

a similar arrangement on Aerojet’s Orbital Maneuvering System engine on the space shuttle. 

There is not a bi-propellant issue. Between the two series valves, there’s always a little propellant trapped on landing, and they have to decontaminate those areas before returning to the 

hangar. Obviously, we didn’t have that situation after an SPS engine flight. Nobody got that 

near to it.

All in all, the total 

testing during 

the development 

and qualification 

totaled about 

twenty-eight hours 

on the engine 

assembly. That’s a 

long time for that 

small engine. 
Clay Boyce, Aerojet - AJ10-137 Apollo Service Module Engine

Apollo Rocket Propulsion Development 71

QUESTION: Out of those 6,400 seconds, what percentage would you say was development 

versus qualification?

BOYCE: I would say probably 50 percent to 60 percent of those seconds were logged in 

development prior to starting qualification. There were some additional development tests in 

parallel with qualification testing to evaluate issues that arose during qualification. Here’s a 

little summary of development and qualification: 216 injectors in the program, 3,200 firings, 

and 700 minutes of time on injectors; nozzle extensions, thirty-one units made, 1,400 firings 

for a total of 700 minutes. We made 274 combustion chambers, with 1,400 firings on the 

thrust chamber, and about 900 minutes of test time. We had twenty-seven total engine assemblies and 4,000 firings of the all-up engine assembly, which totaled about 1,700 minutes.

QUESTION: Early on in the design concept, did you have abort requirements already factored 

in or did those come later? That is, could you use the engine as an abort?

BOYCE: I think in one of the first Earth orbit manned launches, there was an early shutdown 

problem with one of the Saturn stages; they fired the SPS engine to put the Command/Service 

Module into orbit. It wasn’t planned to be that way. During Earth orbit missions, the SPS 

was fired for engine operating characteristics tests, orbit corrections, and de-orbit. Remember, 

the original concept was that it would never fire until the astronauts were ready to come 

home from the moon. It did, of course, take over midcourse corrections; lunar orbit insertion 

and Earth return orbit insertion after they went to the LEM concept. After adopting LEM, 

there were no specification changes that changed the engine capability of fifty starts and 750 

seconds total firing time. They did change the mission duty cycle firing schedule to match 

the LEM scenario. During the qualification program, an “abort” firing of 610 seconds was 

demonstrated. Worst-case mission requirements were thirty-six starts and 546 seconds of burn 

time - well within original specification.

STEVE FISHER: So, apparently it wasn’t part of the requirement, but the capability was there 

in terms of numbers of seconds. 
72 Remembering the Giants

(NASA image number: MSFC-0101206)
Apollo Rocket Propulsion Development 73

Gerard “Jerry” W. Elverum began working at California 

Institute of Technology’s Jet Propulsion Laboratory (JPL) 

in 1949, where he spent ten years performing pioneering 

research and development on propellants and rocket 

propulsion. In 1959, he joined Space Technology 

Laboratories (later called TRW). In May 1963, Grumman 

and NASA selected his patented design concept for 

a deep-throttling liquid bi-propellant rocket engine for a backup development program 

for the descent engine of the Lunar Excursion Module (LEM). In December 1964, NASA 

committed to the Space Technology Laboratories design, and the first flight engine was 

delivered to Grumman in August 1966. Elverum was program director and chief engineer 

for the LEM descent engine throughout this time. He joined the American Rocket Society 

in 1951, received the American Institute of Aeronautics and Astronautics’ (AIAA) James H. 

Wyld Propulsion Award in 1973, and was elected an AIAA Fellow in 1983. He was elected 

to the National Academy of Engineering in 1987. He retired as vice president and general 

manager of TRW’s Applied Technology Division in 1990.

Chapter Six

Gerard Elverum

TRW - Lunar Descent Engine
Chapter Six

74 Remembering the Giants

The rocket engine companies Rocketdyne and Aerojet were the big names back in the late 1950s 

and early 1960s; they got the Apollo astronauts into orbit around the moon. But, how in the 

world did it happen that those guys were sitting in the descent module of the Apollo Program 

on top of an engine that was developed by a few engineers at a space engineering company called 

Space Technology Laboratories (STL) instead of one of the big rocket companies?

In 1960, Space Technology Laboratories was a system engineering company that was doing 

engineering and technical direction on the Atlas, Titan, and Minuteman programs. But, at 

that time, they decided they also had to get involved in space technology because they were 

beginning to build military spacecraft. One of the things they decided was that in order to 

be able to make certain maneuvers in space, and to do some of the things they wanted to do 

with missiles, they were going to have to come with a deep-throttling rocket engine to perform 

those missions. At the time, I was the head of the Advanced Propulsion Development Department at STL. They came to me and said, “We need a 20:1 throttling engine. It has to be storable. It’s going to be pressure-fed. It has to be stable, and we need it in order to move into the 

next phase.” So, I got that job. But, the story of the Lunar Excursion Module descent engine 

really starts back many years before that at Cal Tech’s Jet Propulsion Laboratory (JPL).

At JPL, I was playing around with a pair of concentric capillary tubes. The reason I was doing 

that is that we had responsibility for developing the Corporal, which was the Army’s intermediate, tactical, ballistic missile. At the time, they decided they had to lower the freezing point, 

so they went to something called stable fuming nitric acid (SFNA) from red fuming nitric 

acid (RFNA). The SFNA, we put a lot more NO2 , and then, we put some water in. When 

we did that, the Corporal engine got very rough. It had all kinds of performance problems. 

They turned to me and said, “You get into your laboratory and see if you can figure out what’s 

going on.” Well, I concluded that the liquid phase reactions of hypergolic propellants were 

really controlling what was going in the combustion of that engine, and that when we put the 

water in, we stopped the early reactions that led to the distribution of the propellant from the 

injector we were using on the Corporal. As a result of that, they no longer were mixing the way 

they should. They were being separated, and then, they were detonating in certain parts of the 

engine, and it was causing it to be rough. That meant they had to change the injector considerably and back off on how we were doing it. (See Slide 7, Appendix H)

At JPL, I was working on the liquid phase reactions with concentric tubes, and we were able 

to demonstrate that these early, fast reactions would occur as fast as we could mix a propellant. (See Slide 8, Appendix H) There was basically zero activation energy that was driving the 

system. When you have zero activation energy reactions like that, you can separate the propellants, and that was giving us problems on the Corporal. When you put some force mixing into 
Gerard Elverum, TRW - Lunar Descent Engine

Apollo Rocket Propulsion Development 75

it, you can get the reactions to go in less than a millisecond. (See Slide 6, Appendix H) But, in 

building that thing, it seemed like a very good idea that as long as we could get those reactions 

to go, we’d put that concentric tube together and see if we could make an injector out of it. If 

we had these early reactions going very rapidly, we didn’t need distributed injection in order 

to be able to run the thing as a rocket engine. When we ran this with dinitrogen tetroxide 

(nitrogen tetroxide or nitrogen peroxide or N2

O4) and hydrazine, we were unable to pull the 

inner tube back away from the face. As soon as we pulled it back in about fifty thousandths 

of an inch, the whole end of the tube would just detonate. That showed how fast those reactions could be because that was going off even when just a small percentage of the material was 

mixed. But by pushing that forward a little bit, we were able to run this engine, run it at very 

high combustion efficiency, and run it stably. 

We were running it with N2

H4 -hydrazine that everybody else was having terrible problems 

with, because it was very difficult to handle that combination from a stability point of view. 

When I went down to Space Technology Laboratories and they said, “Let’s start a throttling 

rocket engine program,” I determined that a fixed area injector with a separate flow control 

valve was one way you could throttle an engine. (See Slide 4, Appendix H) You could have a 

variable area injector - that was another way that you could throttle it - or you could have a 

variable area injector with a separate flow control valve. If you were going to vary the area of the 

injector, you had to be able to do that in a reasonable fashion. If you had 1,000 holes, trying 

to vary the area of those was going to be very difficult. But, there are problems with throttling 

with a fixed area injector and a separate flow control valve. If you back off on the flow rate and 

let the pressure drop, you need some minimum delta pressure across your injector at the low 

end. Then, that drives the delta pressure up so fast that you can’t throttle 10:1. You just can’t 

get there. It required tank pressures of thousands of pounds per square inch (psi), or if you start 

at 300 psi, you would have to have like one psi across the injector at the low end and that isn’t 

going to work either.

The second way that you could talk about throttling an engine is with just a variable area 

injector. (See Slide 5, Appendix H) But, that has its own set of problems if you look at the tank 

pressure. Let’s say we set it at 450, and we wanted a 300 psi chamber pressure. If you are going 

to throttle it just 10:1, this would be the chamber pressure, you get an enormous pressure drop 

across the injector, and the injector has to be very, very tiny. In small thrust engines, which we 

were working on at the time, 500-pound thrust, the dimensions around that throttling injector 

would be so small that we had no possibility of building it and still be in uniform flow. That 

was a major drawback to that type of system. But, if we go to a system that combines both a 

valve and a throttling injector, then we have an entirely different kind of process going on in 

an engine. 
Chapter Six

76 Remembering the Giants

If you have a 300 psi chamber pressure, you are going to throttle it down to thirty psi at 10 

percent thrust. You would like to be able to set the delta pressure across the oxidizer and the 

fuel in the injector at every part of this throttling diagram to be optimum in order to get high 

performance and to become stable. So, you’d like to be able to do that independently. One 

way to do that is to look at the pressure drop across the valve; you can go to cavitating flow. If 

you drop the delta pressure below the inlet to the valve below a given number, it will go into 

cavitation and at that point, we can set the delta pressures across the oxidizer and fuel to be 

any value that we want for purposes of combustion. That will totally isolate the problem of 

controlling the flow rate from the problems of optimizing how the injector will work. So, the 

one fundamental principle that came out of this was – separate the functional things whenever 

you can, so that you are free to optimize the function in this case, controlling the flow rates 

and the mixture ratio to an exact number from the functions and the processes of optimizing 

the injector. Then, we had what is called a cavitating venturi valve, and we designed it to be a 

throttling cavitating venturi. Then, we had it linked 1:1 with a single element injector in the 

middle of the chamber, which was optimized as far as pressure drop across the oxidizer and 

fuel at every flow rate to maximize the performance in the chamber. By separating these basic 

functions, we were able to spend our time optimizing the injector for the chamber and the 

mixture ratio control and the flow rate. So, over this total complex duty cycle of the Apollo 

descent, we knew we would have absolute control of the mixture ratio, no matter what profile 

they decided to run.

In a nutshell, this became the basic concept of the Lunar Excursion Module (LEM) descent 

engine. The injector has a sliding sleeve that controls both the oxidizer in the middle and the 

fuel coming down the outside. That was linked 1:1 with two cavitating venturi valves, so that 

for every single flow rate that was going through those cavitating venturi valves, there was 

an optimized delta pressure and injector configuration, totally linked together. (See Slide 9, 

Appendix H)

In 1962, we heard that the Apollo Program was probably going to the descent stage concept, 

and Rocketdyne had been given the role of building the descent engine because they were 

the engine guys in the United States at the time. But, we also heard that NASA was getting 

concerned about the basic throttling engine design and were going to go to a backup program. 

They wanted somebody to do a backup. Well, I took this diagram to NASA and said, “We want 

to bid on the Apollo engine.” They said, “Who are you, and why are you in our office?” I came 

back with some data on this engine. It showed that we were using hydrazine and N2

O4, which 

nobody else had been able to run effectively. We showed them that we had throttled it over not 

10:1 but 20:1, that we’d kept a flat performance profile over the whole 20:1 throttling range, 

and that we never had any indication of instability or any other behavior. So, they gave us a 

Request for Proposal (RFP) and said, “Okay, we’ll let you bid on the backup.” 
Gerard Elverum, TRW - Lunar Descent Engine

Apollo Rocket Propulsion Development 77

We had to scale that up by a factor of ten because that turned out to be the biggest thing we could 

put in our test stand. We only had a small test stand in the middle of Inglewood, California, at 

that time. (See Slide 13, Appendix H) We scaled it up to 5,000 pounds from the 500-pound 

engine. We built an ablative lined chamber, so that it would have to be uncooled and ran that 

at 5,000 pounds of thrust. Everybody was saying, “Okay, it’s going to go unstable and that will 

put you out of business.” It didn’t go unstable. So, that gave NASA and the Grumman Aircraft 

Engineering Corporation, who was the prime contractor, a problem because now we were still 

in the middle of competing, and they weren’t sure they wanted a little outfit like us competing. 

But there we were with this engine running and giving good performance and throttling 10:1 

in our facility. (See Slide 13, Appendix H)

The competitors for the backup engine were Reaction Motors Inc., a liquid rocket engine 

company that had done the Vanguard engine, and Aerojet. One of the things they said was: 

“It’s okay to have a 5,000-pound engine, but the LEM is going to have to be 10,000 pounds.” 

That meant it was going to be this diameter, and if it was that diameter, it was going to go 

unstable and that’s it. So, they said, “You’ve got to demonstrate with a bomb, in the right 

diameter chamber and show that you are going to be stable.” Well, it took us a while to build 

that big diameter chamber, and it turned out that by the time we got it all set up on the test 

stand, NASA said, “We’re coming out.” Grumman said, “We’re coming out.” It was a Saturday 

morning, and we had never been able to fire that engine in order to see whether it was going 

to work. We used steam-blow in the stand in order to be able to run 10:1, and run full on the 

2:1 nozzle that we were using.

We built what we called the “iron pig.” (See Slides 14 and 15, Appendix H) It was seventeen 

inches in diameter. We ran the 5,000-pound injector into it and filled it up with bombs around 

in various different locations. I had to fire the first one of those tests in front of NASA and in 

front of Grumman, because we had just barely got it on the stand Saturday morning. Obviously, my management was not too happy with the idea of firing for the first time in front 

of the customer, but that was it; you either go or you didn’t go. Well, we fired it, and it was 

completely stable. At 100 psi, we fired a bomb, went up to 210 psi. Within about ten milliseconds, we had immediately stabilized the engine at that size. Well, as a result of this test, and as a 

result of the fact that we were throttling 10:1 with good performance, they gave us the backup 

contract. It took us about a year and one-half of very intense competition with Rocketdyne, 

but at the end of that time, NASA came back to us and said, “You’re now on for being the 

descent engine contractor.” Well, what we had to do was promise that we would put a whole 

facility together at San Juan Capistrano, and we built a whole series of test stands there. (See 

Slides 16 and 17, Appendix H)
Chapter Six

78 Remembering the Giants

For the duty cycle for the LEM, we started out at 10 percent each time. They wanted to start 

the duty cycle at 10 percent in order to get the vehicle stabilized. We started at 10 percent for 

about thirty seconds, and then kicked it up for about six or seven seconds. This was the duty 

cycle in order to go from the orbit they were in, into the orbit that was required to start the 

descent down to the surface of the moon. Then, we waited for about one hour and had the 

restart. We restarted again at 10 percent for about thirty seconds to stabilize all their stuff. 

Then, we went up to full thrust. This was the braking phase coming along on the descent. 

Then, we dropped down to what we called the “flare out.” When we dropped from the full 

thrust position, we went from a calibrated full-thrust engine into the cavitating venturi region 

of the valves. From that point on, we could do whatever we wanted in this duty cycle. We never 

knew exactly what they were going to do, particularly once they got into this phase. (See Slide 

18, Appendix H)

It turned out that the most critical mission we ever ran was Apollo 11. That was the most difficult landing. It exercised the engine to the absolute maximum in terms of the ablative liners 

capability, fuel control, and everything else. One of the things we had to do was to be able to 

set down on the moon and land on top of a rock. It turned out, like everything else, NASA 

came up with a specification for the rock this engine would have to land on and for the way 

the nozzle would have to crush on it. When I was sitting at Grumman during that first landing, 

the first thing that Neil Armstrong did when it came down and he found out they were over a 

boulder field, was to begin hunting around to figure out where he could find a NASA specification rock on which to land. That’s because we had demonstrated the nozzle would crush okay 

on a rock of the specified size. At least, that’s the way I felt back at Grumman. So, Armstrong 

kept looking and looking and looking, and I said, “Find the damn rock and set it down.” We 

were facing problems with propellant getting ready to run out. Armstrong finally set it down 

with less than a fifteen-second margin. If we had gone fifteen seconds more, he would have 

had to abort, but Armstrong was not about to abort that flight, so he went clear to the end. 

We went almost to the absolute maximum of everything we could do, and I was probably the 

only guy in the world that knew where all these margins were and how close we were getting to 

them. They talked about a few guys at Houston getting blue when they finally set down. I told 

them, “You’ve got one back here at Grumman who was blue a long time ago.”

The overall landing mission duty cycle involved 60 percent and 10 percent hovering modes. 

Talking about the descent engine, the thrust chamber design had a titanium case lined with 

ablative material out to a nozzle expansion ratio of 16:1. Then, we had a columbium nozzle 

extension from there out to 48:1. We used the Wah Chang Company to get the columbium 

nozzle and get the coating done. Those nozzles were very successful, but they had to be very 

Armstrong finally 

set it down with 

less than a fifteensecond margin. If 

we had gone fifteen 

seconds more, he 

would have 

had to abort
Gerard Elverum, TRW - Lunar Descent Engine

Apollo Rocket Propulsion Development 79

thin. That nozzle had to be like seven-thousandths of an inch at the exit cone, and that was 

out about five feet in diameter. So, we stabilized it with a single flat ring in order to keep its 

shape. But, the reason it had to be that thin was that, otherwise, it could have kicked the whole 

descent stage over if it had hit some of these boulders at whatever angle. It was very successful 

that we were able to make that columbium nozzle work. (See Slide 19, Appendix H)

If you look at a cutaway of the final descent engine injector, you can see how the central element 

injector, the oxidizer, came down the middle and deflected off of the columbium tip. The fuel 

came down the outside of the tube, and a single moving part allowed us to change both the 

area of the oxidizer and the area of the fuel with a single-sleeve, driven and mechanically linked 

up 1:1 with the two cavitating venturi valves. We used a barrier coolant in the engine because 

it was ablatively lined. This engine had to run for 1,000 seconds ablatively, which is a long, 

long time if you are using ablative materials. We barrier cooled this just enough to keep that 

1,000-second capability on the engine, and it cost us a little bit of performance. We lost about 

1.5 percent of overall performance by doing that, but it allowed us to run the engine for up to 

1,000 seconds. (See Slide 21, Appendix H)

When it came to the flow control valve, it was the first time I know of that anybody built 

a controllable cavitating venturi control valve. The reason I decided to go with a cavitating 

venturi on the engine was that I had used them at JPL, trying to run fluorine and hydrazine, 

and chlorine trifluoride and hydrazine. If you know anything about chlorine trifluoride and 

fluorine and hydrazine, you are never sure what’s going to happen in the chamber. We would 

put fixed, cavitating venturis in the line in order to be sure that we were getting the right 

propellant flow at the right mixture ratio. We then could study what was really happening in 

the chamber without having what was happening in the chamber change the mixture ratio 

and change the flow rates all over the place. When we needed a control valve, we said there’s 

no reason we can’t take a cavitating venturi and control the area of it, which would give us 

whatever throttle profile we needed over the entire throttle range. That was the valve that was 

put on the descent engine. On the head end, there were cavitating valves on each side. The 

actuator was an electrical actuator done by Bendix Commercial Vehicle Systems. (See Slide 22, 

Appendix H) 

The upper end of the valves was tied to a cross-arm, which was set on the pivot. The other 

end of the cross-arm goes to drive the sleeve on the center element of the injector. We went to 

quad ball valves on the fuel and oxidizer sides in order to assure that we had a positive chance 

for a shutdown or startup. The injector was spring loaded against the cross-arm, so that if we 

lost any electric power, the engine automatically would go to full-thrust position. That was 
Chapter Six

80 Remembering the Giants

important, because if we lost electric power on that actuator during the descent, the astronauts 

had to be able to go back up into orbit. The engine had to be able to go back up to full thrust 

where they could then turn around and go back up and, hopefully, get clear back into orbit, so 

they could tie up with the command module again. That was a spring-loaded system where the 

actuator had to drive it down to the low-thrust position. (See Slide 23, Appendix H)

The lightweight chamber liner evolved over a whole series of optimizations. We ended up with 

oriented silica fabric on the inside. These were oriented about sixty degrees to the flow. Then, 

we had a lightweight metal felt material as an insulator outside of that in the chamber, and that 

had a titanium structural shield around it. Then, we had a radiation exit heat shield because 

of the way this engine was buried down underneath the landing system where it was only the 

bottom part of the nozzle that really stuck out there. We had to insulate that. (See Slide 24, 

Appendix H)

This was a major development problem and a major lesson that we learned. When we had this 

thing almost finished and designed, the people who made the silica fabric in the United States 

decided they didn’t want to have anything more to do with it. Unbeknownst to us, they got 

their last material from some French company, and when we got that material and ran an ablative liner, everything came to pieces. The silica, even though it was supposed to be the same 

spec material as we had been developing, did not, in fact, perform the way the American silica 

fabric worked. It took us a long time to be able to sort that out and to get back to a supplier 

that would provide a material that was okay. The qualification specifications on something 

like those composite materials were extremely sensitive, and we had to be able to watch every 

single detail. A little thing like somebody deciding to get their fibers from someplace else could 

totally change the characteristic, and that happened to us on this engine development.

The “all-up” engine had a 48:1 exit cone. The injector was in the center, and the cavitating 

valves sat on the outside of it. We had a square gimbal and had to gimbal the engine about six 

degrees. We gimbaled at the throat, not up at the head end. The gimbal worked very well. We 

never really had any problem with our gimbal design. (See Slide 25, Appendix H)

We did combustion stability tests; thirty-one bomb tests were conducted with five- to fiftygrain charges. We induced spikes from up to 175 percent of chamber pressure in all those 

tests. We had to bomb it at 10 percent, 25 percent, 50 percent and 100 percent thrust. The 

recovery was always less than ten milliseconds. We would give it a spike, and it would damp 

out immediately. We would recover in every test. At low thrust, it would take longer to recover 

as the pressure was much lower. But, we never sustained a high-frequency instability. In all of 
Gerard Elverum, TRW - Lunar Descent Engine

Apollo Rocket Propulsion Development 81

the engines and all of the tests throughout development and flight, we had never had the case 

of combustion instability in any of the central pintle injectors. As far as I know, that hasn’t been 

the case for any other engine. In some engines, people put in baffles. They knew that we had to 

put them in on the Atlas engine when it went unstable on Atlas 29 and Atlas 30 in Florida and 

destroyed the launch stand. Those instabilities came out of the blue when people thought we 

had it understood. We went to baffles in the Atlas engine, but that’s a damping system; it is not 

a dynamically stable system. It prevents the instability from growing by damping it out, and if 

you can damp out the disturbance, then you have to wait until it might get a new disturbance. 

Then, it would damp that out, but you won’t racetrack it around and cut the thrust chamber 

off the injector. In all of these other types of engines, they either go to baffles or they go to 

acoustic chambers, which is another method of damping. But in the LEM engine, we have 

never been able to sustain an instability in that engine. (See Slides 26 and 27, Appendix H)

In my opinion, the reason is that this was what we sold to NASA and Grumman after we 

did the iron pig test. Why was it stable? Well, it was basically stable to the first tangential 

mode, because it is like trying to play a violin while your bow is bowing at the nodal point of 

the string. You can’t make music that way. If you put the energy in at the nodal point of the 

tangential mode, you can’t support a dynamic, unstable combustion. So, we had the central 

injection system operating in a stable mode, and the basic flame front was where most of the 

energy was being converted. If we bombed that engine, we drove that energy release zone in 

towards the nodal point and as soon as we drove it into the nodal point, it could not support 

that frequency anymore. It could not couple and drive the instability. We’ve done this at every 

size engine diameter you can imagine, up to three and one-half feet. In fact, there was one at 

NASA’s John C. Stennis Space Center that was a 650,000-pound thrust engine operating with 

hydrogen and oxygen, which everybody said must go unstable. We did it at Edwards Air Force 

Base in California with a 250,000-pound thrust engine, where again we set off bombs, and it 

was stable. In every known case, both for the tangential and the radial, this injector configuration goes into dynamic stability. (See Slide 28, Appendix H)

Through March 1967, we had conducted more than 1,700 injector tests for 70,000 seconds of 

operation. We had twenty-six head-end assemblies overall with 55,000 seconds of operation. 

We did high altitude tests – twenty-seven builds, 195 different starts, and 18,000 seconds. 

We had a lot of experience on this hardware. But, the key to the Apollo Program success, one 

of the keys that I believe, is it was not a hardware-poor program. It was not, “let’s just build 

one engine, and somehow, we’ll solve all the problems as we go.” It was a program that put 

in enough hardware that we could go to the corners of the box and find out where the failure 

modes were. What did it take? What were the design criteria that would, in fact, provide you 

But, the key to the 

Apollo Program 

success, one of the 

keys that I believe, 

is it was not a 

hardware-poor 

program.
Chapter Six

82 Remembering the Giants

with margin against the critical failure modes? You had to have enough hardware to demonstrate that was true. I think that was true of the F-1 engine program and on the service module 

program. We had hardware. We weren’t nursing a few little pieces of hardware on some test 

stand some place and trying to figure out whether or not we had really uncovered all of the 

failure modes and demonstrated the technical design criteria necessary to assure that those 

modes were adequate. (See Slide 29, Appendix H)

This program had a lot of hardware in it even though it went very, very fast. The program was 

an extremely fast program. The Rocketdyne engine program started in February 1963. We 

were told that we could start our backup program in July 1963. We had to bring on a whole 

Capistrano1

 test facility, build all of those sites. We did our first, full throttling range test down 

at that site in November 1964. That was after bringing in the steam systems, and the ejectors 

and everything else. We finished Phase A qualification in 1966, completed Phase B qualification in 1967, and delivered our nine engines to the customer by May 1967. (See Slide 30, 

Appendix H)

It was a very fast program. We had a lot of hardware in it. But, in fact, they gave us the resources 

that were necessary in order to carry out a program as we saw we had to do. One reason for that 

was we had a guy named Joe Shea, who was the NASA chief engineer of the Apollo Program 

during most of the early critical development. Shea was one of the best system engineers I’ve 

ever come across. He would make the decisions that said, “We’re not going to stick with some 

specification that we arbitrarily set up at the beginning of the program. I’m going to balance 

my specifications and my requirements, and so, if I have a limit on this component, I’m going 

to find out if I can change that around and take the load over here someplace else.” And he did 

that in real time many, many times, which is one of the reasons I think we were able to get a 

successful program in a short amount of time, as we did with Apollo. 

For any new program, one of the things you really must have is a guy who is the system engineer and is in charge of the system engineering. You have to build that discipline in such a way 

that you have a complete system engineering process, and you know how requirement allocation decisions are being made. The same thing is true for the Air Force now. They are relearning 

that same thing – you must have a system engineering process and a leader with authority. In 

1

 San Juan Capistrano, near San Diego, is where most of TRW’s propulsion test areas were located. Today, Northrop 

Grumman Corporation (NGC) has begun dismantling facility assets at Capistrano Test Site (CTS) after taking over 

the former TRW Corp. in a merger.
Gerard Elverum, TRW - Lunar Descent Engine

Apollo Rocket Propulsion Development 83

the end, that is what controls these kinds of things. Otherwise, you back yourself into a corner 

with a whole set of arbitrary things, and you don’t have the money to get out from that corner. 

I would argue that one of the first things you must think about is the structure of your system 

engineering process and the authority of your chief engineer. Everything else falls out from 

that. In our case, Joe Shea was that guy. He did the job very well. I think he got tagged with a 

bum rap that was not necessarily his fault at the end of things, but during this critical part of 

the development, he was the guy that let it go ahead, let it go forward.

We got kind of on the wrong side with some of the guys at Houston. The astronauts didn’t 

mind what we did, but NASA wasn’t too happy. We put this billboard up on NASA Road One 

going into Johnson Space Center, where every time people would go in, they’d see this sign 

saying, “The last ten miles are on us.” That’s because the astronauts went into orbit for lunar 

descent at 50,000 feet or about 10 miles. That’s where we started our descent firing. (See Slide 

32, Appendix H)

Gene Cernan was the guy assigned in the astronaut corps to follow the descent engine development. He was out at our facility many times. He knew everything about the engine. He was 

commander of Apollo 17, the final Apollo mission. He was the last man to step off the surface 

of the moon. When he got back to Houston, I sent him a letter with a little drawing I had 

done for him. It showed two astronauts hanging onto to a descent engine as it travels to the 

surface of the moon. One of the astronauts is saying, “When TRW said, ‘The last ten miles are 

on us,’ I didn’t realize that they were going to ride down on the engine instead of the lander.” 

(See Slide 33, Appendix H) 

---------------------------------------

Editor’s Note: The following information reflects a question-and-answer session held after Elverum’s 

presentation. 

QUESTION: In looking at the statistics chart showing the number of starts, there was a ratio 

of about 20 percent sea level tests and 80 percent altitude tests in the descent engine program. 

Maybe that ratio was comparable for the other engine, the Aerojet engine. But, given that, let’s 

say that thrust is an important parameter that you had to measure, especially with the throttling 10:1. We often say, “Test like you fly, and fly like you test.” In terms of this engine, how 

did you know that the facility thrust measurement and the facility altitude simulation were 

sufficient to replicate or represent flight conditions?
Chapter Six

84 Remembering the Giants

ELVERUM: As we went through the program, what we determined, and what we all agreed 

on, was that the thrust coefficient (Cf

) of the nozzle, after you get past a certain point, is really 

an engineering parameter. It’s not a fundamental parameter that is going to be highly variable. 

Once we knew what the contour of the nozzle was, and once we knew what its characteristic 

was out to 2:1, we could calculate what the 48:1 thrust coefficient was going to be. In every 

case that we made a test, the calculation was precise. We weren’t looking for a problem out at 

48:1. Once we crushed the nozzle and said, “Yeah, we can land on the boulder,” and once we 

had the thermal profile of that columbium nozzle, we did not require a lot of effort there. The 

real characterization was done in throttling over the 10:1 with the injector and controlling the 

mixture ratio on that – the whole head-end assembly – out to 2:1. I think everybody at NASA 

and Grumman agreed that flying like you test is great, particularly if you are using an aircraft 

engine. But, in this case, the thrust coefficient of the nozzle was not an issue.

QUESTION: How many engines did you test in the altitude chamber?

ELVERUM: I can’t remember. We had twenty-seven complete 48:1 engine builds that were 

acceptance tested and tested in various types of throttling programs, which is a lot of hardware. 

I’ve been following the SpaceX program, for example, as part of the role that I do for the U.S. 

Department of Defense. There, they just have a few engines. And, a whole bunch of other 

commercial rocket companies who want to get in the business talk about, “Well, if I could get 

a little more money, I would be able to build two engines.” They think that’s going to demonstrate that we ought to put a 500 million-dollar payload on top of a space vehicle that’s had 

two engines tested and only tested to nominal conditions. They never had a chance to burn out 

three of the engines and go way off nominal. 

It takes a lot of hardware to be able to demonstrate that you have understood the criteria that 

established the risk of an engine. The risks of an engine means you have to define the failure 

modes and you have to do some real “honest-to-God” risk assessment. I mean quantitative risk 

assessment, not just qualitative. By quantitative, I mean you assess what the profile of the risk 

is, then you test that profile and see if you can collapse the uncertainty of that risk. That takes a 

lot of hardware. It’s something that has to be put into a new program. We did all kinds of stuff 

like that, but it is very expensive, very expensive.

QUESTION: Do you have any comments on how the descent engine was used in Apollo 13?

ELVERUM: In Apollo 13, it was very strange. I’ve always had a kind of an eerie feeling about 

the whole program because when I was at JPL as a young engineer working with hypergolic 

propellants and cavitating venturis, I had no feeling or concept of an Apollo Program. Apollo 
Gerard Elverum, TRW - Lunar Descent Engine

Apollo Rocket Propulsion Development 85

5 was the first time that we ran the descent engine in space, and it was hooked up in space after 

separating from the S-II.

We had the tandem configuration of the service module, the command module, and the LEM 

sitting out there, and we were to fire the LEM. On Apollo 5, we were firing the LEM to show 

how it would work. There was a problem. I can’t remember where the problem was, but something caused a problem before that engine had finished its burn. It was not in the engine, but 

there was some other problem, and NASA made a controlled shutdown. Then, they came to 

us and asked, “Hey, we’re up there. We want to finish this test program. Is it okay if we restart 

that engine again in space with this tandem configuration?” We said, “As long as it has been 

more than forty minutes since you shut down, our analysis says that you will be okay in terms 

of the thermal characteristics of the inside of that chamber.” They restarted it and pushed that 

system around in orbit on Apollo 5.

It turned out, that when it came to Apollo 13, we went back into the record, and said, “Hey, we 

have pushed this system around up there on Apollo 5, and we have also restarted this tandem 

configuration.” The requirements on Apollo 13 were to put it back into play. The spacecraft 

was out of free return to the earth at the time of the accident. It would not have come back. 

NASA said, “Okay, we’ll use the descent engine to put the spacecraft in a free trajectory; it will 

go around the moon and be on free trajectory back to Earth.” Then, as it came around the far 

side of the moon, the guys found out that they had an oxygen problem. As you remember, 

things were getting pretty bad in there. They said, “We’ve got to get it back as fast as we can. 

Is it okay if we re-fire the engine? Now, we’re in a free trajectory, so we want to put as much 

delta-v (or change in velocity) in as we can. Can we re-fire right now?” We said, “Yes, the data 

says it has been this period of time.” We could re-fire the engine, run the rest of the duty cycle 

up as far as we needed while preserving enough fluids to make the final correction as the spacecraft got near Earth, and restart the engine. It was pretty fortuitous that we could give them 

those answers.
86 Remembering the Giants

Similar in shape but not size were the three big engines aboard Apollo spacecraft. Two of them had no backup, so they were designed to be the 

most reliable engines ever built. lf the service-propulsion engine failed in lunar orbit, three astronauts would be unable to return home; if the 

ascent engine failed on the Moon, it would leave two explorers stranded. (A descent-engine failure would not be as critical, because the ascent 

engine might be used to save the crew members.) (Source: http://history.nasa.gov/SP-350 - Apollo Expeditions to the Moon”, published 1975)
Apollo Rocket Propulsion Development 87

Chapter Seven

Tim Harmon

Rocketdyne – Lunar Ascent Engine1

1

 Please see Section 4 for Mr. Harmon’s biographical information.

The ascent engine was the last one from the moon, and I want to focus on the idea of redundancy 

and teams in regard to the engine. By teams, I mean teamwork – not just within Rocketdyne. 

It was teamwork within Rocketdyne; it was teamwork within Grumman; it was teamwork 

within NASA. These were all important elements leading to the successful development of 

the lunar excursion module (LEM) engine. Communication, rapid response, and cooperation 

were all important. Another aspect that went into the development of the ascent engine was 

the integration of technology and of lessons learned. We pushed all the above, plus technology 

and lessons learned, into a program, and that led to a successful result. One of the things that I 

like to think about – again in retrospect – is how it is very “in” now to have integrated product 

and process teams. These are buzzwords for teamwork in all program phases. That’s where you 

combine a lot of groups into a single organization to get a job done. The ascent engine program 

epitomized that kind of integration and focus, and because this was the mid- to late-1960s; this 

was new ground for Rocketdyne, Grumman, and NASA.

Redundancy was really a major hallmark of the Apollo Program. Everything was redundant. 

Once you got the rocket going, you could even lose one of the big F-1 engines, and it would 

still make it to orbit. And once the first stage separated from the rest of the vehicle, the second 

stage could do without an engine and still make a mission. This redundancy was demonstrated 

when an early Apollo launch shut down a J-2 second-stage engine. Actually, they shut down 

two J-2 engines on that flight. Even the third stage, with its single J-2 engine, was backed up 

because the first two stages could toss it into a recoverable orbit. If the third stage didn’t work, 

you were circling the earth, and you had time to recover the command module and crew. 

Remember how on the Apollo 13 flight, there was sufficient system redundancy even when we 

lost the service module. That was a magnificent effort. TRW Inc. really ought to be proud of 

their engine for that. (See Slide 2, Appendix I)

We had planned for redundancy; we had landed on the moon. However, weight restrictions 

in the architecture said, “You can’t have redundancy for ascent from the moon. You’ve got one 

engine. It’s got to work. There is no second chance. If that ascent engine doesn’t work, you’re 

stuck there.” It would not have looked good for NASA. It wouldn’t have looked good for the 

country. There was a letter written that President Richard Nixon would read if the astronauts 

got stuck on the moon, expressing how sorry we were and so forth. It was a scary letter, really. 

The ascent engine was an engine that had to work. (See Slide 3, Appendix I)
Chapter Seven

88 Remembering the Giants

The lunar excursion module vehicle was an ungainly, lightweight vehicle. The vehicle walls were 

really thin pieces of metal. The ascent engine was buried inside. All you could see was a little 

bit of the nozzle extending from the bottom of the vehicle. (See Slide 3, Appendix I) The LEM 

vehicle ascent stage height was twelve feet, and the diameter was thirty-one feet. The LEM 

weighed about 11,000 pounds. Now, the ascent engine was fifty-one inches high; it wasn’t a 

very big thing. It produced 3,500 pounds of force (lbf) thrust. It was a very “cozy” environment, 

this LEM vehicle. By cozy, I mean it had to be lightweight. By cozy, I mean they couldn’t even 

install seats due to the weight penalty. The astronauts stood! They had parachute-like harnesses 

to provide stability. The astronauts stood and looked out “tilted” windows to see the landing 

area. The descent engine had a backup, the ascent engine, and the reaction control engines 

on the vehicle (there were four sets of four) were redundant. Again, the situation was: we did 

not have a backup for the ascent engine. When you don’t have a backup, you make the thing 

simple. NASA was smart in this. They said, “All we want you to do is one on/off switch.... You 

don’t have to throttle – constant thrust is okay. We are going to fix the mount and we are not 

going to gimbal this engine. We are going to have redundant valves and hypergolic ignition. 

Get the two propellants together, it lights.” This is the plan for a simple, reliable engine. That 

was the approach, and I think it was a sensible idea. (See Slide 4, Appendix I)

The specifications for the engine were: pressure-fed, 3,500 pounds of force vacuum, a 310 specific 

impulse (Isp), and hypergolic propellants. One possible hypergolic fuel, hydrazine, freezes like 

water at two degrees Celsius. This was unacceptable, so unsymmetrical dimethylhydrazine – I 

don’t know who comes up with these names – suppressed the freezing temperature. Reliability 

was the name of the game; keep it simple. That was the architecture and the plan. (See Slide 

5, Appendix I)

It’s important to remember Rocketdyne did not start out with this contract. This was the 

Bell Aerospace contract, which began in July 1963. Bell had challenges. Every development 

program does. Their biggest issues were thrust chamber erosion and combustion instability. 

These continued in the program all the way up to 1967. If you followed the descent engine 

program, you knew it was through qualification testing (also known as “qual”) in 1967. In 

1967, Bell hadn’t even started qualification testing. So, suddenly NASA realized they could get 

to the moon, but they might not be able to get off the moon – not a good situation. NASA 

awarded a backup program to Rocketdyne in August 1967. We were successful with the main 

issues. By 1968, we had won the complete contract and were off and running. A calendar was 

established to set the major development schedule milestones. (See Slide 7, Appendix I) The 

Bell program was given to NASA in July 1963. The Agena engine was Bell’s basis for the ascent 

engine. It was a reasonable upper stage engine and had several successful satellite launches. I 

think it was a 16,000-pound thrust engine. In late 1964, I believe, NASA said, “We want to 

do stability testing. We want the engine duration to be longer.” So, the specifications changed 

Reliability was the 

name of the game; 

keep it simple. 

That was the 

architecture and 

the plan. 
Tim Harmon, Rocketdyne – Lunar Ascent Engine

Apollo Rocket Propulsion Development 89

a little bit, and Bell began stability testing. The original injector was unbaffled. As soon as 

they bombed it to test stability, they had instability. They put a baffled injector in to improve 

stability, but were not successful in eliminating the instability problem. That created a situation 

in which Rocketdyne became a backup program. (See Slides 6, 7 and 8, Appendix I)

Because I did not work for Bell Aerospace, I can only provide my impressions of the ascent engine 

development problems based on NASA’s historical documents. One of Bell’s development 

issues was their program had limited test hardware early in the development cycle. When 

Bell had instabilities, there were manufacturing issues that could be blamed. You also get the 

impression there was some denial that there was a problem. When engine combustion went 

rough, they were never sure of its cause. “Was it manufacturing; did we do something wrong 

there? If we had an ablative wall thrust chamber, instead of steel battleship-type thrust chamber, 

maybe the ablative would change the acoustics, and it might absorb it.” Instabilities tend to 

do a lot of hardware damage. I think their rough combustion cutoff was a little slower than 

Rocketdyne’s, or what would be typically used, so they sustained a fair amount of hardware 

damage by the time they shut the engine down. It didn’t quite tell them what the issues were. 

All those variables added up to mask the root cause of their problem. 

Rocketdyne had heads-up information the ascent stage was in trouble. There’s a story I can’t 

deny or confirm: that Rocketdyne marketing people at NASA’s Johnson Space Center in 

Texas had the ability to read documents upside down on a person’s desk. Doing so, they 

realized there were problems with the Bell ascent engine. It did not take a rocket scientist to 

realize that problem, because in 1967, Bell hadn’t started qualification testing. Rocketdyne 

thought, “There’s something wrong here. Maybe this is an opportunity.” It did prove to be an 

opportunity. NASA recognized the issue and said, “We need a backup ascent engine system, 

just in case.” They put out a Request for Proposal in June 1967 for a backup injector system 

that would address the stability issue. Three companies were looked at and down-selected. 

TRW might have been one; but I’m not positive about that. The others were Aerojet and 

Rocketdyne. Rocketdyne’s approach was to put everything we could think of into stability. We 

decided we had better design in baffles. We incorporated the best manufacturing technology 

that was available. We went with electron beam (EB) welding. We incorporated electron 

discharge machining (EDM) drilling, which is a real nice, advanced, repeatable system for 

drilling orifices. Then, we put in acoustic cavities.

Acoustic cavities were another stability aid. Rocketdyne had, in 1965 or 1966, an independent 

research and development program on a small rocket engine. It used a beryllium thrust chamber 

with a stainless steel injector. Beryllium is a brittle material. If the injector gets warm, it will 

expand, and the brittle beryllium material will break. To make up for that, engineers put a gap 

between the injector and the thrust chamber. We tested this in a pulse mode multiple times. 
Chapter Seven

90 Remembering the Giants

This was a small engine, just a little pulse mode engine. It turned out this little beryllium engine 

with a stainless steel injector occasionally had low performance pulses; i.e., the pulses wouldn’t 

go quite as high on the Datelink Independent Gateway Retrofit (DIGR). Mostly, pulses were 

to be a certain height. Occasionally, 3 percent of the time it turns out, pulses wouldn’t go quite 

as high. We didn’t understand that. We didn’t know what was going on there. There was no 

high-frequency instrumentation. It was just that we noticed this data anomaly. The hardware 

didn’t break, and there wasn’t a problem with the hardware. We thought, “It is working and it 

isn’t broke; let’s let it go.” But management said, “We have to understand the problem.” During 

the analysis, we thought, “Well, maybe it’s that gap between the injector and the beryllium 

chambers, and propellants are sneaking up there and detonating and causing some sort of 

instability.” The question then became: “Make the gap smaller or make it larger?” We chose to 

make the gap larger, and cut the diameter of the injector down. The instability disappeared. 

They were all now the same height. Engineering said, “Aha! We’ve got the problem solved.” 

Management said, “We don’t think you do because you don’t know what you did.” We had to 

analyze it. 

We then decided maybe it was a Helmholtz-type resonator disrupting the instability. Most of 

our engines in 1965 and 1966 were in production. We were not going to add acoustic cavities, 

which was a precursor of this type of stability aid. That idea went on the shelf. Then, along 

came this opportunity. We had come up with this injector and put a baffle in it. The president 

of the company, Sam Hoffman, said, “Let’s put those acoustic cavities in there.” I’m not sure 

we had all the analytic background, but that seemed to work on this little internal research 

and development (IR&D) engine. Hoffman said, “Put it in there.” Of course, we already had 

our design, and we had to do some redesign to get the acoustics cavities in there, but it was 

a brilliant idea, and it helped sell the program to NASA as the backup program for the Bell 

problem.

People talk about the competition at that time. It’s true we were in a space race, and we 

were doing pretty well maybe by 1966 or 1967. But NASA said, “Rocketdyne, you are still 

in competition with Bell.” To set the stage, Bell was conducting their development program 

in parallel after we won the backup contract, which put us in competition. We had the 

competition of the space race and the competition with the schedule. Near the end of 1967, 

we hadn’t even started qualification testing on the ascent stage. Rocketdyne had to respond to 

this competition. They recognized right off that this program had to be an integrated effort, 

and put together a team consisting of engineering, development engineering, design, and other 

personnel. It included inspection and quality teams. It included NASA and Grumman. It 

included a member of the test stand team. The test organization was represented there. They 

were all in one place, on one team.
Tim Harmon, Rocketdyne – Lunar Ascent Engine

Apollo Rocket Propulsion Development 91

Every morning at eight o’clock, we had a standup meeting. It was truly a standup meeting. The 

offices in those days were fairly small. At eight o’clock, people piled in there. Every morning, it 

was like trying to see how many people you could stuff in a Volkswagen. For fifteen minutes, 

they discussed the results of yesterday, what we would do today, and marched on. It was 

a twenty-four/seven type of operation. Rocketdyne really went on this twenty-four/seven 

operation with a vengeance. Because of schedule constraints and schedule pressure, the team 

pushed down the point at which decisions were made. This was 1967; I had only been with the 

company four years, a fairly junior engineer. I was in charge of stability testing, which was run 

in two shifts. The first shift and second shift were stability testing. The third shift cleaned up 

the mess we made in the first and second shifts; then, it started all over the next day. I was the 

development engineer, and my job was to get stability tests done as rapidly as possible. They 

said to me, “If we can gain schedule, you can work the crew weekends – Saturday and Sunday.” 

That part of my charter was a lot of responsibility for a guy of four years’ experience because 

the test and operation staff was thirty people a shift. I was a junior engineer able to say, “We’re 

going to spend some money this weekend to gain schedule.” We were really pushing. We had 

the long pole in the tent for 1967. I don’t mean that I was particularly special; that was the way 

it was for everybody on the team.

This was one of the most fun and fulfilling programs I ever worked on; it was a very intense 

program, but let me describe what Rocketdyne was able to accomplish. The contract proposal 

was submitted in June 1967. It was awarded as a backup in July 1967. We got a go-ahead in 

August. However, once we submitted the proposal in July, we recognized the need for speed. 

We started assuming we would win the contract, and we started building hardware. It’s a 

phenomenal thought, but we were able to design and get ready to go in two months.

We conducted our first altitude test, then our first full duration test two months later. We 

completed our preliminary design review in November. We completed our critical design 

review in December, the first engine system full duration altitude test in January. The term 

“design feasibility tests” meant we knew enough of what we were doing regarding stability, to 

finish proving that in March. Our development program milestones give some sense of what 

we were able to accomplish, and we were not hardware poor. (See Slide 8, Appendix I) We were 

able to get a lot done in a very short period of time. (See Slide 9, Appendix I)

For the design feasibility program, we had twelve injectors. We accomplished 872 tests, 

including 302 combustion stability tests. The combustion stability test used a little bomb. It 

was more like a blasting cap or a firecracker, but it was a little bomb. We built a battleship, just 

a plain steel thrust chamber. We could run this chamber about ten seconds before it got too 

hot. I could get five bomb tests off during a ten-second run. With every hot-fire, we could get 

It was a twentyfour/seven type of 

operation.
Chapter Seven

92 Remembering the Giants

five stability tests accomplished. NASA and Grumman were coming in to review the program 

one Monday. We were going to work through the weekend to get as many bomb tests off as we 

could. We wanted to have 100 bomb tests completed. We got ninety-nine tests done, and then, 

it went unstable. I couldn’t believe it. I was crushed. It turned out the instability was caused 

by the bomb shrapnel damaging the injector. In post-test, we disassembled the injector and 

recognized we had damaged the injector with shrapnel from the multiple bomb detonations. 

The wiring to the blasting cap was going backwards, hitting the injector face, changing the 

injector orifices’ shapes, and denting the injector. After that, we would take the injector apart 

after each test, and if we saw any damage to any orifice we would hand drill it back out. That 

eliminated the instability problem. But, it made for a very interesting NASA meeting that 

Monday because the injector went unstable Saturday night about eight o’clock, and we only 

had Sunday to figure out what caused this instability and whether we understood it. (See Slides 

10 and 11, Appendix I)

To give an idea of the twenty-four/seven intensity of the schedule, Rocketdyne decided to 

appoint only one development engineer for stability, and it was me. I don’t know why. But, 

to accomplish testing like that, I was to be available whenever the engine ran. In those days, 

I wore contact lenses and really couldn’t wear contact lenses much more than ten hours. I 

just couldn’t see beyond that. I had to do something to be able to see twenty-four hours with 

contact lenses. So, I would wear one contact lens at a time. I would look this way, or I would 

look that way. By three or four o’clock in the morning I knew it was time to change because my 

eyes were really getting tired. I was so tired, I managed to stick both contact lenses in one eye. 

It took a while to straighten myself out, but I did get the contacts sorted out and completed the 

test program. That shows you some of the dedication of the team. We really worked at it.

Another program issue was combustion chamber erosion. There were a couple of others as well. 

One was baffle weld cracks. Last, we had to make the engine lighter. Within nine months of 

becoming a parallel effort with Bell, Rocketdyne had gotten to the point of showing NASA 

that we knew what we were doing, and we became the single contractor for the ascent engine. 

Within nine months, we had caught up with Bell, who was probably four years of testing 

ahead of us, and we became the prime contractor for the ascent engine. That was quite an 

accomplishment.

Finally, with combustion instability, we did one test where we verified that the acoustic cavities 

were doing their job. On one test, we blocked the acoustic cavities, and sure enough, on the 

first bomb the injector went unstable. The acoustic cavities were validated with that test.

The chamber erosion issue was very difficult to solve due to the need of getting uniform flow 

from the injector. One way to protect ablative material is to put film cooling material down 
Tim Harmon, Rocketdyne – Lunar Ascent Engine

Apollo Rocket Propulsion Development 93

the walls of the thrust chamber. But the engine loses performance when you do that. There 

is always a trade off. We wanted to protect the thrust chamber, but we couldn’t do too much 

because the specific impulse performance goes down. After we had several tests that showed 

erosion, we discovered it was always occurring in the same general area. With the conventional 

wisdom of putting film cooling all around, we would end up with unacceptable performance. 

My manager, Mike Yost, said, “Well, you know, we don’t need to put film cooling all the way 

around, we just need to put it where it is eroding.” That really sounds like an easy solution now, 

but it wasn’t conventional at the time. In those days, you put film cooling in a nice symmetrical 

pattern. But Yost said, “Let’s just do it where we need it.” That eliminated most of the film 

cooling performance loss. This was a good example of thinking out of the box.

I mentioned that we had hard starts. That happens when the fuel gets to the combustion 

chamber before the oxidizer. When the oxidizer hits, it goes “bang,” which is not a good 

situation. Our solution was to make the fuel pipe large and add an extra little volume at the 

end of the fuel pipe, so that the fuel had to prime a large duct. When we did that, the oxidizer 

reached the combustion chamber first, before the fuel. The design looks pretty unconventional 

and is not real elegant. (See Slide 12, Appendix I) 

Testing was key to demonstrating high engine reliability. We had test facilities everywhere. 

There were two test sites at Santa Susana in California. We used White Sands, New Mexico; 

Reno, Nevada; and even Bell Aircraft test facilities. We had to take our engine to Bell and run 

it there. Bell was, of course, extremely interested in our design. We were still in a competitive 

situation. After they installed it, they thought it would be a good idea to X-ray the engine. We 

said, “No way.” They did focus on our unconventional inlet duct, thinking it was a stability 

aid. It was only a priming aid to keep the hard starts from occurring. If they thought it was the 

solution to stability, it wasn’t – perhaps it led them down an incorrect thinking path. It was an 

accidental thing, not done on purpose on our part. But it did cause Bell a lot of consternation. 

We ran successfully in their facility, partly to verify that our performance was equal to theirs. 

It was.

The LEM engine was on the critical path. It really was a twenty-four/seven operation. We 

had this team that included everybody. It really was an integrated effort. Another milestone or 

key part of the effort was communication. It was rapid. It was effective. Decisions were made in 

timely order for anything. Decisions were down to the lowest person. It was interesting that 

we had this close working relationship with Grumman Aircraft, as they had chosen Bell, but 

they recognized the issue and so did Bell Aircraft. I should also say that Bell was very helpful 

in the program. We had the injector. We assembled the engine. But the valves and the thrust 

chamber were Bell Aircraft’s. Everybody worked in concert, because the goal was to get to the 

moon and to get back. (See Slide 13, Appendix I)

Testing 

was key to 

demonstrating 

high engine 

reliability.
Chapter Seven

94 Remembering the Giants

When trying to understand how fast we were able to respond, it’s important to remember 

we were not hardware poor. We essentially got going within a month of the contract award. 

Our feasibility testing was completed. Then, the solution was demonstrated literally in eight 

months. This was a prototype integrated product and process development (IPPD) team. It 

was a totally committed, very goal-focused team, in which all the disciplines were involved. I 

think Rocketdyne really showed what they could do in a critical situation. I think we did put the 

lessons learned in this engine. We implemented them in an effective manner, and we integrated 

technologies within that environment. Once we got the engine working, NASA asked us to 

reduce its weight. This engine only weighed 171 pounds, but still, the weight had to be reduced. 

The reason was that NASA wanted to bring home more rocks from the moon. If the engine 

weighed less, they could do that. I was put in charge, and we took a 171-pound engine and 

reduced its weight by thirty pounds. That meant an additional thirty pounds of moon rocks 

could come back to Earth. That was a significant amount of improvement. That was a 15 

percent weight reduction on a pretty well-developed engine. (See Slide 14, Appendix I)

---------------------------------------

Editor’s Note: The following information reflects a question-and-answer session held after Harmon’s 

presentation. 

QUESTION: One of the things that we run into with the “can-do” attitude that you showed 

on your last slide is you don’t want to sacrifice quality, and that’s always the issue. How do you 

balance that? How did you balance that back then to make schedule and still maintain quality? 

What were you doing? What advice would you give?

HARMON: One way was that quality was part of the team. When you put manufacturing, 

quality, and engineering personnel together, quality can look at your design work and say, 

“Man, that is really hard to see whether that is going to be a good weld or not.” They are right 

in tune. I was talking recently with some gentlemen who had testing personnel as part of their 

team. They were there every day. They were there in the standup meeting. When an engineer 

said, “I need 100 strain gages,” the test guy said, “I got thirty. You are going to have to make do 

with thirty, or it is going to cost you ‘X’ bucks to go to sixty and it’s going to take ‘X’ months to 

get it available for you.” We had this immediate feedback of what we were asking for, whether 

it was doable, whether quality staffers could do it, and that was the virtue of that team. The 

quality check certainly was built into the design as best we could do.

. . . we took 

a 171-pound 

engine and 

reduced its 

weight by 

thirty pounds. 
Tim Harmon, Rocketdyne – Lunar Ascent Engine

Apollo Rocket Propulsion Development 95

There was one instance where we found weld cracks on this system. There is an art to welding. 

This piece was machine-welded, but there’s still an art to it in the setup. What was the problem? 

Well, it turned out, we were working two shifts, and the machinist on the night shift didn’t 

quite make the setup right. It was a difficult weld. It was difficult enough that we would 

sometimes overstress that weld if the setup was not perfect. It was a weld penetration issue. We 

solved the weld problem by removing the second shift welder from the task. 

I got to do the inspections on that weld, to see whether I felt the penetration was correct. We 

actually had one of our LEM engines in a “bird” (referring to an Apollo vehicle on a Kennedy 

Space Center launch pad) at Florida. Since it was not a big engine, it was hard to look from the 

outside and see how good the weld penetration was. It was supposed to be a one-day trip. 

The welding, though, was interesting because of how we were able to solve that issue and figure 

it out quickly and verify the welds on the LEM were fine – the same way I verified the quality 

of that paper swimsuit.

Rocketdyne met the challenges. The Lunar Excursion Module Ascent Engine was a prototype 

IPPD program, but it was probably the best yet for integrating lessons learned and technology. 

It was one of the best programs I ever worked on. I think almost every team member who 

worked on the LEM looks back on it, and is very pleased with the results.

Apollo Rocket Propulsion Development 97

Appendix A

Event Program
Appendix A

98 Remembering the Giants
 Speaker’s Biographies and Events Program

 Apollo Rocket Propulsion Development 99

 Apollo Rocket Propulsion Development 101

Pictured (left to right): Steve Fisher, Clay Boyce, Bob Biggs, Gerald Pfeifer, Tim Harmon, Gerard Elverum, 

Paul Coffman, and Shamim Rahman.

An actual F-1 Engine is shown in the background; a display model in front of the NASA SSC onsite visitor 

center, April 2006.

Appendix B

Speakers’ Group Photograph

Apollo Rocket Propulsion Development 103

Appendix C

Robert Biggs’ Presentation Viewgraphs
Appendix C

104 Remembering the Giants
Robert Biggs’ Presentation Viewgraphs

 Apollo Rocket Propulsion Development 105
Appendix C

106 Remembering the Giants
Robert Biggs’ Presentation Viewgraphs

 Apollo Rocket Propulsion Development 107
Appendix C

108 Remembering the Giants
Robert Biggs’ Presentation Viewgraphs

 Apollo Rocket Propulsion Development 109
Appendix C

110 Remembering the Giants
Robert Biggs’ Presentation Viewgraphs

 Apollo Rocket Propulsion Development 111

Apollo Rocket Propulsion Development 113

Appendix D

Paul Coffman’s Presentation Viewgraphs
Appendix D

114 Remembering the Giants
Paul Coffman’s Presentation Viewgraphs

Apollo Rocket Propulsion Development 115
Appendix D

116 Remembering the Giants
Paul Coffman’s Presentation Viewgraphs

Apollo Rocket Propulsion Development 117
Appendix D

118 Remembering the Giants
Paul Coffman’s Presentation Viewgraphs

Apollo Rocket Propulsion Development 119
Appendix D

120 Remembering the Giants
Paul Coffman’s Presentation Viewgraphs

Apollo Rocket Propulsion Development 121
Appendix D

122 Remembering the Giants
 Apollo Rocket Propulsion Development 123

Appendix E

G. R. “Jerry” Pfeifer’s Presentation Viewgraphs
Appendix E

124 Remembering the Giants
G. R. “Jerry” Pfeifer’s Presentation Viewgraphs 

Apollo Rocket Propulsion Development 125
Appendix E

126 Remembering the Giants
G. R. “Jerry” Pfeifer’s Presentation Viewgraphs 

Apollo Rocket Propulsion Development 127
Appendix E

128 Remembering the Giants
G. R. “Jerry” Pfeifer’s Presentation Viewgraphs 

Apollo Rocket Propulsion Development 129
Appendix E

130 Remembering the Giants
G. R. “Jerry” Pfeifer’s Presentation Viewgraphs 

Apollo Rocket Propulsion Development 131
Appendix E

132 Remembering the Giants
Apollo Rocket Propulsion Development 133

Appendix F

Tim Harmon’s SE-7 & SE-8 Presentation 

Viewgraphs
Appendix F

134 Remembering the Giants
Tim Harmon’s SE-7 & SE-8 Presentation Viewgraphs 

Apollo Rocket Propulsion Development 135
Appendix F

136 Remembering the Giants
Tim Harmon’s SE-7 & SE-8 Presentation Viewgraphs 

Apollo Rocket Propulsion Development 137
Appendix F

138 Remembering the Giants
Tim Harmon’s SE-7 & SE-8 Presentation Viewgraphs 

Apollo Rocket Propulsion Development 139
Appendix F

140 Remembering the Giants
Tim Harmon’s SE-7 & SE-8 Presentation Viewgraphs 

Apollo Rocket Propulsion Development 141

Apollo Rocket Propulsion Development 143

Appendix G

Clay Boyce’s Presentation Viewgraphs
Appendix G

144 Remembering the Giants
Clay Boyce’s Presentation Viewgraphs

Apollo Rocket Propulsion Development 145
Appendix G

146 Remembering the Giants
Clay Boyce’s Presentation Viewgraphs

Apollo Rocket Propulsion Development 147
Appendix G

148 Remembering the Giants
Clay Boyce’s Presentation Viewgraphs

Apollo Rocket Propulsion Development 149
Appendix G

150 Remembering the Giants
Apollo Rocket Propulsion Development 151

Appendix H

Jerry Elverum’s Presentation Viewgraphs
152 Remembering the Giants

Appendix H
 Apollo Rocket Propulsion Development 153

Jerry Elverum’s Presentation Viewgraphs
154 Remembering the Giants

Appendix H
 Apollo Rocket Propulsion Development 155

Jerry Elverum’s Presentation Viewgraphs
156 Remembering the Giants

Appendix H

Design of the Pintle Mechanism.
 Apollo Rocket Propulsion Development 157

Jerry Elverum’s Presentation Viewgraphs

Pintle injection thrust chamber illustration.
158 Remembering the Giants

Appendix H

Hot fire test facility
 Apollo Rocket Propulsion Development 159

Jerry Elverum’s Presentation Viewgraphs
160 Remembering the Giants

Appendix H
 Apollo Rocket Propulsion Development 161

Jerry Elverum’s Presentation Viewgraphs

Altitude simulation test stand
162 Remembering the Giants

Appendix H
 Apollo Rocket Propulsion Development 163

Jerry Elverum’s Presentation Viewgraphs
164 Remembering the Giants

Appendix H
 Apollo Rocket Propulsion Development 165

Jerry Elverum’s Presentation Viewgraphs
166 Remembering the Giants

Appendix H
 Apollo Rocket Propulsion Development 167

Jerry Elverum’s Presentation Viewgraphs
168 Remembering the Giants

Appendix H
 Apollo Rocket Propulsion Development 169

Jerry Elverum’s Presentation Viewgraphs
170 Remembering the Giants

Appendix H
Apollo Rocket Propulsion Development 171

Appendix I

Tim Harmon’s Lunar Ascent Engine 

Presentation Viewgraphs
Appendix I

172 Remembering the Giants
Tim Harmon’s Lunar Ascent Engine Presentation Viewgraphs

Apollo Rocket Propulsion Development 173
Appendix I

174 Remembering the Giants
Tim Harmon’s Lunar Ascent Engine Presentation Viewgraphs

Apollo Rocket Propulsion Development 175
Appendix I

176 Remembering the Giants
Tim Harmon’s Lunar Ascent Engine Presentation Viewgraphs

Apollo Rocket Propulsion Development 177
Appendix I

178 Remembering the Giants
Apollo Rocket Propulsion Development 179

Dr. Shamim Rahman of NASA Stennis Space 

Center opens the event by introducing NASA SSC 

Center Director Dr. Richard Gilbrech.

NASA SSC Center Director Dr. Richard Gilbrech 

welcomes attendees mostly from three NASA 

Centers (Stennis, Marshall, and Johnson).

Dr. Shamim Rahman of NASA and Mr. Steve 

Fisher of Rocketdyne introduce the Apollo 

propulsion speakers.

Appendix J

Event Photos and Highlights

April 25, 2006
Appendix J

180 Remembering the Giants

Mr. Fisher describes the background of this event.

Apollo propulsion speakers accompanied by 

their wives.

Mr. Bob Biggs describes the development of the 

Apollo F-1 booster engine.

Mr. Bob Biggs describes the start sequence of the 

Apollo F-1 booster engine. 
Event Photos and Highlights 

 Apollo Rocket Propulsion Development 181

Mr. Bob Biggs takes audience questions.

Mr. Pete Rodriguez, Director of NASA Marshall 

Test Lab, poses a question.

Mr. Paul Coffman, engineer on Saturn V J-2 

engine, enjoys some of the audience interactions.

Mr. Paul Coffman discusses development 

challenges on the J-2 engine.
Appendix J

182 Remembering the Giants

Audience members taking in the lessons from 

Apollo.

Mr. Paul Coffman takes audience questions.

Mr. Paul Coffman addresses questions.

Mr. Preston Jones, Manager of NASA Marshall 

Propulsion Department, poses a question on the 

J-2 engine.
Event Photos and Highlights 

 Apollo Rocket Propulsion Development 183

Mr. Bob Biggs, developer of the F-1 engine.

Mr. Jerry Pfeifer discusses development of Apollo 

hypergolic reaction control thrusters.

Mr. Jerry Pfeifer discusses Apollo reaction control 

thrusters variants.

NASA Stennis Test Complex engineers absorb 

lessons learned.
Appendix J

184 Remembering the Giants

Mr. Jerry Pfeifer takes audience questions.

Mr. Tim Harmon discusses his team’s propulsion 

contributions at Rocketdyne during Apollo.

Mr. Tim Harmon shares with audience a picture 

of a “younger Tim.”

Mr. Tim Harmon discusses his team’s propulsion 

contributions at Rocketdyne during Apollo.
Event Photos and Highlights 

 Apollo Rocket Propulsion Development 185

Mr. Clay Boyce of Aerojet is introduced to begin his 

discussion of the Apollo service module engine.

Mr. Clay Boyce, developer of the Apollo service 

propulsion system.

Mr. Clay Boyce discusses the engine value assembly.

Test engineer poses a question.
Appendix J

186 Remembering the Giants

Mr. Clay Boyce takes audience questions on 

Apollo service module propulsion.

Mr. Jerry Elverum discusses the Apollo lunar 

lander engine.

Mr. Robert Cort of NASA White Sands Test 

Facility attends as WSTF Propulsion Test Office 

chief.

Mr. Jerry Elverum, developer of the lunar 

lander engine.
Event Photos and Highlights 

 Apollo Rocket Propulsion Development 187

Mr. Miguel Rodriguez (left), Director of NASA 

SSC Engineering & Science Directorate, and his 

Deputy Randy Galloway (right) participate in the 

session along with engineer Jim Ryan (center).

Mr. Bill Arceneaux, center, head of Constellation 

Program Test and Verification Office.

Mr. Jerry Elverum describes the engine flow control 

valve and throttling ability.

Mr. Jerry Elverum takes audience questions.
Appendix J

188 Remembering the Giants

Mr. Tim Harmon gives a second presentation, 

this time on the lunar ascent engine.

Mr. Tim Harmon takes questions.

Audience members at the closing of the 

presentations.

Audience upon closing of the lecture that day.
Event Photos and Highlights 

 Apollo Rocket Propulsion Development 189

Several seminar presenters: Coffman, Elverum, Biggs and Boyce with their wives at the luncheon.

Tim Harmon with members of NASA SSC Senior Management at Von Braun Tower luncheon.

Shamim Rahman and wife, Shaheen, 

at luncheon.

R. Bilstein, Stages to Saturn: A Technological History of the Apollo/Saturn Launch Vehicles, 

NASA SP-4206, 1996 (out of print), now published by University of Florida Press, 2003.

 

R. S. Kraemer, and V. Wheelock, Rocketdyne: 

Powering Humans into Space (AIAA Education), 

American Institute of Aeronautics and Astronautics, 

2006.

A. Lawrie, Saturn 1/1B: The Complete 

Manufacturing and Test Records, Apogee Books, 

2002.

J. L. Sloop, Liquid Hydrogen as a Propulsion Fuel, 

1945-1959. NASA SP-4404, 1978. Out of print. 

G. P. Sutton, History of Liquid Propellant Rocket 

Engines, American Institute of Aeronautics and 

Astronautics, 2006.

G. P. Sutton and Biblarz, Rocket Propulsion 

Elements, 7th Edition, John Wiley and Sons, 2001.

A. Young, The Saturn V F-1 Engine: Powering 

Apollo into History, Springer Praxis, 2008.

Further Reading

Remembering the Giants 191

Saturn V, AS-501 Vehicle 

(NASA Image number: S67-50531)

 Apollo Rocket Propulsion Development 193

Ablative – Material that evaporates or vaporizes. 

Ablestar – The Thor-Ablestar, or Thor AbleStar, was an American expendable launch system 

consisting of a Thor missile, with an Ablestar-powered upper stage.

Actuate – To put into motion or action. 

Aerozine-50 – A 50/50 mix of hydrazine and unsymmetrical dimethylhydrazine, it was developed in the late 1950s as a storable, high-energy, hypergolic rocket fuel.

APG – Adverse pressure gradient.

Attitude – The orientation of a spacecraft relative to its direction of motion.

Cf - Thrust coefficient; a measure of the amplification of thrust due to gas expansion in a 

particular nozzle as compared with the thrust that would be exerted if the chamber pressure 

acted only over the throat area of the nozzle.

Calrod® – An elongated electrical heater unit, much like the coiled burner/heating element on 

an electric stove. The trademark name has become a generic moniker for any electric heater in 

the shape of a straight rod. 

Cavitation – The formation of vapor bubbles of a flowing liquid in a region where the pressure 

of the liquid falls below its vapor pressure. Mechanical forces - such as the moving blades of a 

ship’s propeller - can cause this.

Chlorine trifluoride (ClF3) – A colorless, poisonous, corrosive and very reactive gas that is 

primarily of interest as a component in rocket fuels; in industrial cleaning and etching operations in the semiconductor industry; in nuclear reactor fuel processing; and in other industrial 

operations.

Columbium – Also known as niobium, this chemical element is used as an alloying element 

in steels and superalloys.

Deep throttling – Refers to drastically reducing the thrust of an engine by restricting flow of 

propellants to less than 100 percent of capacity. The challenge is to do so without creating 

instability in the engine. Most engines are designed to deliver constant thrust. Many can be 

throttled down somewhat. For instance, the space shuttle main engine can be throttled down 

to 60 percent thrust without instability occurring. However, to design an engine that can be 

throttled down to 10 percent or less of optimal thrust is very challenging. 

Delta-P – Delta pressure refers to the difference between standard atmospheric air pressure 

and the air pressure read by a gauge or test equipment.

DIGR – Datelink Independent Gateway Retrofit 

Glossary
194

Glossary

Remembering the Giants

F-1 – An engine developed by Rocketdyne and used to power the Saturn V vehicle. Five F-1 

engines were used in the S-IC first stage of each Saturn V, which served as the main launch 

vehicle in the Apollo program. The F-1 is still the most powerful liquid-fueled rocket engine 

developed. 

Grumman – The Grumman Aircraft Engineering Corp. was founded in 1929 and grew into 

a leading twentieth-century American producer of military and civilian aircraft. It later was 

renamed the Grumman Aerospace Corp.. In 1994, it was acquired by Northrop Corp. to form 

Northrop Grumman.

Helmholtz resonator – Named for German physician and physicist Hermann Ludwig Ferdinand von Helmholtz, it is a container of gas (usually air) with an open hole (or neck or port). 

A volume of air in and near the open hole vibrates because of the “springiness” of the air inside. 

A common example is an empty bottle: the air inside vibrates when one blows across the top. 

Source: www.phys.unsw.edu.au/jw/Helmholtz.html

Hypergolic – Refers to an ignition process that involves a rocket propellant mixing with an 

oxidizer and igniting upon contact. No ignition spark is needed. In the process, a hypergolic 

fluid is hermetically sealed in a cartridge. As long as it is sealed, the fluid is stable. However, the 

fluid spontaneously ignites when it comes into contact with any form of oxygen. In starting an 

engine, increased fuel pressure within the igniter fuel system ruptures the cartridge. Fuel and 

the hypergolic fluid enter the chamber via separate lines. When the hypergolic fluid comes in 

contact with an oxidizer already present, it ignites, initiating fuel combustion.

Hysteresis – A lag of effect when the forces acting on a body are changed 

IPPD – Integrated Product and Process Development

IR&D – Independent Research and Development

ISP (specific impulse) – A measure of the “fuel efficiency” of a rocket or jet engine. It represents 

the impulse (change in momentum) per unit of propellant. The higher the specific impulse, 

the less propellant is needed to gain a given amount of momentum.Specific impulse is a useful 

value to compare engines, much like “miles per gallon” for cars. A propulsion method with a 

higher specific impulse is more propellant-efficient.

J-2 – An engine used for the second and third stages of the Saturn V moon rocket. The J-2 

engine used liquid hydrogen and liquid oxygen propellants for a maximum thrust of 225,000 

pounds. It was the first manned booster engine to use liquid hydrogen as a propellant and the 

first large booster engine designed to be restarted multiple times during a mission.

LEM – Lunar Excursion Module
195

Glossary

 Apollo Rocket Propulsion Development

lbf – Pounds of force

Mach – The speed of an object moving through air, divided by the speed of sound. Thus, 

something traveling Mach 7 is traveling seven times the speed of sound. According to the 

aerospaceweb.org Web site, the generally accepted speed of sound is 1,116.4 feet per second 

(or 761.2 miles per hour)

Mandrel – A spindle or axle used to support material being machined or milled; a core around 

which materials may be cast and shaped, or a shaft on which a working tool is mounted

Man-rated – A term used to signify that a spacecraft, launch vehicle or airplane is worthy of 

transporting humans. Referred to as “human-rated” in most instances now.

MMH – Monomethylhydrazine, a volatile hydrazine chemical used as a fuel in bi-propellant 

rocket engines and frequently in hypergolic mixtures.

Molybdenum – A silvery metal that has the sixth-highest melting point of any element. It readily forms hard, stable carbides and often is used in high-strength steel alloys. 

N2

O4 – Dinitrogen tetroxide (nitrogen tetroxide or nitrogen peroxide). It is a powerful oxidizer, highly toxic and corrosive. N2

O4 is hypergolic with various forms of hydrazine, i.e., they 

burn on contact without a separate ignition source, making them popular bi-propellant rocket 

fuels

North American Aviation Inc. – Founded in 1928, North American was a major American 

aircraft manufacturer, responsible for such crafts as the P-51 Mustang fighter, the F-86 Savre 

jet fighter, the Apollo Command and Service modules and the space shuttle orbiter. Through 

various mergers, it now is part of The Boeing Company.

OME – Orbital maneuvering engine

Pintle – A a pin or bolt used as part of a pivot or hinge. In rocketry, a pintle engine uses a 

single-feed fuel injector rather than the hundreds of smaller holes used in a typical rocket 

engine. That lowers the cost of engine manufacture while surrendering some performance. 

The concept originally was developed by the Grumman Aerospace Corp. for the Apollo Lunar 

Module. 

PFRT – Preliminary Flight Rating Test

Psi – Pounds per square inch (measure of pressure) 

Reaction Motors Inc. – Formed in 1941 by four members of the American Rocket Society, the 

company developed the rocket engines used by the early X planes, including the Bell X-1 that 

first broke the sound barrier in 1947 and the famous North American X-15. The company 
196

Glossary

Remembering the Giants

was taken over by Thiokol Chemical in 1958 to become the Reaction Motor Division. Later, 

the division became Thiokol Propulsion, which was acquired by ATK (Alliant Techsystems) 

in 2001.

Retort – A closed laboratory vessel with an outlet tube, used for distillation, sublimation, or 

decomposition of substances by heat. 

RFNA – Red fuming nitric acid, a storable oxidizer used as a rocket propellant. It consists 

mainly of nitric acid, also containing dinitrogen tetroxide and water. 

RP-1 – Rocket propellant 1, a highly refined form of kerosene used as a rocket fuel, usually 

with liquid oxygen as an oxidizer. 

SSME – Space shuttle main engine

Scarfed nozzle – A two-dimensional, non-symmetric nozzle with a flat lower wall (cowl) that 

is truncated to save weight. 

SPS – Service Propulsion System. In the Apollo Program, it was used to place spacecraft into 

and out of lunar orbit, and for mid-course corrections between the Earth and the moon. The 

Apollo SPS used a single AJ10-137 engine.

Theater High Altitude Area Defense (THAAD) – Formerly known as the Terminal High Altitude Area Defense, ut is a U.S. Army project to develop a system to shoot down short- and 

medium-range ballistic missiles. 

TRW Inc. – a company active in the early development of missile systems and spacecraft, most 

notably the NASA deep space satellites Pioneer 10 and 11. TRW also pioneered systems engineering. The 1958 merger of Thompson Products Co. with the Ramo-Wooldridge Corp. was 

named Thompson Ramo Wooldridge Inc., then shortened to TRW Inc. in 1965.

Unsymmetrical dimethylhydrazine (UDMH) – A toxic, volatile, clear liquid that turns yellowish on exposure to air and absorbs oxygen and carbon dioxide. It is also hygroscopic, meaning 

it attracts water molecules from the surrounding environment. It mixes completely with water, 

ethanol, and kerosene. In concentrations between 2.5 percent and 95 percent in air, its vapors 

are flammable. It often is used in hypergolic rocket fuels because it is stable and can be kept 

loaded in rocket fuel systems for long periods. UDMH functions as a starter fuel to start combustion and warm the rocket engine prior to switching to kerosene.

Venturi – A tube with a tapered constriction that causes an increase in the velocity of flow of 

a fluid and a corresponding decrease in fluid pressure; a specially shaped tube that affects the 

speed and pressure of a flowing fluid
 Apollo Rocket Propulsion Development 197

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Mudgway, Douglas J. Uplink-Downlink: A History of the Deep Space Network, 1957–1997. 

NASA SP-2001-4227.

Dawson, Virginia P., and Mark D. Bowles. Taming Liquid Hydrogen: The Centaur Upper Stage 

Rocket, 1958–2002. NASA SP-2004-4230.

Meltzer, Michael. Mission to Jupiter: A History of the Galileo Project. NASA SP-2007-4231.

Heppenheimer, T. A. Facing the Heat Barrier: A History of Hypersonics. NASA SP-2007-4232.

Tsiao, Sunny. “Read You Loud and Clear!” The Story of NASA’s Spaceflight Tracking and Data 

Network. NASA SP-2007-4233.

CENTER HISTORIES, NASA SP-4300:

Rosenthal, Alfred. Venture into Space: Early Years of Goddard Space Flight Center. 

NASA SP-4301, 1985. 

Hartman, Edwin, P. Adventures in Research: A History of Ames Research Center, 1940–1965. 

NASA SP-4302, 1970. 

Hallion, Richard P. On the Frontier: Flight Research at Dryden, 1946–1981. NASA SP-4303, 

1984. 

Muenger, Elizabeth A. Searching the Horizon: A History of Ames Research Center, 1940–1976. 

NASA SP-4304, 1985. 

Hansen, James R. Engineer in Charge: A History of the Langley Aeronautical Laboratory, 

1917–1958. NASA SP-4305, 1987. 

Dawson, Virginia P. Engines and Innovation: Lewis Laboratory and American Propulsion 

Technology. NASA SP-4306, 1991.

Dethloff, Henry C. “Suddenly Tomorrow Came . . .”: A History of the Johnson Space Center, 

1957–1990. NASA SP-4307, 1993. 
Hansen, James R. Spaceflight Revolution: NASA Langley Research Center from Sputnik to 

Apollo. NASA SP-4308, 1995.

Wallace, Lane E. Flights of Discovery: An Illustrated History of the Dryden Flight Research 

Center. NASA SP-4309, 1996.

Herring, Mack R. Way Station to Space: A History of the John C. Stennis Space Center. 

NASA SP-4310, 1997.

Wallace, Harold D., Jr. Wallops Station and the Creation of an American Space Program. 

NASA SP-4311, 1997.

Wallace, Lane E. Dreams, Hopes, Realities. NASA’s Goddard Space Flight Center: The First Forty 

Years. NASA SP-4312, 1999.

Dunar, Andrew J., and Stephen P. Waring. Power to Explore: A History of Marshall Space 

Flight Center, 1960–1990. NASA SP-4313, 1999.

Bugos, Glenn E. Atmosphere of Freedom: Sixty Years at the NASA Ames Research Center. 

NASA SP-2000-4314, 2000.

Schultz, James. Crafting Flight: Aircraft Pioneers and the Contributions of the Men and Women 

of NASA Langley Research Center. NASA SP-2003-4316, 2003.

Bowles, Mark D. Science in Flux: NASA’s Nuclear Program at Plum Brook Station, 1955–

2005. NASA SP-2006-4317.

Wallace, Lane E. Flights of Discovery: An Illustrated History of the Dryden Flight Research 

Center. NASA SP-4318, 2007. Revised version of SP-4309.

GENERAL HISTORIES, NASA SP-4400:

Corliss, William R. NASA Sounding Rockets, 1958–1968: A Historical Summary. 

NASA SP-4401, 1971. 

Wells, Helen T., Susan H. Whiteley, and Carrie Karegeannes. Origins of NASA Names. 

NASA SP-4402, 1976. 

Anderson, Frank W., Jr. Orders of Magnitude: A History of NACA and NASA, 1915–1980. 

NASA SP-4403, 1981. 

Sloop, John L. Liquid Hydrogen as a Propulsion Fuel, 1945–1959. NASA SP-4404, 1978. 

Roland, Alex. A Spacefaring People: Perspectives on Early Spaceflight. NASA SP-4405, 1985. 

Bilstein, Roger E. Orders of Magnitude: A History of the NACA and NASA, 1915–1990. 

NASA SP-4406, 1989. 

Logsdon, John M., ed., with Linda J. Lear, Jannelle Warren Findley, Ray A. Williamson, and 

Dwayne A. Day. Exploring the Unknown: Selected Documents in the History of the U.S. Civil 

Space Program, Volume I, Organizing for Exploration. NASA SP-4407, 1995.
Logsdon, John M., ed, with Dwayne A. Day, and Roger D. Launius. Exploring the Unknown: 

Selected Documents in the History of the U.S. Civil Space Program, Volume II, External 

Relationships. NASA SP-4407, 1996.

Logsdon, John M., ed., with Roger D. Launius, David H. Onkst, and Stephen J. Garber. 

Exploring the Unknown: Selected Documents in the History of the U.S. Civil Space Program, 

Volume III, Using Space. NASA SP-4407,1998.

Logsdon, John M., ed., with Ray A. Williamson, Roger D. Launius, Russell J. Acker, Stephen 

J. Garber, and Jonathan L. Friedman. Exploring the Unknown: Selected Documents in the 

History of the U.S. Civil Space Program, Volume IV, Accessing Space. NASA SP-4407, 1999.

Logsdon, John M., ed., with Amy Paige Snyder, Roger D. Launius, Stephen J. Garber, and 

Regan Anne Newport. Exploring the Unknown: Selected Documents in the History of the U.S. 

Civil Space Program, Volume V, Exploring the Cosmos. NASA SP-4407, 2001.

Logsdon, John M., ed., with Stephen J. Garber, Roger D. Launius, and Ray A. Williamson. 

Exploring the Unknown: Selected Documents in the History of the U.S. Civil Space Program, 

Volume VI: Space and Earth Science. NASA SP-2004-4407, 2004.

Logsdon, John M., ed., with Roger D. Launius. Exploring the Unknown: Selected Documents 

in the History of the U.S. Civil Space Program, Volume VII: Human Spaceflight: Projects 

Mercury, Gemini, and Apollo. NASA SP-2008-4407, 2008.

Siddiqi, Asif A., Challenge to Apollo: The Soviet Union and the Space Race, 1945–1974. 

NASA SP-2000-4408, 2000.

Hansen, James R., ed. The Wind and Beyond: Journey into the History of Aerodynamics in 

America, Volume 1, The Ascent of the Airplane. NASA SP-2003-4409, 2003.

Hansen, James R., ed. The Wind and Beyond: Journey into the History of Aerodynamics in 

America, Volume 2, Reinventing the Airplane. NASA SP-2007-4409, 2007.

Hogan, Thor. Mars Wars: The Rise and Fall of the Space Exploration Initiative. NASA 

SP-2007-4410, 2007.

MONOGRAPHS IN AEROSPACE HISTORY (SP-4500 SERIES):

Launius, Roger D., and Aaron K. Gillette, compilers. Toward a History of the Space Shuttle: 

An Annotated Bibliography. Monograph in Aerospace History, No. 1, 1992.

Launius, Roger D., and J. D. Hunley, compilers. An Annotated Bibliography of the Apollo 

Program. Monograph in Aerospace History No. 2, 1994.

Launius, Roger D. Apollo: A Retrospective Analysis. Monograph in Aerospace History, No. 3, 

1994.

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

Rendezvous Concept. Monograph in Aerospace History, No. 4, 1995.
204

NASA History Series

Remembering the Giants

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

History, No. 5, 1996.

Powers, Sheryll Goecke. Women in Flight Research at NASA Dryden Flight Research Center 

from 1946 to 1995. Monograph in Aerospace History, No. 6, 1997.

Portree, David S. F., and Robert C. Trevino. Walking to Olympus: An EVA Chronology. 

Monograph in Aerospace History, No. 7, 1997.

Logsdon, John M., moderator. Legislative Origins of the National Aeronautics and Space Act 

of 1958: Proceedings of an Oral History Workshop. Monograph in Aerospace History, No. 8, 

1998. 

Rumerman, Judy A., compiler. U.S. Human Spaceflight, A Record of Achievement 1961–1998. 

Monograph in Aerospace History, No. 9, 1998.

Portree, David S. F. NASA’s Origins and the Dawn of the Space Age. Monograph in Aerospace 

History, No. 10, 1998.

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

Station. Monograph in Aerospace History, No. 11, 1998.

Phillips, W. Hewitt. Journey in Aeronautical Research: A Career at NASA Langley Research 

Center. Monograph in Aerospace History, No. 12, 1998.

Braslow, Albert L. A History of Suction-Type Laminar-Flow Control with Emphasis on Flight 

Research. Monograph in Aerospace History, No. 13, 1999.

Logsdon, John M., moderator. Managing the Moon Program: Lessons Learned From Apollo. 

Monograph in Aerospace History, No. 14, 1999.

Perminov, V. G. The Difficult Road to Mars: A Brief History of Mars Exploration in the Soviet 

Union. Monograph in Aerospace History, No. 15, 1999. 

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

Monograph in Aerospace History, No. 16, 1999.

Maisel, Martin, Demo J.Giulanetti, and Daniel C. Dugan. The History of the XV-15 Tilt Rotor 

Research Aircraft: From Concept to Flight. Monograph in Aerospace History, No. 17, 2000. 

NASA SP-2000-4517.

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

Airplane. Monograph in Aerospace History, No. 18, 2000. NASA SP-2000-4518.

Chambers, Joseph R. Partners in Freedom: Contributions of the Langley Research Center to U.S. 

Military Aircraft of the 1990s. Monograph in Aerospace History, No. 19, 2000. 

NASA SP-2000-4519.

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

Research Center. Monograph in Aerospace History, No. 20, 2000. NASA SP-2000-4520.
205

NASA History Series

Apollo Rocket Propulsion Seminar

Portree, David S. F. Humans to Mars: Fifty Years of Mission Planning, 1950–2000. 

Monograph in Aerospace History, No. 21, 2001. NASA SP-2001-4521.

Thompson, Milton O., with J. D. Hunley. Flight Research: Problems Encountered and What 

they Should Teach Us. Monograph in Aerospace History, No. 22, 2001. NASA SP-2001-

4522.

Tucker, Tom. The Eclipse Project. Monograph in Aerospace History, No. 23, 2001. 

NASA SP-2001-4523.

Siddiqi, Asif A. Deep Space Chronicle: A Chronology of Deep Space and Planetary Probes 

1958–2000. Monograph in Aerospace History, No. 24, 2002. NASA SP-2002-4524.

Merlin, Peter W. Mach 3+: NASA/USAF YF-12 Flight Research, 1969–1979. Monograph in 

Aerospace History, No. 25, 2001. NASA SP-2001-4525.

Anderson, Seth B. Memoirs of an Aeronautical Engineer: Flight Tests at Ames Research Center: 

1940–1970. Monograph in Aerospace History, No. 26, 2002. NASA SP-2002-4526.

Renstrom, Arthur G. Wilbur and Orville Wright: A Bibliography Commemorating the OneHundredth Anniversary of the First Powered Flight on December 17, 1903. Monograph in 

Aerospace History, No. 27, 2002. NASA SP-2002-4527.

No monograph 28.

Chambers, Joseph R. Concept to Reality: Contributions of the NASA Langley Research Center to 

U.S. Civil Aircraft of the 1990s. Monograph in Aerospace History, No. 29, 2003. SP-2003-

4529.

Peebles, Curtis, editor. The Spoken Word: Recollections of Dryden History, The Early Years. 

Monograph in Aerospace History, No. 30, 2003. SP-2003-4530.

Jenkins, Dennis R., Tony Landis, and Jay Miller. American X-Vehicles: An Inventory- X-1 to 

X-50. Monograph in Aerospace History, No. 31, 2003. SP-2003-4531.

Renstrom, Arthur G. Wilbur and Orville Wright: A Chronology Commemorating the OneHundredth Anniversary of the First Powered Flight on December 17, 1903. Monograph in 

Aerospace History, No. 32, 2003. NASA SP-2003-4532.

Bowles, Mark D., and Robert S. Arrighi. NASA’s Nuclear Frontier: The Plum Brook Research 

Reactor. Monograph in Aerospace History, No. 33, 2004. (SP-2004-4533).

Wallace, Lane and Christian Gelzer. Nose Up: High Angle-of-Attack and Thrust Vectoring 

Research at NASA Dryden, 1979-2001. Monograph in Aerospace History No. 34, 2009. 

NASA SP-2009-4534.

Matranga, Gene J., C. Wayne Ottinger, Calvin R. Jarvis, and D. Christian Gelzer. 

Unconventional, Contrary, and Ugly: The Lunar Landing Research Vehicle. Monograph in 

Aerospace History, No. 35, 2006. NASA SP-2004-4535.
206

NASA History Series

Remembering the Giants

McCurdy, Howard E. Low Cost Innovation in Spaceflight: The History of the Near Earth 

Asteroid Rendezvous (NEAR) Mission. Monograph in Aerospace History, No. 36, 2005. NASA 

SP-2005-4536.

Seamans, Robert C., Jr. Project Apollo: The Tough Decisions. Monograph in Aerospace History, 

No. 37, 2005. NASA SP-2005-4537.

Lambright, W. Henry. NASA and the Environment: The Case of Ozone Depletion. Monograph 

in Aerospace History, No. 38, 2005. NASA SP-2005-4538.

Chambers, Joseph R. Innovation in Flight: Research of the NASA Langley Research Center on 

Revolutionary Advanced Concepts for Aeronautics. Monograph in Aerospace History, No. 39, 

2005. NASA SP-2005-4539. 

Phillips, W. Hewitt. Journey Into Space Research: Continuation of a Career at NASA Langley 

Research Center. Monograph in Aerospace History, No. 40, 2005. NASA SP-2005-4540. 

Rumerman, Judy A., Chris Gamble, and Gabriel Okolski, compilers. U.S. Human 

Spaceflight: A Record of Achievement, 1961–2006. Monograph in Aerospace History No. 41, 

2007. NASA SP-2007-4541.

Dick, Steven J.; Garber, Stephen J.; and Odom, Jane H. Research in NASA History. 

Monograph in Aerospace History No. 43, 2009. NASA SP-2009-4543.

Merlin, Peter W., Ikhana: Unmanned Aircraft System Western States Fire Missions. 

Monograph in Aerospace History #44. NASA SP-2009-4544.

ELECTRONIC MEDIA (SP-4600 SERIES)

Remembering Apollo 11: The 30th Anniversary Data Archive CD-ROM. NASA SP-4601, 

1999. 

Remembering Apollo 11: The 35th Anniversary Data Archive CD-ROM. NASA SP-2004-4601, 

2004. This is an update of the 1999 edition.

The Mission Transcript Collection: U.S. Human Spaceflight Missions from Mercury Redstone 3 to 

Apollo 17. SP-2000-4602, 2001. 

Shuttle-Mir: the United States and Russia Share History’s Highest Stage. NASA SP-2001-4603, 

2002. 

U.S. Centennial of Flight Commission presents Born of Dreams ~ Inspired by Freedom. 

NASA SP-2004-4604, 2004.

Of Ashes and Atoms: A Documentary on the NASA Plum Brook Reactor Facility. 

NASA SP-2005-4605.

Taming Liquid Hydrogen: The Centaur Upper Stage Rocket Interactive CD-ROM. 

NASA SP-2004-4606, 2004.
207

NASA History Series

Apollo Rocket Propulsion Seminar

Fueling Space Exploration: The History of NASA’s Rocket Engine Test Facility DVD. 

NASA SP-2005-4607.

Altitude Wind Tunnel at NASA Glenn Research Center: An Interactive History CD-ROM. 

NASA SP-2008-4608.

CONFERENCE PROCEEDINGS (SP-4700 SERIES)

Dick, Steven J., and Keith Cowing, ed. Risk and Exploration: Earth, Sea and the Stars. 

NASA SP-2005-4701.

Dick, Steven J., and Roger D. Launius. Critical Issues in the History of Spaceflight. 

NASA SP-2006-4702.

Dick, Steven J., ed. Remembering the Space Age: Proceedings of the 50th Anniversary 

Conference. NASA SP-2008-4703.

SOCIETAL IMPACT (SP-4800 SERIES)

Dick, Steven J., and Roger D. Launius. Societal Impact of Spaceflight. NASA SP-2007-4801.