AMPTIAC is a DOD Information Analysis Center Administered by the 

Defense Information Systems Agency, Defense Technical Information Center 
History is replete with examples of armies and empires whose

military success was a direct result of their technical supremacy.

The Roman Empire’s expertise in road building allowed them

to keep their armies supplied during campaigns. Half of the

world’s people were subjects of the British Empire at its pinnacle; due in large part to their ships, which enabled them to

project force around the globe. Most recently, America assumed

the promontory of military superiority at the beginning of the

twentieth century. This was a direct product of America’s industrial and technology edge over competitor nations; and was

concurrent with the ascendancy of our air power. This first

became apparent in World War I, and then again in World War

II, when the Army discovered it could greatly extend its reach

into enemy territory using the Army Air Corps. The Navy

experienced similar success in the Pacific War when it shifted its

emphasis from battleships to aircraft carriers, which helped

turn the tide against Japan.

In the years following the Second World War, the arrival of the

Space Age was highly anticipated. The pioneering work by the

US Military and Werner Von Braun in the 1950s (along with

Sergei Korolev’s parallel efforts in the Soviet Union) was performed to be the “first to space.” Even before the first successful

launch in 1957, the strategic value of space was understood.

Commanding the high ground of space has long been an Air

Force priority. Shortly after the launch of Sputnik, Air Force

Chief of Staff General Thomas D. White declared that the Air

Force “must win the capability to control space.” Since that

time, space systems have played an increasingly larger role in

the Air Force’s overall mission, culminating with the establishment of the Space Command in 1982.

The Space Command, and by extension much of the Air

Force’s space technology got its first “test under fire” during the

1991 Persian Gulf War; when its satellites provided critical

communications for the US Central Command, both in-theater and with the Pentagon. Global Positioning Satellites (GPS)

were able to provide our armed forces with precise position

information of key targets. Furthermore, Space Command’s

GPS and weather satellites guided ground units across the featureless expanses of the Iraqi desert in all weather. In addition,

the Air Force’s early warning satellites gave a much needed

“heads-up” to Army Patriot missile batteries repelling incoming

Scud attacks. Space technology had come of age – the Gulf

War was truly the “First Space War.”

This brings us to the present day – the impact and versatility of military space systems continue to grow, and just in time.

If the Gulf War was the first space war, then the current Global

War on Terror is most certainly the second. In addition to the

functions of weather, navigation, communications, and early

warning, the latest generation of tactical satellites are also capable of providing guidance, and even control of certain military

systems. The new wave of unmanned aerial vehicles, such as

Predator and Global Hawk, utilize satellites for guidance and

communications. The same is true for weaponry; like the new

Joint Direct Attack Munition (JDAM), recently used in

Afghanistan and Iraq. American Special Forces depend on satellites to provide navigation, surveillance, communication, and

intelligence while on missions. In the past decade, military

space capabilities have evolved from a supporting role to an

indispensable part of the warfighter. In the future, some space

systems may even become the warfighter.

How will America maintain its ‘Space Superiority’ in the

future? The best way is to preserve the technology edge that has

brought us to this point. The Air Force Research Laboratory’s

Materials and Manufacturing Directorate (AFRL/ML) has

demonstrated its value as the focal point for originating new

materials technologies for space applications. While war tends

to draw our attention away from technology development, it

also makes us thankful for those same technologies, which have

now been called upon to protect the nation. It is only through

the continued development of new and enabling technologies

that we will continue to command the high ground in the space

wars to come.

Christian E. Grethlein

Editor-in-Chief

Editorial: Seizing the High Ground

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Section I: The Air Force’s Role in Space Materials Research – Past, Present, and Future

The Air Force has a long and recognized history of advancing the state-of-the-art in space technology. The Air Force Research

Laboratory’s Materials and Manufacturing Directorate has played a major role in this effort by developing the technologies to enable

advanced space systems.

Defending the Nation through Excellence in Space: 

A Conversation with AFRL/ML’s Director ... 5

Dr. Charles Browning, the Director of the Air Force Research Laboratory’s Materials and Manufacturing Directorate

(AFRL/ML), was recently able to share his perspective on the nature and direction of AFRL/ML’s contributions to space

technology development.

A Legacy of Achievement ... 11

Mr. Paul M. Propp was a witness to nearly four decades of space-related materials research at AFRL/ML, where he worked

from the early 1960s through the mid-1990s. He shares some of his recollections about AFRL/ML’s achievements and

major milestones.

Section II: The Space Environment and Materials Testing

This is probably the best starting place for the uninitiated, who want to learn more about space, and the unique challenges it poses

to professionals developing the materials and systems that space vehicles need to carry out their missions.

Materials and the ‘Final Frontier’ ... 17

Satellites, probes, and launch vehicles share very few common design and performance issues with earth-based systems

operating on land, sea, or in the air. This article will introduce the reader to the basic concepts of space and space technology, and thus may be the right place to start their reading.

MaterialEASE: Testing in the Space Environment ... 25

How do we know that a selected material will work in space? The only way to know is to test it. This article explores some

of the conditions of space environments and how materials and components are tested, both at earthbound facilities, and

in space.

Laser Hardened Materials Evaluation Laboratory Simulates 

Space Environment for Advance Materials, Space Systems Testing ... 31

Short of taking material samples into space, evaluation of their properties must be performed in earthbound test facilities.

The Air Force Research Laboratory (AFRL) is home to one such state of the art facility that serves as a resource to both

the military and contractors.

Section III: Technology Development for Launch and Reentry Applications

For space systems, “getting there” is more than half the battle. Anything to reduce a launch vehicle’s weight or increase its performance can go a long way to extending the capabilities of satellites or reducing the costs of putting them up. Moreover, mission

success is also a function of how well systems are protected from the extremes of the launch, orbital, and reentry environments. 

Doubling our Reach to Space ... 35

The Integrated High-Payoff Rocket Propulsion Technology (IHPRPT) program is a collaborative effort initiated by

DOD, NASA, and industry; with the primary objective of doubling rocket propulsion capability by 2010. The program

leverages the commonalities of military, civil, and commercial rockets to enhance and increase the opportunities for joint

research and development.

Issue Focus: 

Air Force Research 

Enables Enhanced Launch 

and Space Systems
Ceramic Materials for Reusable Liquid Fueled Rocket Engine 

Combustion Devices ... 39

AFRL/ML and its industry partners are developing lightweight ceramic materials and manufacturing processes for use 

in liquid fueled rocket engine combustion devices (thrust chambers and nozzles) for next-generation, reusable launch 

vehicles. These lightweight ceramics could potentially reduce the weight of combustion devices up to 50%.

Cooperative Planning is Accelerating the Development of Advanced TPS ... 45

Thermal protection systems (TPS) are materials that are used to protect launch vehicles, reentry vehicles, and spacecraft

from exposure to high temperatures experienced during operation. Future launch vehicles and hypersonic weapon systems

are envisioned that will be far easier to maintain than the current state-of-the-art.

Section IV: Technology Development for Satellite Applications

Systems that operate in the extremes of space have robust and innovative materials and manufacturing processes to thank for 

their success. AFRL/ML has been a leader in developing, testing, and fielding the state of the art materials and processes needed to

fabricate the myriad of subsystems that make up the modern military satellite.

Composites for Orbiting Platforms ... 51

The cost of placing objects into low earth orbit is roughly $10,000 per pound, thus weight is closely managed during

design. Lighter-weight support subsystems allow additional useful payload mass, which means more performance or service life for satellites. The Nonmetallic Materials Division’s Structural Materials Branch (AFRL/MLBC) has led the way by

developing composite materials to reduce the structural weight of satellites and their subsystems.

Spacecraft Materials Development Programs for Thermal Control Coatings 

and Space Environmental Testing ... 59

Protecting spacecraft and their working components from the harmful effects of extreme temperatures is no simple task.

This is the main function of a thermal control system (TCS). A well-designed and properly operating TCS is essential to

protecting the critical subsystems of any spacecraft.

Advanced Materials and Processes for Large, Lightweight, 

Space-Based Mirrors ... 67

New material solutions and manufacturing processes are required to meet the Air Force’s directed energy weapons, 

reconnaissance/surveillance, and secured communication needs. This article reviews some traditional materials and

processes used to fabricate mirrors for space applications, and also examines new material concepts being considered 

for use in space-based mirrors.

Precision Tooling for Thin Film (Membrane) Reflectors ... 75

This article describes the innovative design approach, fabrication processes, and materials that SRS Technologies and

AFRL/ML are developing for precision tooling. These developments are directed toward the production of polymer-based

thin film (membrane) reflectors for use as inflatable light weight antennas, solar power concentrators, and imaging optics.

Acknowledgements

Does our cover look familiar? It should to many of our readers. It also happens to be the poster artwork

for this year’s National Space and Missile Materials Symposium (see the announcement later in this

issue). We would like to thank the Anteon Corporation for allowing us to use it, and more specifically, to Ms. Rae Parks who created it.

We would also like to thank the following AFRL personnel for their support and guidance in the preparation of this issue: Mr. George Schmitt, Chief of the Integration and

Operations Division; Mr. Fredrick Coleman at the Materials and Manufacturing

Directorate's Technology Information Center; and Mr. Dan Cleyrat at the Materials

and Manufacturing Directorate's Space Office.

Lastly, AMPTIAC would like to extend our sincere appreciation to

all our authors and contributors who helped to make this another great

issue!
The AMPTIAC Quarterly, Volume 8, Number 1 5

Unlike aircraft applications where hundreds of systems might

be procured, most space applications do not require many

systems. Hence, far fewer numbers of components or structures are needed – yet materials are critical to those applications. How does ML

balance the need for developing critical

technologies intended for a single space

system vs. a host of air vehicles or engines?

If a specific benefit greatly enhances, or in

many cases enables a desired capability for

the customer, then obviously we may want

to pursue investment in the area. It can’t be

viewed strictly as a numbers game. The

strategic or tactical value/impact of a space

system can be immense: so much so that the

return on investment is significant, even

when there’s only one system procured.

Materials are fundamental to a component’s capability – that is, they endow components with the characteristics needed to

meet the demands of the application. This is

very evident in the case of spacecraft, which

are produced in very limited numbers.

When a required mission capability is determined to be dependent on new materials technology (and thus drives the need for

improved materials), then further investment is justified.

Again comparing to aircraft systems, ML has historically been

involved in developing new manufacturing methods that help

reduce the cost of structures for these applications. This is 

driven by the fact that production runs for aircraft are normally

quite large. Do you have any manufacturing programs in-place

to help reduce the cost of structures and components for space

systems that will be procured in very limited quantities?

ML has been active in space materials since the days immediately after Sputnik was launched. Our past and present range of

activities includes research and development of structural 

materials based on our expertise in composites, nonstructural

materials such as thermal control coatings, lubricants, and polymeric films. In addition, it also includes electronic packaging and

low cost manufacturing of space components, as well as sensor

materials for surveillance systems. Our lean manufacturing 

initiatives to address affordable components

and systems (when few are being made) are

particularly well suited for space systems,

although many are applicable to aircraft

production as well.

Balancing limited resources is always a

challenge, especially when multiple issues

must be addressed. Within the space 

sector itself, what relative emphasis does

ML place on launch, satellite, and force

application technologies?

In the launch area, we have heavily invested in rocket propulsion materials for boost

and orbit transfer. For liquids, this includes

materials for turbopumps, combustion

chambers, nozzles, and lines/ducts/valve.

For solids, insulation and nozzles.

Moreover, we have invested extensively in

both chemical and electric on-board

propulsion systems for spacecraft. The propulsion area represents the largest stake in a single area within ML’s space enterprise. ML also supports spacecraft with investments spread

across bus and payload technologies. In the force application

area, ML has a long legacy in reentry thermal protection 

materials and continues to support the area with materials development, analysis, and testing.

To follow up on our last question, what are the critical 

material and process technologies that the Materials and

Manufacturing Directorate is currently developing for space?

The Materials and Manufacturing Directorate has a robust

technology development program that spans all materials 

classes for both satellite and space launch applications. ML is

addressing many needs for the functional areas of satellites. For

satellite propulsion, we are developing refractory metals, 

Recently, Dr. Charles Browning, the Director of the Materials and Manufacturing Directorate, at the Air Force Research Laboratory

(AFRL/ML), was able to share with AMPTIAC some of his reflections and insights on the contributions of AFRL/ML (and to a greater

extent, that of the Air Force) to space technology. Our conversation with Dr. Browning delves into the various aspects of ML’s history, 

philosophy, accomplishments, and strategies to meet and surmount the challenges of prevailing in the space environment; and how these

technologies are critical to national defense.

AMPTIAC would like to thank Dr. Browning for his time and consideration in consenting to this interview. After reading this interview and the rest of this issue of AMPTIAC Quarterly, we believe that our readers will take away a greater appreciation of the critical role

that materials play in enabling space technology, and how AFRL/ML, as a technology leader, contributes to the nation’s defense. - Editor 

DR. CHARLES E. BROWNING

Director of the Materials 

and Manufacturing Directorate

The Air Force’s Role in Space Materials Research – Past, Present, and Future Section I
6 The AMPTIAC Quarterly, Volume 8, Number 1

ceramic matrix composites, and carbon-carbon materials in

both traditional and new propulsion systems. We are addressing the continued need for lightweight power generation and

distribution. ML is supporting surveillance and reconnaissance

missions with the development of a variety of sensor materials.

We have a rich and ongoing history of providing materials for

structures and mechanisms, including space lubricants. For

communication applications, we are developing wide band-gap

semiconductors and electro-optic polymers. Thermal management is increasingly difficult for satellite components with

higher power densities. As such, we are continuing to develop

highly thermally conductive materials and advanced thermal

control coatings.

For launch vehicles, we are focusing our development efforts

on the subsystems that will be enabled by advanced materials.

For example, ML is developing a number of thermal protection

materials such as carbon-carbon, ceramic, and metal matrix

composites. Rocket propulsion is another area where materials

will define the performance of various propulsion concepts.

Cryogenic tanks will require lightweight materials that are

compatible with liquid oxygen and hydrogen. The recent

Shuttle disaster has highlighted the need for vehicle health

monitoring and we are just beginning to examine the requirements of needed materials.

Over the past several years there seems to be a drive towards

larger satellites, but at the same time micro- or nano-satellites

are being considered. What emphasis does ML place on

these dual trends?

Our program is well suited to support both areas, as materials

are fundamental to all applications. Our historical and continued work in organic matrix composites has an area of emphasis

to provide for lighter-weight stable structures, as well as multifunctionality. These developments will not only be of benefit to

large spacecraft, but to smaller ones as well. 

Additionally, we have efforts supporting large, lightweight

deployable structures with application to mirrors or antennas.

For small satellites, our work in thermal management materials

is crucial. We are developing carbon-carbon materials that will

provide improved thermal efficiency for thermal planes, heat

sinks, and radiators. This has particular impact on denser 

electronic systems as spacecraft become smaller. Additionally,

our work in advanced coatings and lubrication will provide

added benefit to either.

Our biomimetics and nanomaterials research has great

potential benefit to small spacecraft. Technology impact areas

include smaller lightweight sensors, adaptive materials, room

temperature IR detectors, EM shielding, and high performance

tethers to name a few.

It seems as if obtaining material requirements for future 

aircraft systems would be a straightforward process, since 

the Air Force’s aircraft program offices are located with you

at Wright Patterson AFB. How do you successfully gather the

requirements needed to establish new ‘space’ materials 

development programs when the program offices dealing 

with space related technologies are scattered among 

several locations around the country?

As you know, we established Integrated Application Areas

(IAAs) that are in part charged with working closely with customers to obtain requirements. The Space IAA has initiated

and fostered many initiatives to broaden communication and

interaction across both customers and other technology development organizations. Some of these activities include the

National Space and Missile Materials Symposium, the

IHPRPT Materials Working Group, and the TPS Working

Group. One of the common benefits of these activities is the

opportunity to validate requirements and provide organizations

with insight to formulate recommended strategies for materials

development programs at large. There are also many other

activities and organizations with whom the Space IAA and its

extended team are engaged; such as the National Aerospace

Initiative, AFSPC (Air Force Space Command) reviews, the

AFRL Space Sector Office, sister directorates, Warfighter

Technology Areas, etc. Additionally, we have ML collocates at

Kirtland AFB (AFRL Space Vehicles Directorate) and Los

Angeles AFB (Space and Missile Center) who continuously

maintain a finger on the pulse of requirements.

As more nations begin developing space systems, rapid

access to space becomes essential for our Nation’s security.

Can you project a single enabling technology that must be

developed if the Air Force is ever to develop a single-stage-toorbit (SSTO) system that can maintain aircraft-like operations?

With respect to materials impact, highly robust, operable 

thermal protection systems are the most enabling. Propulsion

materials closely follow this, where we currently have a reasonably robust materials program. The development of highly

durable materials and processes that will provide options for the

advancement of the various vehicle TPS and hot structures is

equally critical. The TPS must be capable of protecting the

vehicle from not only high heat loads, but also other compromising environments associated with flight and ground operations. The TPS technology is applicable to both advanced

access to space vehicles and high mach/hypersonic air breathing

vehicles. Additionally, the technology has potential payoff to

both rocket and air breathing engine components, such as nozzles. The temperature capability, service life, and turnaround

time are the most critical for the TPS overall. The durability

and damage tolerance of the TPS system will govern the ability of advanced reusable space vehicles to perform with aircraftlike operations.

Over the past several decades, ML has been responsible for

developing many of the materials technologies that have

enabled advanced space systems. A few examples that 

come to mind include carbon-carbon and polymeric matrix

composites for high temperature and structural applications,

and mercury cadmium telluride materials for infrared sensors.

What technologies are you working on today that may 

provide similar advances in capabilities?

Much of the cutting edge work in photonic materials may provide space capabilities that do not exist today. For example, one

application of photonics would provide for multi-gigahertz

modulators; which due to their speed, weight, and size, could

enable a space-based radar system. Another application might be
The AMPTIAC Quarterly, Volume 8, Number 1 7

for non-volatile memory. Our work in biomimetics and nanomaterials could potentially result in lighter, smaller, flat panel

sensors or room temperature infrared detectors to name a few.

Maturing and transitioning technologies is a process that can

sometimes take decades to accomplish. What actions does

your organization take to accelerate the process so that your

new technologies become recognized as viable approaches

by industry?

Here are two examples: one for near or current technology, and

the other a paradigm change in the insertion process of new

materials technologies:

The Space Systems Support and Affordability Effort (S3AE)

is a unique program that we initiated approximately six years

ago to address materials and manufacturing issues of spacecraft,

and provides an avenue for rapid transition. S3AE is a partnership between ML and major spacecraft prime contractors to

address and resolve common issues in materials, processes and

manufacturing. The project duration has a near-term focus, 12

- 18 months, to resolve key shop floor issues common across the

industry. A major product is documentation on issue resolution

and best practices available to the entire spacecraft community

(non-proprietary). It provides a means for near real-time transition to spacecraft production lines. The program also provides a

key connectivity between AFRL and spacecraft primes. 

For new materials technology, we are working on a program

with DARPA called Accelerated Insertion of Materials (AIM).

The time required to procure new systems is often linked to the

ability to make new materials and processes ready for implementation. Consequently, progress in the aerospace sector will

to a large extent be driven by the development and insertion of

new materials and more efficient manufacturing processes. The

AIM program is establishing and validating new approaches for

materials development that will accelerate the insertion of 

these materials into the production hardware for aerospace

applications. The AIM program is showing that the time scale

between development of a new material and its implementation into production can be significantly shortened. The 

resulting pay-off is a substantial reduction in the time required

for the development of cost-competitive aerospace systems with

higher performance and greater efficiency.

In these days of limited budgets, it seems as if collaboration

and joint programs are possibly the only way that technologies can be developed quickly. We know that the Integrated

High Payoff Rocket Propulsion Technology (IHPRPT) Materials

Working Group and the Thermal Protection Systems Working

Groups are ongoing DOD/NASA/industry initiatives to

increase the rate at which new thermal protection and 

propulsion materials are developed. Do you see this

approach being used in other areas?

Collaborative programs and joint organization sponsorship

have been in effect for many years in ML, including examples

such as the Integrated High Performance Turbine Engine

Technology (IHPTET) program and the Composites

Affordability Initiative, both of which include several government agencies, industry OEMs and materials suppliers. Many

other programs are ongoing and ML develops and partners on

new ones continually. Organizing and providing leadership for

collaboration is a routine approach to doing business for us as

we look to leverage resources, develop and transition the

enabling materials technologies, and being able to afford these

costly new developments in a world of ever limited resources.

Integrated Vehicle Health Monitoring (IVHM) might be a

new candidate area. There are many aspects to IVHM that cut

across numerous disciplines and must be tied together to form

a working capability. Currently there is work ongoing within

AFRL, NASA, the Army, Navy, and other agencies. Pulling

together a working group to share common requirements and

solutions is an initiative worth serious consideration. AFRL is

currently addressing many of the constituent technology areas

that would comprise IVHM. These include degradation

processes of materials, vehicle damage caused by the assembly

process, new and novel sensing technology development,

miniaturized control and power-scavenging, modeling of NDE

response, and dynamical systems engineering optimization.

ML is also engaged in identifying critical failure modes that

would be used to define material thresholds in TPS, beyond

which the loss of the vehicle would be considered imminent.

Any thrust in ML would be focused on structures, TPS, wiring,

and propulsion in that order.

The Long Duration Exposure Facility (LDEF) launched by the

Space Shuttle in April 1984 was initially planned to be in

space for a year, but because of the loss of Challenger it

remained aloft for nearly six years. When it was recovered,

the results from exposure of the test specimens to the space

environment were mixed. Some of the tests came out better

than expected since the materials received extended exposures. However, other materials, such as the polymer thin film

specimens, were severely eroded or even destroyed, which

limited the amount of useful data that were gathered. Today

we see a similar prospect occurring. In August of 2001, the

Materials International Space Station Experiment (MISSE)

was attached to the International Space Station (ISS), where

it was scheduled to return to Earth last year. Because of the

grounding of the Shuttle fleet, it will be spring 2005 at the

earliest before MISSE can be returned. Are you worried that

the some of the materials currently being tested on MISSE

may see a similar fate as those on LDEF?

We are somewhat concerned, mainly because atomic oxygen

(AO) erosion is a very severe problem for some materials.

However, the fluence of AO which the specimens on the ram

side of LDEF experienced was significantly greater than that

anticipated on the MISSE experiment. LDEF’s orbital configuration was designed to maximize the AO exposure for the

ram-facing specimens, while totally shielding those on the wake

side of the satellite. Because MISSE is attached to the ISS, we

were not afforded that same luxury, thus none of the MISSE

specimens are either constantly exposed to or constantly shielded from AO. Additionally, LDEF was on-orbit for 5.7 years

and its altitude had decayed from 275 to 175 nautical miles,

and was about a month away from reentry when it was recovered. Thus, most of LDEF’s AO exposure actually occurred in

its last 6-12 months of orbiting the earth. The ISS is in a stable

orbit at about 217 nautical miles and if things go as planned,
8 The AMPTIAC Quarterly, Volume 8, Number 1

MISSE will have been in orbit for about 3.5 years. Thus, the

possibility of totally destroying specimens is not as likely to be

a concern when compared to LDEF. 

Two other thoughts concerning this question: First, MISSE

was originally planned to be a two part experiment with one set

of PECs (passive experimental carriers) being on-orbit for 1 year

while a second identical set were to be on-orbit for 3 years which

is about what the current PECs are anticipated to receive. Thus,

we may just be able to reverse the experiment plan without as

significant of an impact to the specimens as might be expected.

Second, a lot of work has been done to improve the AO resistance of various polymeric materials since LDEF. While MISSE

still is utilizing polyimide materials as AO reference materials

and monitors, the newly engineered polymers on MISSE should

not be as impacted by the AO environment as those on LDEF.

Over the past ten years or so, ML has increased its efforts 

in researching materials for space applications to the levels

seen today. Barring any unforeseen circumstances, do you

envision that this focus will remain for the foreseeable future?

Most likely. AFRL as a whole plans to increase its percent

investment in space technology to the 25-30% level by 2009.

Given the importance of space capabilities in the last two conflicts, a sustained or increased investment in space is justified.

Since many of the engineers and scientists involved in 

the early days of space research have already retired, do 

you find it much of a challenge to obtain the experienced

technical support needed to foster the technology areas 

the lab is pursuing? 

The availability of experienced technical support as well as new

scientists and engineers to train to be experts in space materials

and other disciplines is a national issue. Senior leadership in the

DOD, other government agencies such as NSF and NASA, as

well as those in academia have identified the need for promoting technical careers to entice student to pursue courses of

study to enter those fields. Materials science and engineering

are fundamental across the space sector so that we share these

needs for bright young people. We are attracting them and in

many cases are growing our own by training and providing

exciting opportunities in space materials research. 

Considering that the space environment is so different 

than that facing aircraft, do you find it difficult to train the

engineers that transition into space research? 

From a technology perspective, space can be broken into three

major sectors, spacecraft, space access vehicles, and missiles.

Each sector has its own set of customers, processes, and 

requirements. Additionally, each sector is at a different level of

maturity. Currently, major platforms of each of these sectors are

either one-time use or not recoverable. These aspects inherently present a major challenge for anyone working in the space

sector. The air area provides somewhat of a more comfortable

environment for engineers, as customers and requirements are

more straightforward. Conversely, the space sector is relatively

in its infancy and has very diverse customers and requirements

spread across the three major sectors mentioned above.

ML’s Space IAA Office leads or offers many opportunities for

educating ML personnel in space and missile systems. These

include classes, symposiums, workshops, and meetings. It

developed and offered several classes on spacecraft systems

taught by renowned experts. The IAA Office leads the National

Space and Missile Materials Symposium that is held every June.

Part of this forum includes tutorials and workshops, including

presentations by Nobel laureates. Tutorials include subjects

such as nanotechnology, spacecraft design, smart materials,

space environments, etc. Workshops on propulsion materials

and thermal protection are also held annually. The IAA Office

also holds a bi-weekly meeting at the Directorate to keep personnel abreast of events in the space arena. As an additional

part of that meeting, cameo presentations are offered on 

various materials technologies and activities.

Recently, President Bush proposed extending manned space

flight further into our solar system, including returning to 

the moon and eventually going to Mars. During the early

days of manned space flight, ML was quite involved in space

technology development. A notable example of Air Force

technology that transitioned to manned space flight was 

the heat shield technology used to keep the capsules from

burning up during reentry. Do you see a role for ML in 

developing technologies to enable systems to meet the 

president’s challenging goals?

ML has many technologies that would be applicable to the

President’s new space vision. For example, all of our materials

work for spacecraft would be applicable, as we discussed 

earlier. Depending on the concepts that are ultimately identified, many of the ML programs for space access vehicles could

also be enabling, particularly in propulsion and thermal 

protection. Additionally, our cutting edge work in biomimetics

and nanomaterials for transformational capabilities would 

definitely support the new space vision. What has been true

historically is still true: materials are the principal enablers of

emerging technologies.

Dr. Charles E. Browning, a member of the Senior Executive Service, is Director, Materials and Manufacturing

Directorate, Air Force Research Laboratory, Air Force Materiel Command, Wright-Patterson Air Force Base, OH.

Dr. Browning is responsible for the planning and execution of the Air Force’s advanced materials, processes, and

manufacturing and environmental technology programs to support all elements of Air Force acquisition and sustainment. He is also responsible for interfacing these specific areas throughout the corporate Air Force and

Department of Defense. He leads an organization of approximately 530 government employees with a yearly

budget of nearly $250 million.

Dr. Browning began his career with the Air Force in 1966 and has held various senior technical and 

management positions within the laboratories. He was appointed to the SES in 1998.
The AMPTIAC Quarterly, Volume 8, Number 1 9

Materials for Spacecraft Propulsion Systems

• High durability catalyst materials, such as advanced nanocomposites and hexaluminates with various

coatings, IR Platinum etc. 

• Oxidation resistant chambers - Rhenium/Iridium alloys and processing; cost effective 

manufacture of combustion chambers, and powder metallurgy.

• Hall Thrusters – Alumina, Aluminum Nitride, and Silicon Nitride used to make erosion resistant 

chambers with rapid prototyping operations.

• Electrostatic carbon-carbon ion engine grids. 

Materials for Electric Power 

• Currently conducting industrial base analysis of solar cell manufacturing.

• Polymer wiring – Polymer embedded with metal (copper and nickel coating). Currently 

with application to aircraft but potential for space.

Sensor Materials 

• High performance mercury cadmium telluride (MCT) materials for long wave infrared detection. 

• Early efforts to develop materials growth and processes for quantum dots and carbon 

nanotubes for future hyperspectral systems. 

• Wide band-gap semiconductors; Gallium nitride (GaN) for SSPAs (Solid State Power Amplifiers) 

and LNA (Low Noise Amplifiers) for radar and communications. 

• Developed electro-optic polymer modulators enabling high bandwidth data links and straightforward

solutions to data fusion.

• Developed laser source materials for space-based Light Detection and Ranging (LIDAR).

Spacecraft Structures and Assemblies

• Developing advanced lubrication formulations for application to control momentum gyros (CMGs) that

will result in their meeting 15-year satellite life and for application to moving mechanical assemblies.

• Developed high speed bearing races and pennzane formulation to provide longer lifetimes for future

satellites. TiCN coated races outperformed the state of the art.

• Developing lubrication techniques for micromechanical systems (MEMS).

• Applying nanocomposite research to the development of multifunctional materials for improved 

electrical/thermal conductivity, reduced CTE, and microcracking. 

• Developing advanced polymer materials for application to space-deployable structures and membranes.

Researching nano-processing techniques for ultra-light structures. 

• Collaborating with other DOD agencies, NRO, NASA, and industry for the development of advanced

mirror and optics technology. The thrust of the program is to develop structural substrates, optical 

surfacing techniques, and processing tailorability. 

• Manufacturing: A partnership (S3AE) was developed between ML and major spacecraft prime 

contractors to address and resolve common issues in materials, processes, and manufacturing. The 

project duration has a near-term focus, 12 to 18 months, to resolve key shop floor issues common 

across the industry. A major product is documentation on issue resolution and best practices available 

to the entire spacecraft community (non-proprietary). The program provides a key connectivity 

between AFRL and spacecraft primes. Embedded heat pipes are an example of recent work.

Hardening 

• AFRL/ML is developing a number of technologies that provide filtering and limiting of energy 

from laser irradiation on optical systems.

Materials for Communications, Power Control, and Microwave 

• Developed electro-optic polymer modulators for high-bandwidth optical communications. Working 

for even higher performance modulators for the photonic control of phased array radar for weight 

and power savings.

• Developing improved wide band-gap semiconductor materials (silicon carbide), which will enable 

compact megawatt power switching devices for on-board power management and distribution. 

Thermal Management Materials for the removal of heat from on-board components 

and structures

• Developing advanced coatings; conducting space environment testing.

• Developing materials for thermal planes and radiators.

• Developed C-C radiator material and transitioned to Earth Orbiter 1 (EO-1) satellite, now in orbit.

(Completed)

• Developing carbon-carbon thermal planes and heat sinks for heat removal in advanced electronics.

• Developing carbon-carbon heat pipes for radiators.

• Developing higher conductivity carbon foams for advanced high performance heat exchangers 

and radiators. 

In The 

Spacecraft 

Area

While many of the leading

edge material and process

technologies that the

Materials and

Manufacturing Directorate

are currently developing for

space are discussed in this

issue of the AMPTIAC

Quarterly, the spectrum of

covered topics is far short of

exhaustive. For the sake of

completeness, we offer this

comprehensive directory 

of AFRL/ML’s material

activities related to space

technology.
10 The AMPTIAC Quarterly, Volume 8, Number 1

Thermal Protection System Materials for protection of vehicle structures from high thermal loads.

• Hybrid & Cooled Leading Edges: A hybrid solution will involve the right combination of materials for

management of thermal and structural loads. The hybrid concept being developed by ML is utilizing a

structurally integrated non-parasitic approach with options for a novel combination of materials such 

as C-C, ceramics, metals, foams, aerogels, and/or phase change materials. The focus is on reliability,

durability, and supportability, as well as cost and manufacturability. Another approach is heat pipe 

cooling, which is accomplished by converting the leading edge of an airframe component and a specified

chordwise distance to the rear into a liquid metal heat pipe. This program’s focus is also on reliability,

durability, and supportability, cost, and manufacturability.

• Highly-Operable Ceramic Acreage Panels: ML has two cooperative programs with the Air Vehicles

Directorate (AFRL/VA) to develop and demonstrate large C/SiC standoff thermal protection system

panels. This approach provides for simplified removal that facilitates maintainability and operability of

the vehicle system. 

• Carbon-Carbon Nosetips & Aeroshells: ML is continuing advancement in C-C nosetips and aeroshells.

The current focus is on materials for a modified C-C aeroshell for reentry, which is thermally efficient,

structural, and low cost. The effort also has potential for transition to leading edges and heatshields.

Current materials being investigated are carbon-carbon infiltrated with RTV resin.

• Gamma Titanium Aluminide for Acreage Panels: A metallic approach for acreage TPS is being worked 

by ML in collaboration with AFRL/VA and NASA. This work involves the processing and testing 

of gamma-titanium aluminide sheets for application to large, windward surfaces of a reusable launch

vehicle.

• High Temperature Ceramics for Control Surfaces: ML has planning in place for process and composition

development for fiber reinforced high temperature ceramics. It will include sub-element fabrication and

test of monolithic and composite ultra-high temperature ceramics for durability assessment. Oxidation

protection methods will also be addressed. Materials will be C/SiC.

Materials for Liquid Rocket Engines for the Integrated High Payoff Rocket Propulsion Technology

(IHPRPT) program:

• Turbopump Housing: Nickel based super alloys for LOX. Mondalloy improved Ni based Ox 

compatibility over Haynes 214. 

• Rotating Turbomachinery: Powder metallurgy super alloys. HIP bonded surface layering. Carbon steel

tooling used to clad with HIP process and Ox resistant materials.

• Lines, ducts, and valves: Nanophase aluminum, LCP lined, graphite epoxy ducts, and PMCs for liquid

hydrogen feedlines.

• Nozzles: Radiation and actively cooled approaches using CMCs. SiC/SiC, C/SiC coated with various

materials. Open woven Ceramic for a dual bell. 

Materials for Cryotanks, both organic matrix composites and metallic

• For organic matrix composites current efforts are focused on reducing microcracking, demonstrating 

liquid oxygen compatible OMCs, and assessing modeling/test methods.

• For metallic tanks, collaborating with NASA on the scale-up of an Air Force formulation (2099) 

of Aluminum-Lithium. Provides lighter weight and improved characteristics over current metallic 

formulations.

Vehicle Health Monitoring

• ML is engaged with AFRL in a teaming approach, to address issues associated with operable spacelift

vehicles. ML is also engaged in identifying critical failure modes that define material thresholds in 

TPS with loss of vehicle consequences. ML is also part of AF oversight team on a DARPA Integrated

health monitoring (IVHM) program for propulsion.

Materials for Airframe Structure and Subsystems 

• ML is researching high temperature PMCs that are suitable for airframe applications capable of 

continued service temperatures beyond 450°F. Potential payoff is the reduction in the amount 

of TPS required on a vehicle. 

• Thermal management – Applying higher conductivity carbon foams for advanced high performance

heat exchangers and carbon-carbon thermal planes and heat sinks for advanced electronics.

In The 

Spacelift 

Area

Similarly, the full panorama

of AFRL/ML’s material

development efforts for

launch and re-entry vehicles

far exceeds the scope (or

capacity) of this issue of 

the Quarterly. Here is the

complete array of activities.
The AMPTIAC Quarterly, Volume 8, Number 1 11

Manufacturing Technology for Filament Winding Large 

Solid Rocket Motor Cases: 1960-65

Early Minuteman motor cases were constructed of steel and

titanium. Research was aimed at replacing the Minuteman II

(MM II) second stage metal motor cases with lighter-weight

fiberglass filament-wound cases. Although the MM II system

program office (SPO) did not adopt this technology, the work

provided experience and confidence; which eventually led to

the introduction of a composite third stage motor in the

Minuteman III, and composites in all three stages of the

Peacekeeper.

Fabrication of Large-Diameter Ablative Solid Rocket 

Motor Throats: 1964-70

This research was part of a series of programs to explore alternative designs and materials for very large motors. In the same

timeframe, the Titan III system program had suffered repeated

nozzle failures in its Solid Rocket Motor (SRM) development

efforts. The results of this research and successful motor firings

provided a capability and confidence that allowed the replacement of graphite nozzles in the Titan III SRM system and all

subsequent designs.

Manufacturing Technology for Shear-Forming Large-Diameter,

Metal Solid Rocket Motor (SRM) Cases: Mid 1960-70

Ring-formed and welded metalworking techniques produced

the steel motor cases for the Minuteman and the first block of

Titan III SRMs. Our research resulted in the design, assembly,

and demonstration of a shear-forming machine capable of producing a Titan III-size motor case without welds. The program

produced an expensive but capable, large shear forming

machine for producing case up to 120 inches in diameter and

demonstrated them in both D6AC and 9Ni-4Co steels.

Although it never went into production, it motivated suppliers

to develop shear forming processes that resulted in significant

case cost savings, enhanced performance, and reliability.

Manufacturing Technology for Radiation-Hardened 

Integrated Circuits: Mid 1960’s - Late 1980’s

Early work in this area addressed the introduction of dielectric

isolation into the fabrication sequences for hardened integrated

circuit (IC) manufacture. As radiation effects became better

understood, and circuits became more complex; circuit

redesigns required the introduction of new and or modified circuit fabrication processes, which became the focus of several

Manufacturing Technology (MT) contracts. Hardened memories, logic circuits, and mass storage were all subjects of MT

contracts. Prior to the involvement of the Phillips Laboratory

(now AFRL/VS) in hardened circuit design and fabrication; the

MT program had produced the first hardened chip set that was

MilStd compliant.

Infrared Detector Development: Late 1960’s - Present

At the time of their early system application, certain important

infrared (IR) responsive materials were known empirically, but

not scientifically. The processes used to produce the materials

involved a good deal of art. To help mature IR material technology, the mechanisms of photoconductivity in lead sulfide

(PbS) were systematically studied by the materials laboratory

(ML) to build a stable process that reduced manufacturing

problems. However, ML believed that the ultimate answer lay

in an alternate IR sensor material, namely mercury-cadmiumtelluride (HgCdTe). A major accomplishment was the singlehanded development of HgCdTe IR detectors for short wave

Paul M. Propp, Private Consultant

Retired, Materials and Manufacturing Directorate

Air Force Research Laboratory

The Air Force Research Laboratory has been a leader in space technology since the beginning of the space race. Over the past four decades,

many technical professionals have dedicated substantial portions of their careers to bring some of these technologies to realization. During

this time span, many contributed to some number of these milestones, but few were fortunate enough to bear witness to the entire history

of AFRL’s space efforts. During the preparation of this issue, AMPTIAC was privileged to speak with Mr. Paul Propp, who is one of the

few people to have gained that unique perspective. Mr. Propp worked at AFRL from the early 1960s through the mid-1990s. During the

bulk of his tenure, he was the materials liaison for the Space Systems Division (now the Space & Missile Center), which afforded him a

unique view of developments in both the materials and systems areas. Since his retirement almost 10 years ago, he has continued to serve

AFRL and its technology programs as a private consultant.

Many staples of space technology can trace their roots to groundbreaking work at the Air Force’s laboratories. Bringing some of these more

notable achievements to light, Mr. Propp shares his recollections of the major strides AFRL has made in space technology over the last forty

years. – Editor
12 The AMPTIAC Quarterly, Volume 8, Number 1

infrared (SWIR), medium wave (MW), and long wave (LW) IR

spectrums. ML-directed work resulted in HgCdTe becoming

the detector for the second color DSP focal plane (a Defense

Support Program’s early warning satellite able to simultaneously detect at two different wavelengths) and subsequently replace

PbS as the primary SWIR detector as well.

Advanced Composite Structures for Space: 

Late 1960’s – Present

Following initial work in the use of fiberglass-reinforced polymer materials, ML sought fibers with better mechanical properties. Boron fibers, followed shortly by carbon and graphite

fibers, became the basis for an Advanced Development

Program (ADP). After early advancements directed at aircraft

structures, the ML turned ADP attention to consider space and

missile applications. Aluminum, beryllium, and titanium were

the standard materials of construction at the time. Now, after

many AFRL-programs supporting composite development,

graphite reinforced composite materials have become one of

the industry’s basic materials of construction for space structures, dimensionally stable space structures, and propulsion

tankage (for station keeping). Moreover, they have replaced

fiberglass for filament wound motor cases and other missile

structures in many Air Force systems, including Peacekeeper.

Temperature Compensated Rare Earth Magnets: Early 1970’s

High strength, rare earth magnetic materials were an ML development. Recognizing the potential impact of these materials,

ML was encouraged to seek out applications which would meet

Air Force needs. MT followed the ML development of rare

earth magnets with programs directed toward scaled-up

processes and controls for producing these materials in reproducible and high quality volumes; and subsequently demonstrating them in critical design applications. MT programs

directed at space applications were conducted on magnetic

bearings and traveling wave tubes (TWTs). The TWT work

was immediately successful. Rare earth-focused TWTs are

today the standard in the industry. Magnetic bearings have

found limited space applications, but have been used on DSP

satellites.

3D Carbon Carbon Composites for Reentry Vehicle 

Nosetips: 1970’s

Carbon-carbon (C-C) composites were an ML-conceived and

sponsored development. The ML sought out and established a

strong technology program in support of the Advanced Ballistic

Reentry Systems Program (ABRES). The ABRES program was

responsible for the advanced development of next generation

ballistic missile reentry vehicle technologies. Great emphasis

was placed on improved accuracy and reliability of the warheads reentering under extreme environment conditions. Bulk

graphite (in many varieties) and pyrolitic graphite were the

products of ML and others, but all were found inadequate for

high precision reentry vehicle nosetips. ML focused on a coordinated effort across the Lab and in conjunction with ABRES

to develop C-C composites for this application. This technology was enabling to achieve precision ballistic reentry vehicles

(RVs) for MIRV warheads. This was a totally successful effort

which resulted in the redesign of the MK 12a (Minuteman III)

and subsequent RVs.

High Performance, Lower Cost, Reproducible Carbon Carbon

for Rocket Motor Throat Inserts: Mid 1970’s - Early 80’s

In the afterglow of the success of carbon carbon (C-C) for RV

nosetips, ML turned to rocket propulsion applications. Large

diameter solid rocket nozzles, as in the Titan SRM (note earlier), could accept the lower efficiency of ablative nozzles; while

high performance tactical missiles motors could not. Frequent

failures of graphite and ceramic throats, plus the inefficiency of

ablative nozzles demanded a new approach to motor designs be

SPACE & MISSILE HERITAGE

Traditional and Current

Research Areas

AFRL has a long, proud heritage 

of developing leading-edge space

technology. As shown here, the

Materials and Manufacturing

Directorate (ML) has played a

major role in this legacy.

Structural Materials Thermal Control Materials Laser Hardened Materials

Organic Composites 

Metal Composites 

Carbon-Carbon Composites 

Design Methods & Data 

Joining Technology

Tribomaterials

Missiles/Hypersonics

Space Environment

Power Materials

Industrial Base Issues

Structural Coatings 

Multi-Layer Insulation 

Radiator Coatings 

High Conductivity Substrates

Sensor Materials 

Optical Substrates 

Baffles 

Coatings/Filters 

Antenna Materials 

High Temp Adhesives 

Hardened T/C Coatings 

Hardened Multilayer Insulation Advanced Practices 

Process Optimization 

M&P Transition

Electronic Materials

Detector Materials 

Microwave & Microelectronic Materials 

Packaging 

Nose Tips 

Heat Shields 

Antenna Windows 

Rocket Motors 

Exit Cones

Systems Support

Failure Analysis 

Consultation 

Collocated Engineering

Solid Film Lubricants 

Advanced Liquid Lubricants 

Life Prediction Methods 

Bearings

Spacecraft Coatings 

Space Environmental Effects 

on Materials 

Space Effects Simulation/ 

Modeling 

LDEF Materials Evaluation

Long Life Battery Materials 

High Efficiency Solar 

Cell Materials 

Radiator
The AMPTIAC Quarterly, Volume 8, Number 1 13

taken. The Air Force’s Rocket Propulsion Lab (RPL) had established advanced tactical missile technology as a thrust. ML

advanced a plan to RPL to demonstrate C-C composites for

high performance solid motors. A new basis for designing tactical missiles resulted from a coordinated program involving

ML and RPL. ML led the efforts related to material development and characterizing MT for producibility, while RPL

headed up the motor design and ground firing tests. This work

also helped pave the way for C-C to be introduced into the

Peacekeeper missile motor design and others to follow.

Propellant Compatible Elastomeric Seals and Bladders:

1970’s - Early 80’s

Spacecraft and missiles have a requirement for long-time containment of energetic propellant fluids. Expulsion of these 

fluids in the zero-g environment requires positive pressure to be

imposed. Early approaches employed metal bellows or Teflon

seals and surfaces, but both approaches had problems. ML 

pursued the development of propellant-compatible elastomers

- first for fuels, and later for more aggressive oxidizers. The

development of AFE332, a hydrazine-compatible elastomer

solved that fuel containment need and became an industry

standard used for seals and bladders. The subsequent development of AFE124 provided an elastomer to be used with 

nitrogen tetroxide (NTO). This material found use in Titan II

missile seals.

Contamination and Optical Surface Material Data: 

1970’s - Present

Accurate data for the degradation of spacecraft thermal 

control surfaces and optical surfaces has a major influence on

reliably predicting on-orbit lifetimes and sizing thermal radiators. Degradation of these surfaces results from radiation

effects, and contamination from outgassing products from the

satellite itself. In addition to supporting the development of

space-stable thermal control coating materials, ML has sought

to provide designers with accurate in-space performance

numbers on many of the commonly used thermal control

materials. Space experiments ML 101, ML 12 and very

recently the MISSE experiment, have provided spacecraft

designers with fundamental information that has become

standard references in the industry. ML had also supported

the definition of the first contamination effects model for the

analysis of this phenomenon.

Manufacturing Technology for Nickel Hydrogen 

Space Batteries: 1980’s

The relative cost of nickel hydrogen (Ni/H2) space battery 

cells was cited as an obstacle to their broader usage. This MT

program examined the design and fabrication of pressurized

Ni/H2 cylindrical battery cells with lower cost and enhanced

reproducible performance in mind. At the successful conclusion of this program, the resultant redesign and assembly techniques enabled what was referred to for several years as the 

Air Force standard MT cell. This cell and others were entered

into the national life test program ongoing at the time and

became the initial baseline technology for the transition from

NiCd to Ni/H2 space batteries.

GaAs Solar Cells: Mid 1980-1990’s

GaAs has long been recognized as a candidate solar cell material.

The Aero Propulsion Laboratory (APL, now the Propulsion

Directorate, AFRL/PR) began the development of GaAs for solar

cell application and demonstrated its potential to significantly

exceed the performance of the most advanced silicon cells. In a

series of programs, MT met the demands of performance, quantity, and schedule on this classified program and established the

industrial base for early GaAs solar cell production (for several years,

the industry referred to GaAs solar cells as the “MT cell”). MT 

subsequently supported one or two additional programs dealing

with multijunction cells that have advanced solar cell technology to

the highest level of production cell efficiencies achieved to date.

MT for HgCdTe IR Detector Module Producibility: 

Late 80’s - Early 90’s

The transition of HgCdTe detectors to space systems presented major producibility challenges, especially with the large

production quantities associated with national missile defense.

Low yields, lack of uniformity, and high cost caused major 

system concerns. A sizable MT program was launched with

remarkable results. Yields were improved over tenfold, performance and uniformity was enhanced; and a strong manufacturing base was established.

Surface Analysis for Traveling Wave Tubes: 1980’s

Premature failures of traveling wave tubes (TWTs) have caused

major upsets to Space and Missile Center (SMC) and other

programs. More than one failure mode has been identified for

TWTs. ML was asked to study those modes related to emitter

failures. Working in close association with tube manufactures

and the program office (supported by the Aerospace

Corporation), ML put together a sophisticated in-house team

of surface scientists. Supported with state-of-the-art instrumentation, the team established new analysis techniques, which

resulted in a vastly improved understanding of the surface

physics involved with emission and emitter contamination.

The data and models developed were used to modify tube 

fabrication and material acceptance criteria.

Mr. Paul M. Propp has over 40 years experience in working with materials for space and missile applications. He

was first assigned to the then Air Force Materials Laboratory as a Military officer in 1960, where he served as a

chemist. Mr. Propp rejoined the Air Force in 1964 as a civilian and served as the Materials Directorate’s representative at the Air Force Space and Missile Systems Center until his retirement in 1993. He has participated in

many significant materials advancement throughout the course of his career.

2004 National Space & Missile Materials Symposium

Developing Materials to Transform the Future

21 – 25 June

DoubleTree Hotel Seattle Airport, Seattle, Washington

An ever-increasing reliance on space and missile assets, over the full spectrum of applications, has strengthened the call for reusable launch 

systems and advanced spacecraft and payloads that are both more reliable and more affordable. These needs will only be met by commensurate

advances in space and missile technology. The fundamental sciences of designing materials and the related areas of material processing and 

evaluating material performance in space environments represent the cornerstones of these enabling technologies. 

The National Space & Missile Materials Symposium (NSMMS) can trace its origins back to the 1996 AFRL Materials Directorate Program

Review. In September of that year, over 250 engineers & project managers from the Air Force, NASA, and Industry participated in the Review.

Building on the success of that review and an overwhelming request for expansion by Tri-Service and NASA attendees, Dr. Vincent Russo, former Director of AFRL/ML and current Executive Director, Aeronautical Systems Center, founded the National Tri-Service and NASA-sponsored

Symposium in 1998. With the Air Force continuing to act as lead, the NSMMS provides a national forum on the technical challenges facing

future generations of space and missile systems in materials and processes. 

NSMMS presents information on the state-of-the-art issues in materials and process technology development for space and missiles. It provides

attendees a global perspective for aiding investment planning and contributes to the exchange of technical information and interaction of 

managers, project engineers, and scientists. Additionally, users of materials and advanced technologies impacting the future direction of science

and technology in space related fields gain insight into the thinking of leaders in space materials and processing technology. 

The focus of the 2004 NSMMS is on critical enabling materials, materials processing technologies, and the integral role 

of advanced materials in space, missile and surveillance applications. Distinguished speakers for the plenary sessions 

this year include:

• Mr. Sean O’Keefe, Administrator, National Aeronautics and Space Administration 

• General Lance Lord, Commander, Air Force Space Command 

• Mr. Peter Teets, Director, National Reconnaissance Office

• Dr. Anthony Tether, Director, Defense Advanced Research Projects Agency

• Major General John Urias, Deputy Commander, Research, Development and Acquisition, US Army Space and Missile Defense Command

• Dr. Steven Chu, Nobel Laureate

• Mr. Storey Musgrave, NASA Astronaut (Dinner Speaker)

This year’s conference has four major technical thrusts; each of which is the topic of one of the technical session:

Technical Session 1, New Directions in Materials Science, focuses on the frontier of the next materials science revolution. Papers address needs

for future space and missile applications, discuss new directions in materials science, and identify new opportunities where new technologies can

address those needs. Applicable system applications include spacecraft, reusable launch vehicles, hypersonic systems, missiles, and reentry 

systems. Breakthroughs and developments in this arena will depend on many disciplines, including materials scientists, chemists, biologists, 

computational scientists, and engineers 

Technical Session 2, Access to Space, addresses critical material and process technologies for enabling access to space, focusing on both 

airbreathing and rocket propulsion vehicles. NASA’s goal is affordable, safe space transportation. The Air Force’s goal includes this, as well as 

“aircraft-like operations” for space operations vehicles. 

Technical Session 3, Operating in Space, covers materials used on hardware and systems in various space environments. Emphasis is on 

critical materials for current and planned satellites, spacecraft, and payloads for both commercial and government missions. 

Technical Session 4, Missiles and Missile Defense, addresses material and process technologies for missile defense interceptors, ICBMs, 

SLBMs, and reentry vehicles (ballistic and maneuvering RVs/CAVs). The focus is on critical enabling materials and materials processing 

technologies related to detection, tracking, and engagement of enemy ballistic missiles, as well as missile materials issues relative to delivery and

lethality of offensive weapons.

For more information see http://usasymposium.com/nsmms/default.htm or contact:

Steering Committee Chair: Eric Becker eric.becker@wpafb.af.mil

Technical Information: Daniel Cleyrat daniel.cleyrat@wpafb.af.mil

Symposium Coordinators: Michelle Kubal mkubal@anteon.com

Patrick Sisson psisson@anteon.com

Telephone: 937-254-7950 ext. 1168 • Fax: 937-254-2153 

Postal address: 

National Space & Missile Materials Symposium 

C/O: Anteon Corporation 

5100 Springfield Pike, Suite 509 

Dayton, OH 45431 
NAMIS Focus Areas:

Conferences And Workshops

This area contains modules on conferences and workshops sponsored by 

organizations such as the DOD, DOE, and NASA. Topic areas include high 

temperature polymers, aircraft structures, high temperature composites, and 

electromagnetic window materials.

Material Property Database

This area contains specialized databases. The Infrared Materials Database is 

currently available.

Virtual Libraries

This area features a variety of informational libraries containing electronic 

technical reports addressing research results and lessons learned. Information is

currently offered on thermal protection systems, cryogenic tanks, advanced 

turbine engine development, and the National Aerospace Plane (NASP) initiative. 

http://namis.alionscience.com
The AMPTIAC Quarterly, Volume 8, Number 1 17

INTRODUCTION

America’s legacy of space technology is nearly a half a century

old, yet there have been a multitude of achievements in this

short time. Even with the many successes, much of the technology is still early in the development phase. The United

States continues to lead the development of space technology,

but has seen the emergence of competitors, as several other

countries are working to close this gap.

By far, NASA is the most visible representative of America’s

space program. Many Americans incorrectly perceive that our

efforts to explore space and develop space technology are exclusively the purview of NASA. The US Military has been a leader

in developing space technology since the late 1940’s, a full

decade before NASA’s establishment. In the past several

decades, the commercial presence in space has become a critical element of our nation’s communication infrastructure. It

has grown significantly as well.

The Military, and more specifically the Air Force, has quite

an extensive history in developing and utilizing space technology. In fact, a continued presence in space is critical for the

Military to remain strong and capable of protecting our 

country from all enemies. The military’s presence in space has

recently taken on new importance by providing our fighting

forces with information needed to defeat the continually adapting landscape of terrorist organizations. Therefore, sustained

development of space technology is critical for the Military to

evolve and defend against all threats to our national security.

The following article provides an overview of space vehicle

basics, and provides readers with some of the factors that go

into designing spacecraft.

SPACE MISSIONS

A spacecraft’s design depends entirely on the vehicle’s intended

function. Most spacecraft launched into Earth’s orbit fulfill one

of four primary space missions:

• Weather

• Navigation

• Communications

• Intelligence/Surveillance/Reconnaissance (ISR)

The Military has interests in each of these four mission 

categories. For example, weather and navigation satellites

(Figure 1) can provide valuable information for military units

serving on the ground in foreign territories, and even more 

so for those serving at sea. Communications are an essential

component of coordinating military operations, and thus 

communication satellites are similarly critical. Finally, the vital

importance of ISR satellites for Military purposes requires no

substantiation.

While cost is a strongly prohibitive factor which tapers the

frequency of spacecraft launches, it is not the sole obstacle.

Scheduling, regulations, politics (domestic and international)

and environmental considerations also tend to be contributing

Benjamin D. Craig

Richard A. Lane

AMPTIAC

Rome, NY

If necessity is the mother of invention, then it is equally true that materials are the enablers of technology. Without the constant drive to

develop new and better materials; possessing properties and capabilities above and beyond their predecessors, technological innovation

would grind to a halt. No place is this more true than in the highly demanding (and unforgiving) realm of outer space. Over the past 

several decades, groups of dedicated scientists, engineers, physicists, chemists, astronomers, and many others from all sectors – military, 

civilian government, industry, and academia – have all labored to make space more accessible.

While most Americans, technical and layperson alike, have a general appreciation of space technology’s benefits to humanity and are

aware of many of its historical highpoints, the total number of people who actually make the technology work is an elite few. The truth is,

there are only a handful of technical professionals outside of the space technology field who have first-hand knowledge of, or are well-read

when it comes to what demands are placed on space vehicles. Satellites, probes, and launch vehicles share very few common design and 

performance issues with vehicles and systems operating on land, sea, or air. The articles in this issue of the AMPTIAC Quarterly explore

many of the technical aspects and challenges of accessing, surviving, and succeeding in space. While preparing this issue, it became evident

to us that many of our readers might find it beneficial to do a little ‘catch-up’ reading before diving into the specific technical issues that

make operating in space the challenge it is.

Two articles in this issue address this point. This first article is a ‘Space 101’ piece, which will introduce the reader to the basic concepts

of space and space technology. The second article, our MaterialEASE feature (the next article in this issue), discusses how materials and 

subsystems are tested and evaluated for use in the space environment. So, for the space neophytes among you, we humbly submit that this

might be the right place to start. We trust that you’ll find it helpful and enlightening. Enjoy! - Editor

The Space Environment and Materials Testing Section II
factors. In the late 1960’s, the majority of the 

satellites launched into orbit were used for military

purposes (excluding communications), but since

then, communications satellite launches are being

launched with increasing frequency.

In the past, most launches to send satellites into

orbit were dominated by the US Military and

NASA. In recent years, the commercial sector has

been extremely active in putting satellites into orbit

since space has become a critical aspect of the commercial communication, navigation, and weather infrastructure. Most of the commercial space technology is in the form

of orbiting communication satellites.

ORBITS

Part of defining a spacecraft’s mission is to choose an orbital

path suitable for carrying out the mission. In general, the mission type dictates what orbit must be chosen, and the mission

and orbit together determine many aspects of the spacecraft

design. This is because the type of environment a spacecraft

encounters is highly dependent on the orbital path. Moreover,

the orbit (altitude) and mission type (weight) together will

determine the propulsion requirements.

Humans have been launching objects into Earth orbit since

1957. There is a common misconception that there are an

almost infinite number of orbital paths that can be chosen

when launching spacecraft. In practice, there are only a handful of useful orbits, each offering specific benefits to certain 

mission types. Many of these different orbit types have several

variations, the most common of which are discussed later.

Some of the factors that go into selecting an appropriate orbit

for a spacecraft are mission requirements, Earth coverage, space

environments to be encountered, and whether it will function

as a single satellite or as part of a constellation.

Orbits are characterized by several aspects: altitude, eccentricity, inclination, and synchronization. An orbit can be 

considered as a complete geometric loop around a celestial

body (for the purposes of this article we’ll just consider the

Earth as the center for every orbit), where the farthest point on

the loop from the body is the apogee, the nearest point is the

perigee, the elapsed time for one complete circuit is the orbit

period, and the angle between the orbital plane and the equatorial plane is the inclination. The altitude is simply the height

above the Earth, or the distance of the satellite from the Earth’s

surface. The eccentricity is the degree to which the orbital path

is elliptical (i.e. can be anywhere from highly elliptical to circular). The direction of the orbit, in most cases, is either direct or

retrograde. A direct (or prograde) orbit is any type of orbit that

moves in the same direction as the Earth’s rotation (East with

respect to the Earth). A retrograde orbit is any type of orbit that

moves in a direction opposite to that of the Earth’s rotation

(West with respect to the Earth). Figure 2 is an illustration of

these and other concepts.

Types of Orbits

The four primary types of orbits are the Low Earth Orbit

(LEO), the Medium Earth Orbit (MEO), Geosynchronous

Earth Orbit (GEO), and High Earth Orbit (HEO). While

these terms are all commonly used by the international space

community, the current definitions of LEO, MEO, and GEO

lack specificity, and thus are more rules of thumb. General definitions of these types of orbits are displayed in Table 1.

Low Earth Orbit The low Earth orbit is up to 1500 km above

the Earth, and is mostly circular (an eccentricity near zero).

Communication satellites in LEO have the advantage of a significantly reduced signal delay, as compared to satellites in

18 The AMPTIAC Quarterly, Volume 8, Number 1

Figure 1. A US Military

Navigation Satellite – The

Air Force’s Navstar Global

Positioning System

Satellite.[1]

Figure 2. Illustration of Orbital Concepts and Specific Types of

Orbits (not to scale).

Table 1. Summary of Useful Orbit Types.[3]

Orbit Type Orbital Altitude (km) Orbital Period Space Environment Hazards

Sun Synchronous 800 ~ 100 min Trapped Particles, Cosmic Radiation, Solar Events

Semi Synchronous MEO 27,000 12 hrs Trapped Particles, Cosmic Radiation, Solar Events

Molniya 40,000 (apogee) 12 hrs Trapped Particles, Cosmic Radiation, Solar Events 800 (perigee)

GEO 35,790 km 24 hrs Cosmic Radiation, Solar Events

The Earth’s Axis of

Rotation is Inclined off

Its Orbital Plane

Orbital Period – One

Complete Loop

Earth’s

Rotation

Molniya

Orbit

Apogee

Elliptical

(Eccentric)

Orbit

Retrograde

Orbit

Angle of

Inclination

Equatorial Plane

Polar Orbit

S Inclination=90°

N

Perigee

~63°

~23°

Circular

Orbit

Prograde

Orbit

↔
The AMPTIAC Quarterly, Volume 8, Number 1 19

GEO (as they are much closer to the Earth). Moreover, satellites in LEO can cover every area of the Earth’s surface,

although it may take several orbits. Surveillance satellites in

LEO require less power for sensing since the range is closer.

However, the coverage area is smaller because of this reduced

range. Therefore, more satellites are required in a constellation

to achieve global coverage. For this reason, reconnaissance and

observational satellites are placed in LEO. Satellites in LEO,

however, may experience a harsher radiation environment than

those in GEO.

Medium Earth Orbit The medium or mid-Earth orbit is a

vaguely defined orbit, but is generally considered to be between

the LEO and GEO at approximately 1,500 – 20,000 km.

There is a moderate signal delay for ground antennas receiving

signals from satellites in MEO, but not as significant as those

from GEO. MEO satellites are a balance between power

requirements, coverage, and resolution.

Semi-Synchronous Orbit A special MEO orbit, a semi synchronous orbit has a period of 12 hours, such that it covers an identical ground track every 24 hours. This orbit is uniquely suited

to some communication and navigation missions. Global

Positioning System (GPS) satellites use this special orbit located at approximately 20,200 km.

High Altitude Earth Orbit The high altitude Earth orbit (HEO)

is an orbit that provides the largest coverage with the fewest

satellites. HEO satellites have higher power requirements than

LEO and MEO satellites.

Geosynchronous Orbit Geosynchronous orbits are the most

common HEO orbit, where objects in the geosynchronous

orbit move through space at the same angular rate as the Earth’s

rotation. That is, satellites in GEO have an orbital period of

approximately 23 hours, 56 minutes and 4 seconds. A more

specific type of GEO is the geostationary orbit, which has an

inclination very close to zero and is a circular orbit.

(Geosynchronous and geostationary are often used interchangeably despite the technical difference.) This results in the

orbiting object constantly looking at the same part of the Earth

at all times, and can see up to 1/3 of the Earth’s surface at a

time. The altitude for GEO is approximately 35,790 km

(22,240 miles). Most

GEO satellites are in a

geostationary orbit (which

is on the equatorial

plane). This orbit is used

by communications satellites to transmit signals to

antennas that are at a

fixed location on the

Earth’s surface. At this

transmission distance, the

signal delay is a quarter of

a second for a round trip.

Satellite TV communications satellites and some

weather satellites are typically located in a geosynchronous

orbit. Some surveillance satellites are also parked in geosynchronous orbits. When large area coverage of the Earth’s

surface is needed (except polar regions), geostationary orbits are

preferred.

The primary types of orbits are depicted in Figure 3. Some

common variations of these orbits include sun synchronous,

polar, and Molniya orbits.

Molniya The Molniya orbit is a highly elliptically semisynchronous HEO. Its inclination is about 63 degrees (depicted in Figure 2). A satellite in this orbit spends 23 out of 24

hours over the northern hemisphere, and it will repeat its

ground track every 24 hours. Satellites “hover” near their

apogee over the northern hemisphere because of the highly

elliptical orbit. This orbit is ideal for large coverage of the

northern latitudes. Much like the sun-synchronous orbit, the

Molniya orbit uses gravitational disturbances to keep the

apogee (40,000 km) and perigee (800 km) from rotating.

Polar Orbit A polar orbit (see Figure 2) has a very high angle

of inclination, close to 90 degrees, and is typically at an altitude

of 700 – 800 km above the surface of the Earth; thus it is a

LEO. This type of orbit is useful for being able to see the entire

surface of Earth at a relatively low altitude. When a satellite is

launched to attain a polar orbit, it requires more energy (rocket fuel) for it to achieve its destination orbit, as compared to

satellites sent to an orbit that rotates with the Earth. This is

because the latter uses some inertial velocity provided by the

Earth’s rotation to achieve its final orbital speed, whereas the

former has no assistance from the rotation of the Earth and

thus depends completely on rocket fuel to achieve the final

orbital speed.

Sun Synchronous Orbit The Earth revolves around the Sun

about every 365 days. If the orbital plane of a satellite in a

slightly retrograde, ‘near-polar’ orbit is rotated about 1 degree

per day, this satellite will pass over a point on the Earth’s surface at the same time every day. Since the Earth is not an exact

sphere, satellites get an extra gravitational ‘pull’ as they cross the

equator. This pull is a result of the Earth’s equatorial bulge. A

sun-synchronous orbit uses this extra pull to shift, or ‘twist’ the

orbital plane the needed fraction of a degree to maintain a fixed

Figure 3. Illustration of the Primary Types of Orbits.

Low Earth Orbit

Medium Earth Orbit

Geosynchronous Orbit

S

N
orientation to the Sun. The amount of pull corresponds to the

altitude and inclination. The specific sun-synchronous orbit

resulting in the correct rate of planar rotation occurs at an inclination of 98 degrees (8 degrees off-polar in a retrograde orbit)

and an altitude range of 800 to 1000 km. This orbit is extremely useful for weather satellites, surface mapping, navigation,

and especially surveillance. Surveillance satellites can pass over

a particular location on the Earth during the most optimum

viewing conditions. Figure 4 shows the concept of a sunsynchronous orbit.

SPACECRAFT DESIGN

The design of spacecraft can be divided into space access

(launch, and in some cases, reentry vehicles), and orbital satellites. The following sections describe the various components of

the spacecraft along with their functions.

Launch Vehicles

Launch vehicles, excluding intercontinental ballistic missiles

(ICBMs), have essentially two purposes: first, to protect the

spacecraft from the severe launch and ascent environments, and

secondly, to accelerate a spacecraft (or other mission payload)

into orbit. There are two fundamental types: Expendable

Launch Vehicles (ELVs) and Reusable Launch Vehicles (RLVs).

ELVs, as their name suggests, are used only once to get a payload into space, and are thereafter allowed to fall back to Earth,

to either splash into the ocean, or burn up upon reentering the

atmosphere (this usually happens to some of the smaller components). Very few launch vehicle components are left orbiting

the Earth. The number of orbiting discarded components,

known as space junk, is growing, which poses a hazard to systems in LEO and MEO. In contrast, after RLVs return to Earth

from orbit, their subsystems are either replaced or refurbished

to prepare the vehicle for its next launch. An additional consideration for both ELVs and RLVs is whether the flight will be

manned. Manned flights are constrained to much smaller

launch acceleration loads (about 3g’s versus 6+).

ELVs are used far more often than RLVs because they are

lighter, require less rocket fuel, and consequently, are less

expensive. There are also a number of different ELV designs in

use, all having different mission capabilities. This variety

allows a launch vehicle to be chosen which best suits the payload and mission. The premise of RLVs and their potential

long-term payoff remain very attractive to some for obvious

reasons (e.g. the entire vehicle does not have to be replaced

after every launch).

RLVs encounter all the same conditions that ELVs do during

launch, but additionally they face the arduous task of reentering the Earth’s atmosphere and landing. As a result, they

must be designed to withstand these very severe conditions.

This added capability requires additional systems, which in

turn adds more mass to the vehicle. It has therefore proven 

to be a difficult challenge to design and build a reusable vehicle

that is safe and reliable. To date, the shuttle is the only partially reusable vehicle.

RLVs encounter high wind loads during launch and descent,

and also experience extreme aerodynamic heating during 

re-entry, which requires them to employ thermal protection

systems. Furthermore, the external surface of the RLVs must

have an inherent resistance to erosion and other damage from

sand and dust particles, as well as rain, and possibly even hail;

all of which may be encountered during re-entry. RLVs also

must endure strenuous mechanical loads during landing.

Consequently, the landing systems add significant weight to the

launch vehicle since they are made of steel (they must be able

to support the entire weight of the vehicle upon landing. Other

necessary systems, such as the brakes and parachute, also add a

significant amount of weight.

There are numerous systems that are typically incorporated

into launches. These include wings (on some vehicles), engines,

oxidizer tanks, fuel tanks, intertanks, fuel tank structure, fuel

tank TPS/insulation, fuel tank propulsion systems, fuel tank

electrical systems, and avionics. Some RLVs have wings for 

stabilization while re-entering the atmosphere, although the

wings serve no purpose while in space. This is why RLVs must

have additional systems, such as thrusters, that can provide 

stabilization in space.

The design considerations for launch vehicles differ greatly

from those of aircraft due to the vastly different environmental conditions they experience. Launch vehicles must typically

20 The AMPTIAC Quarterly, Volume 8, Number 1

Figure 4. Depiction of 

a Sun Synchronous

Orbit.[2] 

Photo Courtesy of NASA Dryden Flight Research Center

Figure 5. Artistic Concept of the X-33 on the Launch Pad [7].

30 Minutes

Darkness

Equator

72 Minutes

Sunlight

Orbit Plane

Rotates Eastward

~1° Per Day

25.59° Earth Rotation

Per Orbit

Orbit Path

N

S
endure extreme mechanical and thermal

environments, in addition to the severe

environment of space. For instance, they

experience mechanical stresses from such

conditions as liftoff and ascent acceleration loads, shock loads, vibrational

loads, and aerodynamic loads (wind).

Other stresses and loads include those

imposed by fluid flow and sloshing

(fuel), and internal loads (such as from

turbomachinery).[4] Thermal environments include aerodynamic heating,

heat from the combustion of rocket fuel,

and heat from the rocket exhaust plume.

Some of the most important considerations when selecting materials for use in

such launch vehicle systems are mass

efficiency (strength-to-weight), fatigue

properties, fracture properties, stiffness,

manufacturing processes (eliminating

stress concentrations), and environmental compatibility.

Staged Launch Vehicles

There are other aspects to launch vehicles

besides simply being expendable or

reusable. That is, they can be singlestage-to-orbit (SSTO), or have multiple

stages that boost the payload into its final

orbit. In multi-stage systems, the propellant is contained in several smaller tanks,

as opposed to keeping it all in one large

tank, such as in the SSTO systems. The

multi-stage system, therefore, allows for

the tanks to be jettisoned when the fuel

has been expended, thereby decreasing

the mass of the launch vehicle during the

next stage of its ascent.

Single-Stage-To-Orbit One of the attractions of using SSTO launch vehicle is its

similarity to the relative simplicity and

cost-effective operation of an aircraft.[5]

The primary disadvantage is that

because the entire launch vehicle rides

into orbit, a greater percentage of the dry

mass is required to be the vehicle rather

than the payload. From the SSTO performance equation, the takeoff weight of

the vehicle plus its payload (excluding

rocket fuel) can only be 10% of the total

liftoff weight if the desired vehicle is

intended for a low Earth orbit at 30,000

feet per second.[6] Although there have

been many attempts at developing an

SSTO launch vehicle, none thus far have

been entirely successful, including the

recent X-33, which is shown in Figure 5.

Multi-Stage-To-Orbit The vast majority

of launch vehicles in use today are of the

multi-stage design. This is primarily

because the mass fraction of the payload

can be higher compared to SSTO, since a

significant portion of the weight is discarded during the ascent to orbit. Each

separate propulsion unit is called a stage.

All launch vehicles have at least 2 stages,

with possible additional stages depending

on the mission.

The first stage contains the rocket

engine, fuel, and fuel tanks, as shown in

Figure 6. The center body section contains electronics which control the first

stage performance. The rocket engine

burns the rocket fuel. The strap-on rocket motors provide additional thrust to

help lift the launch vehicle from the

ground. The turbomachinery inside the

rocket motors compress the fuel to

increase the combustion efficiency. The

first stage rockets also provide stabilization to keep the vehicle following its

intended flight path. The main drawback

to using multi-stage launch components

is that the lower stages are used only once

and are disposed of during ascent – never

to be used again. Hence, the long-term

manufacturing costs are higher since new

components must be made for every

launch vehicle.

The second stage is used to boost

what’s remaining of the launch vehicle

through the final phases of ascent into an

orbit, and also can provide some attitude

control. An example of the components

that make up a second stage are shown in

Figure 7. A third stage can be used to

change orbital altitude, or may even be

used to push the spacecraft out of Earth’s

orbit. Both ELVs and RLVs employ orbit

transfer stages to move and position satellites into their final orbit. Reusable orbit

transfer vehicles are envisioned to perform

this same mission. Concepts for these

“space tugs” would use electric propulsion.

An example of a third stage motor is

shown in Figure 8. A payload fairing, as

shown in Figure 9, is used to protect the

spacecraft and upper stage from severe

atmospheric and aerodynamic conditions

(during launch and ascent). This protective structure is discarded (see Figure 10)

after the launch vehicle has cleared Earth’s

atmosphere, usually around the time the

second stage rocket ignites.

The AMPTIAC Quarterly, Volume 8, Number 1 21

Courtesy NASA Jet Propulsion Laboratory

Figure 9. The Upper Stages,

Payload, and Payload Fairing

of a Launch Vehicle[8].

Courtesy NASA Jet Propulsion Laboratory

Figure 10. Artistic Depiction of

the Release of the Payload

Fairing[8].

Courtesy NASA Jet Propulsion Laboratory

Figure 6. Components of the

First Stage of a Launch

Vehicle[8].

Courtesy NASA Jet Propulsion Laboratory

Figure 7. Components of the

Second Stage of a Launch

Vehicle[8].

Courtesy NASA Jet Propulsion Laboratory

Figure 8. Components of the

Third Stage Motor of a Launch

Vehicle[8].

Interstage

Wiring Tunnel

Centerbody Section

Solid Rocket Motors (9)

Oxidizer Tank

Fuel Tank

First Stage

Second Stage Third-Stage Motor

Separation Clamp

Band

Spin Table

Guidance Electronics

Helium Spheres

Nitrogen Spheres

Second-Stage Miniskirt

and Support Truss

Third Stage Motor

Attach

Fitting

Third Stage

Motor

Second

Stage 

Fairing

(2 parts)

Payload

(Mounts on third

stage motor)

Conical Section

Fairing Access Door
22 The AMPTIAC Quarterly, Volume 8, Number 1

Current Launch Vehicles

There are a number of launch vehicles currently in use. Each

one offers a variety of different features; including payload

capacity, reliability, etc. Most importantly, they support different payload masses. Current launch vehicles include the

Pegasus, Atlas, Delta, Space Shuttle, and Titan. Some ELVs are

shown in Figure 11.

PROPULSION

Propulsion systems serve three functions in spacecraft: 1) They

propel the spacecraft from lift-off into low-Earth orbit; 2) they

transfer the spacecraft between orbits, primarily from a lowEarth orbit to a high-Earth orbit; and 3) they provide thrust to

maintain the spacecraft in proper orbit and attitude.

Types of Propulsion

Components of propulsion systems includes turbomachinery,

lines, ducts, valves, control systems, fuel tanks, pressurant storage tanks, rocket nozzles, and exit cones. Liquid propellant is

stored in metallic tanks, usually aluminum, or titanium.

Pressurant tanks are often made of aluminum over-wrapped

with carbon/epoxy composites. The Space Shuttle Orbiter uses

many metallic tanks over-wrapped with aramid fiber/epoxy.

Reaction chambers and exit cones require high temperature

materials. Current materials include niobium (columbium),

other refractory and high temperature metals, and carbon/

carbon composites. Of course, the heart of the propulsion 

system is the type of propellant.[6]

A major consideration when using liquid oxygen and liquid

hydrogen for propellants is that they require very strict monitoring and control. Fuel boil-off, especially when the launch

vehicle is sitting on the launch pad on a hot day, can be significant. Another major concern is the density of hydrogen can

change significantly with small temperature changes. Since

there is an enormous volume of hydrogen propellant, a small

percent change in density can throw off the design weight of

the launch vehicle to a significant extent.

SATELLITE SUBSYSTEMS

Satellites are conceptually divided into two components, the bus

and the payload. The bus is common to all mission types and

consists of the following subsystems: structures, guidance and

navigation, attitude determination and control, electrical power,

communications, command and data handling, and thermal

control. The payload mainly consists of components to fulfill its

mission. The sensors are connected to the command and data

handling as well as the communication bus subsystems. The 

following sections describe the satellite bus subsystems and their

interaction with one another. Figure 12 is a depiction of a 

satellite’s subsystems and payload, although the structure and

placement of the subsystems will vary for the different missions

and even between different satellite manufacturers.

Structures

A satellite’s structure includes both the primary structure,

which supports the main body, as well as secondary structures

which support the various functional components. These structures are dictated by the mission type, and the launch loads.

The structural designs must be able to handle the mechanical

requirements of both on-orbit and launch events. The highest

mechanical loads a satellite will be required to endure are 

typically seen during structural testing, handling, and

launch.[3] Once in orbit, the loads become very small. In most

cases, a satellite is placed in an orbital path for its entire 

existence. In this case, the mechanical requirements are primarily for providing stiffness for dimensional stability, as pointing

accuracy of sensors and communication equipment is crucial.

The only maneuvers are associated with small orbital adjustments and deploying components, such as the solar arrays.

Aluminum alloys and composites are widely used in the 

primary structure due to their combination of low density and

mechanical properties. Secondary structures are typically made

from aluminum, steel, or titanium. Once in orbit, the ideal

weight of structures is zero. Common structural designs

include flat sandwich panels and trusses.

Figure 11. Examples of Expendable Launch Vehicles[3].

MSLS Athena 

I

Athena 

II

Titan 

II

Atlas 

II

Atlas 

IIa

Atlas 

IIas

Atlas 

IIIa

Atlas 

IIIb

Proton

D1e

Proton

M

Titan

IVa Titan

IVb

Atlas 

V

Atlas 

V

Heavy
The AMPTIAC Quarterly, Volume 8, Number 1 23

Guidance and Navigation

The guidance and navigation subsystem is responsible for

maintaining a satellite in orbit. The orbital path is determined

via communications with ground operations and/or other 

satellites. The data is processed by either ground operations or

on-board computers which then command the propulsion 

systems for maneuvering as required for orbital adjustments.

Guidance and navigation has traditionally been controlled by

ground operations. However, with the current GPS satellites

and on-board computers, autonomous orbit control is now an

option.

Attitude Determination and Control

The attitude determination and control subsystem maintains

the satellite’s orientation with respect to its center of gravity. Sun

sensors, horizon sensors and star trackers help determine a 

satellite’s position. There are two concepts used for attitude 

stabilization. The three-axes method may use gyroscopes,

propulsion systems, magnetic torquers, momentum wheels,

and/or reaction wheels to maintain its orientation. Spin stabilized satellites, or spinners, use rotation about the direction of

flight to maintain proper orientation. In most spin stabilized

systems, a component referred to as a ‘de-spun’ rotates with

equal angular velocity in the opposite direction, so as to be fixed

with respect to some reference point such as the Earth or other

satellite. 

With a trend towards higher power satellites, 3-axis stabilized

satellites are preferred since they offer more usable “acreage” for

solar cells. Other stationkeeping and propulsion is done

through propellant systems. Cold gas systems are not typically

used except when hot gases must be avoided or when safety

issues with the solid and liquid systems arise. Solid propellant

systems are used quite extensively, but only for space access.

Liquid bipropellant systems (fuel and oxidizer are separate

components) are attractive since they have the ability to serve

all three propulsion functions. However, they are also more

complex than the traditional solid plus liquid monopropellant

combined systems (fuel and oxidizer are combined). Dual 

and hybrid systems make use of both a monopropellant and 

bipropellant. Typically, hydrazine (N2H4) is used as a monopropellant and can be combined, using onboard systems, with

nitrogen tetroxide N2O4 to produce a bipropellant fuel.

Electrical Power

The electrical power subsystem converts energy, stores, and 

distributes electrical power to the various subsystems of the

satellite. The power system is typically designed to supply 2-3

times the average power required in order to cover peak

demands. It must also be designed to account for degradation

of the power systems over the life of the satellite. The photovoltaic solar cells lose efficiency over time and can suffer 

damage from the space environment. Batteries will lose power

storage capacity with use. Therefore the electrical power subsystem is determined by the average electrical power required at

the end of life. The most common power sources for Earth

orbiting satellites are the solar arrays (silicon or gallium

arsenide photovoltaic cells), and batteries (typically nickelcadmium). The solar arrays provide power and recharge the

batteries during the sunlit period, while the batteries provide

power when the satellite is in the Earth’s shadow.

Communications

The communications subsystem, also known as the telemetry,

tracking, and command subsystem, provides the link between

the satellite and outside systems. Transmitters and receivers,

and/or transponders are used for sending and receiving radio

frequency communications.

Command and Data Handling

The command and data handling subsystem receives and

processes information which is sent to ground operations or

on-board computers. It is also used to control the various 

satellite subsystems. The size and complexity of the command

and data handling system is directly related to the size of the

subsystems and complexity of the mission.

Thermal Control

The thermal control subsystem maintains the temperatures of a

satellite’s subsystems within their acceptable operating ranges.

This challenging task is accomplished by balancing heat 

generated within the spacecraft and exchanged with the environment. Spacecraft thermal control is complicated by the absence

of convection, which is a key method of heat transfer in terrestrial systems.[4] In space, thermal control has to rely on conduction and radiation. Key components of the thermal 

control subsystem are thermal control coatings, insulation,

heaters and radiators. Thermal control coatings include paints

and second-surface mirrors. The latter are polymers or glasses

with reflective coating, such as silver, on the side facing the spacecraft. Multilayer blankets consisting of polymer films with reflective metallic coatings, separated by fibrous mats typically provide

the necessary insulation. Heaters with thermostats maintain the

minimum operating temperatures of the various subsystems.

And radiators are used to expel excess heat into space.

Courtesy of NASA

Figure 12. Typical Satellite Subsystems[9].

Payload (Various Imaging Sensors)

Primary Structure

Solar Array

Communications &

Data Handling

High Gain Antenna

Antenna

Mounting Bench

Propulsion

Module

Guidance &

Control

Power Distribution &

Management
24 The AMPTIAC Quarterly, Volume 8, Number 1

SUMMARY

The US Air Force is a key player in the US space program, and

the Air Force Research Laboratory continues to advance space

technology with fundamental research and development that

applies to the various systems, subsystems and components 

discussed in this article. Many of the articles in this special issue

of the AMPTIAC Quarterly focus on specific components and

technologies used to build the various spacecraft subsystems.

These include rocket nozzles, thin-film membrane reflectors,

high precision mirrors, thermal control coatings, and thermal

protection systems.

REFERENCES

[1] http://www.peterson.af.mil/hqafspc/Library/Image_Library

/gps.htm 

[2] “Sun Synchronous Orbit”, US Centennial of Flight

Commission, http://www.centennialofflight.gov

[3] “Materials Selection and Manufacturing for Spacecraft and

Launch Vehicles”, Short Course, AMPTIAC, January 2000.

[4] C. Zweben and B. Battat, Material Selection for Spacecraft

and Launch Vehicle Applications – A State of the Art Report,

AMPTIAC, August 2002

[5] A. Gurney, “Spotlight on Technology: New Metallic

Thermal Protection Structures Paving the Way for Next

Generation Space Vehicles”, AMPTIAC Newsletter, Vol. 4, No.

2, 2000

[6] Space Mission Analysis and Design - 2nd Edition, eds. W.J.

Larson and J.R. Wertz, Kluwer Academic Press, 1992

[7] “X-33 by Lockheed Martin on Launch Pad – Computer

Graphic”, Dryden Flight Research Center, NASA, http://

www.dfrc.nasa.gov/gallery/photo/X-33/HTML/EC96-

43631-1.html

[8] “Launch Vehicle Parts,” Jet Propulsion Laboratory, NASA,

http://marsrovers.jpl.nasa.gov/mission/launch_vehicle.html

[9] NASA/Goddard Space Flight Center, 1997

Mr. Benjamin D. Craig has been a member of the AMPTIAC technical staff since January 2001. He has authored

or co-authored five articles published in the AMPTIAC Quarterly, as well as a number of AMPTIAC’s technical

products. His technical areas of focus include blast and penetration resistant materials, high temperature superconductors, and materials for sensors and actuators. Mr. Craig also manages the AMPTIAC web site, which serves

as a major resource for the defense materials community. He holds a BS in Chemical Engineering from Syracuse

University, and is currently finishing work on an MS, also in Chemical Engineering, from Syracuse University.

Mr. Richard A. Lane has been a member of AMPTIAC’s technical staff for the past 6 years. He has co-authored

several AMPTIAC newsletter articles and technical reports, on a number of different material technologies, including blast and penetration-resistant materials and nanostructured materials for mechanical enhancement. Mr. Lane

received a BS from SUNY Cortland, majoring in both Physics and Mathematics, and is currently pursuing an MS

in Mechanical Engineering at Syracuse University.

answers materials engineering questions

CALL : 315-339-7090

EMAIL : amptiac_inquiries@alionscience.com

O R

VISIT : http ://amptiac.alionscience.com/Help/inquiry.html

AMPTIAC
THE SPACE ENVIRONMENT

Space is an unforgiving environment – one of harshest known to mankind.

Yet, space is far less understood than other service environments in which

DOD systems operate. It has been known to limit the useful life of satellites,

with 20 to 25% of all mission failures being attributed to the effects of the

space environment.[1] Since access to space is so costly, materials and

components must function throughout the life of a mission without repair

or replacement. Yet, overdesign is not an option, as the cost of launch (driven by payload weight and volume) severely limits the redundancy and size

of components. Adding to these challenges, future spacecraft will be

designed to last longer, their functionality will become ever more complex,

and their power requirements will continue to increase.

The natural space environment (often referred to as space weather)

depends heavily on the satellite orbit, the time of year, and the solar activity. The environment is largely determined by the type of orbit. Low-EarthOrbits (LEO) provide maximum sensor resolution for weather and reconnaissance missions. Satellites and spacecraft in LEO are subjected to atomic oxygen and can experience greater thermal cycling than in other orbits.

Moving farther away from Earth, satellites in geosynchronous orbits

(GEO) around the equator have the same angular velocity as the earth,

thus GEO satellites remain over a fixed point on the Earth. Many communication and weather satellites are parked in GEO. Mid-Earth Orbits

(MEO) are between LEO and GEO. Systems in MEO are subjected to the

trapped radiation in the Van Allen Belts. A polar orbit is a special LEO 

to MEO orbit where the satellite travels over the poles. Since the Earth

turns underneath the satellite, it can cover the entire globe in a short time

period. Polar orbiting satellites typically experience higher radiation 

environments. A sun-synchronous orbit is a special polar orbit where the

satellite passes over the same part of the Earth at roughly the same local

time each day (for more information, please refer to the ‘Orbits’ section of

the article, ‘Materials and the Final Frontier’ in this issue). This can

make communication and various forms of data collection convenient. 

Space weather causes a range of problems to materials on spacecraft.

Thermal/vacuum effects, atomic oxygen, micrometeorites, ultraviolet

radiation, and electron and proton radiation can all adversely impact

space materials. The following sections describe these environmental 

factors and the materials problems they create. 

THERMAL/VACUUM EFFECTS

The atmospheric density at an altitude of 300 km (Space Shuttle orbit) is

ten orders of magnitude less than at sea level. Most organic materials will

outgas under these vacuum conditions. The majority of metallic and

ceramic materials do not outgas unless they have a high vapor pressure,

like cadmium or zinc. Organic outgassing products include moisture and

volatile organic compounds. Moisture sources include both desorption of

moisture trapped inside the organic material and loss of water vapor

adsorbed on the surface. For structural components, outgassing can lead

to changes in the dimensional stability for stiffness-critical components

such as optical benches. The more serious issue is that the outgassing

products can condense and contaminate critical surfaces such as

radiometers, solar cells, and sensors. Cold systems (<100°K), like infrared

detectors, are especially prone to condensing contamination.[2]

Microgravity can compound the effects of contamination products by

keeping the contaminants near the satellite. 

In addition, the vacuum environment eliminates convective cooling

as a form of heat transfer. Structures can only be cooled if the heat is conThe AMPTIAC Quarterly, Volume 8, Number 1 25

Material EASE

TESTING IN THE SPACE ENVIRONMENT

Jeffrey D. Guthrie

AMPTIAC

Rome, NY

Clifford Cerbus

University of Dayton Research Institute

Dayton, OH

Elizabeth S. Berman, Laura Rea

Materials and Manufacturing Directorate

Air Force Research Laboratory

Dayton, OH

25AMPTIAC

The previous article asked (and hopefully answered) the question, “What are the technical challenges of operating military systems in the space 

environment, and how do materials enable many of these missions to succeed?” Subsequently, this article poses the next logical question, “How do

we know that our material, component, assembly, or system will work in space?” Like applications in any other service environment (e.g. land,

sea, or air), the only way to validate material and design choices prior to fielding is to test them. For space systems, this is easier said than done.

The vacuum, microgravity, thermal, and radiation effects of space are extremely difficult to reproduce on the surface of the earth; thus only a handful of facilities exist which can reproduce some of these conditions.

This article explores some of the conditions of space environments that systems are designed to operate in; and how their materials and 

components are tested, both at earthbound facilities, and in space. We are sure you will find this an informative companion piece to the 

previous article. - Editor
26 The AMPTIAC Quarterly, Volume 8, Number 1

Material EASE

ducted to a point where it can be radiated to space. At lower earth orbits,

moving in and out of the earth’s shadow occurs frequently. For a 550 km

circular orbit, the satellite would experience 15 eclipses each day. Under

such circumstances, a number of special materials would be required to

provide thermal management for the satellite. Thermal management

typically accounts for roughly 5% of the mass of a satellite.[1] See our

article on Thermal Control Coatingsand Space Environmental Testing

for more details on thermal management.

Atomic Oxygen

Atomic oxygen is produced in the upper atmosphere when ultraviolet

radiation (UV) from the sun is absorbed by oxygen molecules (O2

) and

causes them to dissociate into negatively charged ions. Atomic oxygen is

prevalent in orbits between 100 and 650 kilometers. The atmospheric

density at these altitudes depends highly on both orbit and solar activity.

While the density is small, the ionic flux encountered is high given the

spacecraft’s velocity of orbit. At Space Shuttle orbits of 250 km to 300 km,

a density of 109 atoms/cm3 yields a flux of 8 x 1014 atoms/cm2

s. The high

velocity of orbit also gives atomic oxygen roughly 5 electron volts (eV)

impact energy per ion.[2] The carbon bonds of many organic materials

are susceptible to impact energies of this magnitude. As an example, the

polyimide film Kapton (a material used extensively for thermal blankets

on spacecraft) shows erosion rates of 150 mils (milli-inches) per year in

a 200 km orbit. Additionally, a few metallic materials (such as silver)

exhibit degradation/erosion in atomic oxygen. Some metals that form

protective oxide films are susceptible to degradation, since low gas pressures can inhibit the formation of adherent oxide films. Atomic oxygen

can also excite some materials to emit radiation. The glow around the

Space Shuttle is attributed to this phenomenon. 

Particle Impact

Spacecraft must also contend with impact of particles either from space

debris or dust. Most of the space debris is found in low earth orbits from

350 km to 2000 km. The debris largely consists of materials ejected from

spacecraft during launch and deorbit, including alumina from solid

rocket motor propellants, aluminum, and oxides from thermal control

coatings (zinc and titanium oxides). All told, it is estimated that there is

over 3,000,000 kg of space debris. The United States Strategic Command

currently tracks over 8,000 objects in space that are over a centimeter in

size - but there are an estimated 1,000 kg of particles under this limit.[3]

Relative impact velocities depend upon the satellite orbit but are on the

order of 10 km/s. Particles as small as 0.1 mm cause erosion on most

materials and particles greater than 1 mm cause significant damage.[1] 

Micrometeoroids and interplanetary dust also can damage spacecraft. Most micrometeoroids are a result of comet ejection. These particles typically have velocities relative to the earth from between 10 to 70

km/s. They are at least 0.2 microns in diameter since smaller particles

are swept away by the solar wind.[4]

Space Radiation

Radiation can wreak havoc on materials, causing either permanent or

transient changes in their physical and electrical properties. Radiation

comes in electromagnetic and particle forms. Nearly 99% of the solar

electromagnetic radiation at one astronomical unit (distance from the

earth to the sun) has a wavelength between 0.3 and 11 μm. In fact, visible light accounts for 41% of the total electromagnetic energy emitted

from the sun. Together, both the ultraviolet and infrared spectrums

account for most of the remaining energy. X-Rays, extreme ultraviolet,

and radio wavelengths account for less than 0.1% of the solar energy.

Solar electromagnetic radiation changes with solar activity and season

(the distance between the earth and the sun changes, as the earth’s

orbital path is slightly elliptical).[5] Reflected radiation (albedo) from

both the earth and the moon contribute to the total irradiance. The albedo is important to low earth orbits and varies with clouds and terrain.

Ultraviolet Radiation: Ultraviolet radiation (UV) is particularly damaging to some materials. Most UV radiation is absorbed by the Earth’s

ozone layer. However, in space, high intensity UV radiation can sever

organic chemical bonds. In addition, UV can lead to discolorations in

polymeric materials and ceramics (e.g. glasses), which change the solar

absorptivity of materials. Changes to a material’s performance, such as

the absorptance of white thermal control coatings must be accounted for

in the design process, since the rate of heat transfer will change proportionately. As Figure 1 shows, changing the solar absorptance from 0.2 to

0.4 over the life of a satellite dramatically increases the needed radiator

area from 375 ft2 to 600 ft2 for a 10 KW satellite. Note, the emissivity (ε)

of the white coating does not change much during the satellite life and it

is shown as a constant 0.80. Designers must account for end-of-life properties by overdesigning components such as radiators. However, increased

Figure 1. Radiator Size vs. End-of-Life Solar Absorptance.

1,100

1,000

900

800

700

600

500

400

300

200

100

10KW

5KW

1KW

Solar Absorptance (αs

)

ε = 0.80

T = 300K

Radiator Area (ft

2)

.1 .2 .3 .4 .5 .6

~375 ft2

~600 ft

2
area translates to extra weight. It is important to note that as the trend to

higher power satellites continues, accounting for the degradation of 

thermal control paints will become increasingly critical during their

design phase. 

Charged Particle Radiation: The sun also emits a continuous stream

of energetic charged particles. These electrons and protons (and some

heavier ions) are called the “solar wind”. On average these particles 

travel at over 400 km/s with a density of 5 particles per cubic centimeter.

During solar flares the charged particles are accelerated. In addition,

particles emitted by other stars pass through our solar system. Charged

particles from a solar flare are referred to as solar cosmic radiation while

particles from other stars are known as galactic cosmic radiation.

Electrons and protons found in cosmic radiation typically have much

higher energy levels than those found in the solar wind.

The earth’s magnetic field traps charged particles in radiation belts

known as the Van Allen Belts. The Van Allen Belts consists of two concentric donut-shaped belts that surround the Earth. Because of the effect

of the solar wind on the magnetic field, both belts are asymmetrically

shaped. On the Earth’s sun-facing side, its magnetic field is compressed

from the solar wind, while the field is expanded on the other side. The

magnetic fields have funnel shaped cusps over the poles that allow penetrating radiation to reach much lower altitudes. In the inner Van Allen

belt, the maximum proton density is found at about 5,000 km above the

Earth’s surface. This belt contains mostly high energy protons produced

by cosmic ray collisions with the Earth’s upper atmosphere. The particles

in this belt are highly energetic, and with prolonged exposure they can

damage materials. The outer belt’s maximum proton density occurs at

16,000 to 20,000 km. This belt contains low to medium energy electrons

and protons mostly generated from the influx of particles during geomagnetic storms, which occur after solar flares. These particles are

deflected by the Earth’s magnetic field, but are captured on the night side

of the planet. Particle densities can increase by a factor of 10 to 1000 in

less than an hour.[6] 

Effect of Energetic Particles: Charged particles can damage an

impacted material by a combination of ionization and atomic displacements. These effects can lead in turn to chemical reactions and changes

in the local morphology. Electrons slow by generating a continuous ionized path in materials. Protons are heavily ionizing and produce higher

ionization near the end of the particle path. The size of damage depends

upon total flux, intensity of radiation, and the angle of impingement.

The interaction of these particles with organic materials can lead to outgassing, shrinkage, cracking, pitting, embrittlement and discoloration.

Energetic particle radiation also influences the absorptivity of materials.

As seen with UV radiation, the solar absorptance of thermal control coatings show an increase over time. For example, when simulating the

Galileo mission, the solar absorptance of Zinc OrthoTitanate (ZOT), 

a thermal control coating, increased from 0.13 to 0.42.[7] Charged 

particle radiation also can lead to “darkening” of optical materials.

Satellite electrical charging (the buildup of static electricity on the

satellite surface) is one of the most common anomalies caused by radiation. Charging occurs when a satellite moves through charged particles

(called “wake charging”). Charging can also occur through direct 

particle bombardment or through a photoelectric effect, with some 

materials subjected to electromagnetic solar radiation. When the 

breakdown voltage of a material is exceeded, an electrostatic charge will

ultimately result in a discharge or arc; which can severely damage 

materials, cause spurious circuit switching, or result in false sensor readings. Materials especially prone to damage include electronic materials,

thermal coatings, and solar cell materials.[5]

The high-energy protons from either cosmic radiation or the inner

Van Allen Belt can completely penetrate a satellite. As they pass through,

they can ionize particles deep inside the system. A single proton can

deposit enough charge to cause a circuit switch, spurious command, or

memory change. These incidents are labeled “single event upsets”. With

a push towards lower voltage commercial-off-the-shelf microelectronics,

it is actually easier to cause electrical upsets than with older satellites.

Additionally, the lessening of the perceived nuclear threat has reduced the

level of nuclear hardening of satellites. While intended to protect systems

from the radiation of a nuclear explosion, this type of hardening also

helped to protect satellites from the radiation naturally present in the

space environment.[5] 

Many satellites use electro-optical sensors to maintain their orientation in space. These sensors reference the direction of certain stars to

achieve precise pointing accuracy. When a high energy proton impacts a

sensor material, a flash of light is produced and misinterpreted as a star.

The computer will then try to find the star in its star catalogue. At this

point the satellite could become disoriented, creating the possibility that

communications antennas, sensors, solar panels all could fail to orient

themselves properly, thus impacting the mission. In extreme cases, the

satellite could even be lost if the batteries were to drain from a lack of

solar power. For this reason, nearly all satellites have a sun sensor. While

not providing the navigational accuracy of a star sensor; a sun sensor can

help the satellite recover its orientation following an upset.[5]

SPACE COMBINED EFFECTS PRIMARY TEST RESEARCH

EQUIPMENT (SCEPTRE)

The extreme conditions present in the space environment and the high

cost associated with placing space systems in orbit require that spacecraft

be thoroughly tested prior to deployment. Every spacecraft program must

address the effect of the space environment on their hardware.

Additionally, all new materials must be space flight qualified prior to use.

Ground-based testing offers the flexibility to study multiple materials in

the appropriate environmental conditions, without the extreme cost and

limited availability of space flight testing. New materials can be subjected to an initial screening, with a rapid turnaround of test results.

Additionally, ground-based testing allows for multiple sample exposures

of a material to gain statistical confidence in its life cycle performance.

Even small modifications in current state-of-the-art materials, such as

changing a raw material supplier or imposing a material processing

change, requires satellite manufacturers to initiate a re-qualification

program to assure that the new version of the same material performs at

least as well as the previous one. Ground-based testing is appropriate for

these re-qualifications.

The space environment is difficult to simulate and no facilities exist

that can test all of its aspects. On January 31 1958, Explorer I (the first

US satellite) discovered charged particle radiation in space using a Geiger

A DVANCED M ATERIALS AND P ROCESSES T ECHNOLOGY

The AMPTIAC Quarterly, Volume 8, Number 1 27

AMPTIAC
28 The AMPTIAC Quarterly, Volume 8, Number 1

Material EASE

counter package designed by Dr. James Van Allen. This discovery led to

the acknowledgment that environmental conditions in space were much

harsher than originally thought. Later in 1958, the predecessor to the Air

Force Research Laboratory’s Materials and Manufacturing Directorate

(AFRL/ML) opened a facility to examine some of the environmental conditions found in space, including the combined effects of vacuum and UV.

Over the years, its capabilities have increased to include the effects of

electron bombardment. Today, it is known as the Space Combined Effects

Primary Test and Research Equipment (SCEPTRE) facility. SCEPTRE is

the only Air Force-owned facility designed specifically for testing and

qualification of spacecraft materials by exposing them to a simulated

space radiation environment (Figure 2). 

SCEPTRE is designed primarily to simulate the conditions of MEO to

GEO with variable temperature, accelerated UV, near visible UV, and electron exposures; all simultaneously in a vacuum environment. SCEPTRE

does not simulate the proton environment, nor the extremely high energy particles associated with cosmic radiation or the inner Van Allen Belt.

With an eight-inch target area, SCEPTRE can expose 19 one-inch 

diameter samples and four reference samples per test. During exposure,

SCEPTRE measures specimen spectral reflectance, solar absorptance,

temperature, electron flux, gross outgassing, and the spectral irradiance

of the UV source. After exposure, SCEPTRE can measure the thermal

emittance. Additional pre- and post-test measurements can be made in a

state-of-the-art optical measurements facility. These optical facilities

measure reflectance, absorptance (emittance), transmittance, and 

scattering as functions of wavelength (0.25 to 15 mm), incidence angle

(0 to 89°), polarization, and temperature (-65°F to +450°F).

Flight test data confirms that SCEPTRE simulates a variety of orbital

environments and accurately reproduces the rate and extent to which a

material will degrade in the actual space environment. The Air Force

thermal control coating development program is predicated on performing in-house space environmental testing and validation. To date, 

SCEPTRE is responsible for greatly advancing the development of 

thermal control materials and coatings. In addition, SCEPTRE has

recently expanded its testing to a wide variety of materials including 

polymeric films, optical thin films, and threat protection materials. 

The facility is also used in charging studies, thin film space stability

studies, and degradation mechanics studies.

The University of Dayton Research Institute (UDRI) is currently

responsible for the operation of the SCEPTRE facility. Under a 1999

Cooperative Research and Development Agreement, UDRI performs

material degradation testing on a fee-for-service basis for both government and non-government sponsors.

SPACE FLIGHT TESTING 

Despite advancements in ground-based testing, flight experiments are

necessary to verify ground-based simulation and analytical modeling

efforts. Additionally, flight testing helps validate new materials technology and demonstrates their performance in components. The Materials

and Manufacturing Directorate has a rich history of space flight testing

despite the often limited availability of such tests. Some of the past experiments include: Spacecraft

Charging at High Altitude

(SCATHA), Skylab, Long

Duration Exposure Facility

(LDEF), and the Materials on

the International Space

Station Experiment (MISSE).

Several examples of 

lessons learned from prominent flight experiments that

AFRL/ML has participated in follow. The objective of these experiments

was to gain understanding of the changes in the properties before and

after exposure to the space environment. This knowledge helped

improve predictions based on laboratory experiments and subsequently

influenced the design of future flight experiments.

Skylab 

Skylab was America’s first space station. It was launched on 14 May 1973

by a Saturn V rocket and placed in a low earth orbit (415 km). Three 

separate astronaut crews occupied the facility for a total of 171 days, 13

hours. The last manned flight returned on 8 February 1974. During this

time, Skylab completed 3900 orbits during a period of low solar activity.

On July 11, 1979, Skylab returned to Earth and scattered its debris from

the Southeastern Indian Ocean across a sparsely populated section of

Western Australia.

The Skylab Thermal Control Coatings and Polymeric Films

Experiment[8, 9] consisted of three duplicate sets of 36 thermal control

coating samples and eight different polymeric film specimens. The thermal control trays were mounted perpendicular to the incident light from

the sun for maximum solar exposure. The polymeric films were located

39 degrees off axis to the solar vector. The first set of specimens was

retrieved after 35 days/550 hours of solar exposure and the second set

after 74 days/1150 hours of solar exposure. A third set of samples was

retrieved after 131 days/2040 hours of exposure. All samples were badly

contaminated during launch. Post flight analysis of the thermal control

coatings indicated that measured changes in specimen thermo-optical

properties were due to a combination of excessive contamination and

solar degradation. The results from this experiment were used in the

design of the Long Duration Exposure Facility M0003-5 Thermal Control

Materials Experiment and offered a direct comparison for many on the

materials between the two flight experiments.

Long Duration Exposure Facility (LDEF)

The Long Duration Exposure Facility was launched into LEO (410 km

orbit) from the Space Shuttle Orbiter Challenger in April 1984 and

Figure 2. The SCEPTRE Chamber.
A DVANCED M ATERIALS AND P ROCESSES T ECHNOLOGY

The AMPTIAC Quarterly, Volume 8, Number 1 29

AMPTIAC

retrieved by the Columbia in January 1990. LDEF contained 57 experiments from the Air Force, Navy, NASA, European Space Agency, academia,

and industry. LDEF contained over 10,000 materials specimens subjected

to 32,422 Earth orbits. The 5.8-year flight greatly enhanced the potential

value of most LDEF materials, compared to that of the original 1-year

flight plan. Most of the exposed surface of LDEF consisted of chromic

anodized aluminum and silver Teflon thermal blankets. 

Upon retrieval, particulate debris was observed trailing LDEF. This

debris mostly consisted of vapor-deposited aluminum backing from

failed thermal-control blankets on the leading edge of the spacecraft. In

general, organic materials such as Mylar, Kapton, paint binders, and

bare composites were severely eroded. The degradation of these materials was expected since they received prolonged exposure to atomic oxygen. Coated composite materials survived and generally maintained

their mechanical properties. 

AFRL/ML and the Aerospace Corporation conducted the M0003

experiment on LDEF which consisted of coatings, thermal paints, polymers, glasses, composites, semiconductors, and detectors that provided

data on environmental parameters. Most of the materials selected were

from various development programs in the 1978 to 1982 time frame.

Materials from previous space flight experiments,

including Skylab, were included to provide data

correlation. The materials were exposed in 

four separate locations on the vehicle. The first

set was exposed on the leading edge of the 

satellite. The second set was exposed on the 

trailing edge of the vehicle. The third and fourth

sets were exposed in environmental exposure

control canisters located 30 degrees off normal 

to the leading and trailing edges. 

The leading edge tray showed atomic oxygen

damaged some materials, especially to two silver

front surface mirrors, which were destroyed.

There was extensive damage to the polymeric film strips. Composite

material samples remained intact, but they were bleached or discolored.

Silver Teflon covers remained intact; however, they had changed from 

silver to milky white, where components were located beneath them; and

to a gold color where there were no components. The trailing edge

showed that contamination discolored many of the specimens. Again, the

polymeric film strips were damaged. The thermal control materials were

discolored from both contamination and radiation.

Materials on International Space Station Experiment (MISSE) 

The first of four Materials International Space Station Experiments

(MISSE-1) was attached to the International Space Station (ISS, 350 km

orbit) on 10 August 2001. MISSE is a cooperative experiment involving

Boeing Phantom Works, AFRL/ML, NASA’s Langley Research Center,

NASA’s Marshall Space Flight Center, and NASA’s Glenn Research Center.

AFRL/ML coordinated the project between the participating agencies and

aerospace companies. In addition, AFRL/ML funded the integration of

the passive experimental carrier “suitcases” that carry and hold the

materials test samples. The AFRL/ML-led “Space Systems Support and

Affordability Effort” (S3AE) consortium, comprised of major spacecraft

manufacturers, shared in the experiment. Over 450 candidate spacecraft

materials on MISSE-1 (1,700 samples) are currently being exposed to 

the space environment during solar maximum conditions (Figure 3).

Some of the materials classes represented include: bulk polymers, composites, thermal control coatings, gossamer films, inflatable materials,

optical materials including mirror materials, and thermal protection

materials. 

In addition to helping coordinate activities under MISSE, Mr. Edward

Stutz from AFRL/ML’s Sensor Materials Branch (AFRL/MLPS) designed

an experiment to evaluate various emerging semiconductor materials on

the space station. The experimental matrix focused on one emerging

class of high performance semiconductors, which may have inherent

radiation tolerance. Known as “Wide Bandgap Semiconductors” due to

their inherent electronic properties, these materials have already led 

to the demonstration of exceptionally high power density devices, as well

as exceptional low noise amplifiers (LNAs) which are critical for space

electronics. 

The MISSE Semiconductor Materials Evaluation Experiment incorporated a broad spectrum of wide bandgap materials, including gallium

nitride (GaN), silicon carbide (SiC) and zinc oxide (ZnO). Mr. Stutz coordinated this activity with a complimentary

program sponsored by AFRL’s Sensors

Directorate (AFRL/SN). AFRL/SN delivered

GaN High Electron Mobility Transistors

(HEMTs), as well as basic epitaxial materials

structures that were included in the MISSE

experiment. The devices used are the result of

a joint collaboration between Cree, Inc

(Durham, NC) and AFRL/SN under a Dual

Use Science and Technology program. The

program focused on the development of wide

bandgap semiconductors for microwave technology applications which are capable of

operating under extreme conditions, such as a crowded or hostile electromagnetic spectrum; as well as under limited environmental controls

due to constraints of mass, volume, or prime power. 

All materials and devices flying on the MISSE payload were 

extensively characterized by structural, optical, and electronic measurement techniques prior to inclusion on the mission. Upon return, the

materials and devices will undergo further testing. The test results will 

be analyzed to provide an assessment of the impact of the actual space

environment on performance. Results of this work will be instrumental

in guiding the future development of semiconductor materials and

devices for space applications.

SUMMARY

Exploiting space requires a thorough understanding of the space 

environment’s effects on materials. The criticality of new materials to Air

Force space missions dictate that all materials be qualified before use.

The Air Force Research Laboratory’s Materials and Manufacturing

Directorate has an extensive 45-plus year history of studying the 

fundamental interactions of the space environment’s effect on materials,

and to qualify new materials for use in space. Using its unique groundFigure 3. MISSE Integrated on the

International Space Station.
based facility, SCEPTRE, AFRL/ML can approximate the MEO and GEO

environments, and thus measure changes in optical properties resulting

from exposure to these conditions. In addition, AFRL/ML continues to

aggressively pursue space flight testing opportunities to strengthen the

understanding of this complex environment. 

For more information on SCEPTRE, please contact Dr. Elizabeth

Berman, elizabeth.berman@wpafb.af.mil, or Mr. Clifford Cerbus, 

clifford.cerbus@wpafb.af.mil. For additional information on MISSE’s

wideband gap semiconductor experiments, please contact Ms. Laura Rea,

Laura.Rea@wpafb.af.mil.

REFERENCES

[1] Carl Zweben, “AMPTIAC/USAF Space Applications of Materials

Training Course”, January 18-21, 2000 Wright-Paterson Air Force Base

[2] Edward M. Silverman, “Graphite / Thermoplastic Material

Development for Space Applications”; 24th International SAMPE

Technical Conference, October 20-22, 1992

[3] M.J. Meshishnek, “Overview of the Space Debris Environment”, SMCTR-95-9, March 1995

[4] W.C. Riley, “Space Environmental Testing of Advanced Materials”,

The Aerospace Corporation, SAMSO-TR-70-218, April 1970

[5] F. Guy, “Space Environmental Impacts on DOD Operations”, Air

Force Space Command Pamphlet 15-2, October 2003

[6] Anon., “Air University Space Primer”, Air University Maxwell Air

Force Base, AL, August 2003, Chapter 6: Space Environment

[7] F.L. Bouquet, “Responses of Dielectrics to Space Radiation”, NASA

Tech Brief Vol. 10, No. 5, Sept 1986

[8] W.L. Lehn and C.J. Hurley. AIAA/AGU Conference on Scientific

Experiments of Skylab, Huntsville, AL, Oct. 30-Nov. 1974, AIAA Paper 

No. 74-1228

[9] W.L. Lehn and C.J. Hurley. AIAA 10th Thermophysics Conference,

Denver, CO, May 27-29, 1975, AIAA Paper NO. 75-689

30 The AMPTIAC Quarterly, Volume 8, Number 1

Material EASE

Mr. Jeffrey D. Guthrie has been a senior materials engineer for AMPTIAC since 1998. Prior to joining AMPTIAC,

he served as a liaison between the Air Force Philips Laboratory (now AFRL/VS and AFRL/DE) and the Materials

Directorate. Mr. Guthrie also spent six years working on Carbon-Carbon Composites for space applications for the

Structural Materials Branch (AFRL/MLBC). Mr. Guthrie holds BS and MS degrees in Metallurgical Engineering

from Montana College of Mineral Science and Technology.

Mr. Clifford Cerbus is a research physicist with the University of Dayton Research Institute (UDRI). For the past 

16-years, he has managed and operated the United States Air Force Research Laboratory’s space simulation 

chamber known as the SCEPTRE Facility. He has also participated in spaceflight experiments including the Long

Duration Exposure Facility (LDEF) and the Materials on the International Space Station Experiment (MISSE). 

Mr. Cerbus received his BS in Physics and Mathematics from Xavier University and his MS in Mathematics from 

the University of Dayton.

Dr. Elizabeth S. Berman is a senior materials research engineer and the space materials direction leader in the

Nonstructural Materials Branch within the Materials and Manufacturing Directorate (AFRL/MLBT) at WrightPatterson AFB, OH. She is in charge of developing both thermal control coatings and tribological materials for

space. She is the government lead for the only Air Force owned facility capable of simulating the space environment for testing of space materials. She also is the AFRL/ML Program Manager for the AFRL MISSE (Materials on

the International Space Station Experiment). Dr. Berman holds a BS in Chemical Engineering from the University of

Minnesota, an MS in Chemical Engineering from Purdue University, and a PhD in Polymer Science from

Pennsylvania State University.

Ms. Laura Rea has served as a program manager for compound semiconductor materials development for the Air

Force Research Laboratory for over fifteen years. Her responsibilities as Technology Advisor for the AFRL Sensor

Materials Branch include oversight of an electronic and optical materials portfolio with an annual budget of over

$20M per year. In 1998, she was selected to serve a one-year assignment as a member of the staff at the Office

of the Secretary of Defense, supporting the Director of Defense Research and Engineering, at the Pentagon.
The AMPTIAC Quarterly, Volume 8, Number 1 31

In an age when shuttles routinely pierce the atmosphere, astronauts visit the International Space Station, and hundreds of

satellites quietly sail thousands of miles overhead; access to

space means more than just exploring the unknown. It means

access to satellite television, Global Positioning and domination

of the battlefield from the highest possible ground.

Unfortunately, access to space also comes with a high price

tag. In fact, the current cost of putting payloads into space is

estimated to be $10,000 per-pound. Given the inherent risks of

launch and deployment, combined with the substantial financial investment in these systems, government agencies and 

corporations make every effort to ensure that their spacecraft

and satellites are built with materials and manufacturing technologies that will ensure the integrity and functionality needed

to accomplish their varied missions. Not surprisingly, as public

and private ventures into space become even more frequent

occurrences, the impetus to knock one or two zeroes off the

price of launching and parking these systems in orbit becomes

more paramount.

American airpower visionary, Gen. H.H. “Hap” Arnold

once said, “The first essential of air power is pre-eminence in

research.” In keeping with this vision, scientists and engineers

at the Air Force Research Laboratory’s Materials and

Manufacturing Directorate’s Laser Hardened Materials

Evaluation Laboratory (LHMEL), are conducting the cutting

edge research General Arnold called for – research that will

allow the Air Force to dominate space both now and in the

future. Established as a world-class material characterization

facility, LHMEL provides results the government and industry need to lower costs and improve spacecraft and satellite

capabilities.

ABOUT LHMEL

The LHMEL is a “one-stop” infrared testing resource; 

providing the Air Force with basic laser-material interaction

data, optical material characterization, laser hardening concept

validation, and thermal simulation capabilities using a unique

collection of laser wavelengths, power levels and operating

modes to deliver high quality material response data. LHMEL

supports the Materials and Manufacturing Directorate’s (ML)

mission to provide laser protection materials and hardening

expertise for Department of Defense (DOD) personnel and

systems. 

LHMEL provides data on laser-materials interaction for 

current and new materials using a wide range of infrared laser

sources. This comprehensive testing resource provides reliable,

economical laser testing in a calibrated National Institute for

Standards and Technology (NIST)-traceable environment. 

HIGH-ENERGY TESTING

The LHMEL uses a collection of high-energy gas and solid

state lasers along with environmental simulation equipment

(such as vacuum chambers, wind tunnels, and structural loading machines), diagnostics, and data acquisition systems to support testing needs for the DOD, the National Aeronautics and

Space Administration (NASA), and commercial industry. Tests

typically involve experiments to investigate material response,

demonstrate system performance, or validate a hardening concept. As a result of efforts to diversify the type and increase the

quantity of facility users, recent efforts also investigated the use

of lasers both as thermal simulation sources and as tools to

assist in materials processing.

The high-energy testing area is currently equipped with three

John J. Eric

Robert J. Hull (Anteon Corporation)

Timothy R. Anderl (Anteon Corporation)

Materials and Manufacturing Directorate

Air Force Research Laboratory

Wright-Patterson AFB, OH

Material and Process Technologies are some of the critical enablers that permit spacecraft and launch vehicles to operate in the harsh 

environs of outer space. While developing and identifying these materials are vital, validating their performance in a prospective service

environment is equally important. Short of taking material samples into space, evaluation of their properties must be performed in earthbound test facilities. Laboratories capable of reliably testing materials in simulated space environments are few. The Air Force Research

Laboratory (AFRL) is home to one such state of the art facility that serves as a resource to both the military and civilian aerospace 

community. - Editor
32 The AMPTIAC Quarterly, Volume 8, Number 1

unique lasers: LHMEL I, LHMEL II, and the neodymium

(Nd):Glass laser (Figure 1). LHMEL I, the workhorse of the

facility, is a 15 kW continuous wave (CW), carbon dioxide

(CO2) laser capable of supporting as many as 50 tests per day.

The turnkey design makes it ideal for low-cost, high-throughput material response testing and laser processing research. Its

primary assets include excellent beam quality, predictable output, and NIST-traceable diagnostics.

With 150 kW of continuous wave output power, LHMEL II

is the nation’s largest CO2 laser (Figure 2). In addition to providing a dynamic range of laser powers and spot sizes for general material testing, this laser serves as a mid-scale bridge

between laboratory level “science” and costly full-scale phenomenology and effects testing. A virtual twin to LHMEL I as

far as reliable performance and testing throughput, LHMEL II

is a one-of-a-kind, high-energy test resource for the DOD.

LHMEL II is routinely used to economically validate material

response models prior to conducting high-cost, full-scale 

validation tests at other facilities.

Donated to the Air Force in 1993, the Nd:Glass laser 

provides LHMEL with a near infrared, pulsed source with up

to 10 MJ of repetitively pulsed energy in either q-switched or

long pulse mode. Its flexible configuration permits the 

simulation of a variety of laser types and pulse trains. When

coupled with in-house optical elements, this device provides

one-of-a-kind, high energy pulses in the near infrared, visible,

and ultraviolet wavelength regions. This device can produce 

5 kJ macropulses in the 0.5 or 5 msec pulse length regime or

100 J pulse streams in the 20 to 100 nsec pulselength regime.

LHMEL EXPANDS ITS SIMULATION CAPABILITIES

A recent cooperative effort between the Materials and

Manufacturing Directorate and the Air Force Space Battlelab

resulted in an upgrade of the Laser Hardened Materials

Evaluation Laboratory. Because of these efforts, the facility has

been expanded beyond laser effects testing to include largescale thermal testing and space environment simulation. The

Space Battlelab, located at Schriever AFB, CO, operates under

the direction of the Air Force Space Command and directly

supports combat operations through space systems.

The collaboration between ML and the Space Battlelab

enabled LHMEL to activate a highly-capable, 27-foot tall, 

20-foot diameter, cryogenically-shrouded vacuum chamber

(Figure 3). In addition to activating the chamber, LHMEL

added a one-sun solar simulator and a cryogenic and high 

vacuum-compatible turntable, which can be used to mount

and rotate samples 360° in the chamber. These new capabilities, which allow researchers to simulate the effects of the space

environment on materials and systems, are valued at over $25

million dollars.

Originally built by the Propulsion Laboratory in the early

1960s (now the Air Force Research Laboratory Propulsion

Directorate), the vacuum chamber was originally used for space

materials and systems qualification work. The Materials and

Manufacturing Directorate Hardened Materials Branch, which

oversees LHMEL operations, and the Space Battlelab shared

the cost of modifications required to make the chamber operational after years of sitting idle. 

In April 2002, after the chamber

became operational, scientists and

engineers at LHMEL began seven

weeks of thermal and environmental

testing of a space-qualified, fully functioning, operational micro-satellite

(Figure 4). Tests simulated conditions

that the satellite would encounter during its lifetime, including regular solar

cycles representing excursions into day

and night. A simulated ground station

collected telemetry data and received

signals produced by the satellite 

during testing.

The largest of three vacuum chambers currently available at LHMEL,

the new space simulation chamber,

“pumps-down” using a series of vacuum pumps and its cryoshroud, an

internal chamber liner that is filled

with liquid nitrogen, from ambient

pressure (760 torr) to 10-6 torr in an

Figure 1. Nd:Glass Laser Laboratory.

Figure 2. The LHMEL II High-Energy CO2 CW Laser.
The AMPTIAC Quarterly, Volume 8, Number 1 33

hour and a half. In addition to aiding the vacuum “pumpdown” process by attracting and trapping stray molecules for

the diffusion pumps to carry out; the cryoshroud simulates the

cold temperature of space, reaching a temperature as low as 

77° Kelvin.

The chamber is also equipped with vacuum “feed-throughs”

so that scientists can monitor samples or systems with diagnostics equipment such as thermocouples, strain gauges,

and infrared (IR) detectors. During testing of the

microsatellite, 32 thermocouples were connected via

these “feed-throughs” to data collection equipment outside of the chamber.

The LHMEL space simulation chamber is situated to

enable laser interaction testing in a large area, highvacuum, simulated space environment. Connected to

LHMEL II, scientists working at the facility can use the

150 kW laser to focus a low energy level beam on the

target, simulating sun-equivalent numbers on a six-foot

diameter area. 

The Laser Hardened Materials Evaluation Laboratory

has two additional vacuum chambers; one measuring

seven feet in diameter by nine feet in length, and the

other measuring 30 inches in diameter by 30 inches in

length. The seven-foot by nine-foot chamber was built

during the LHMEL facility upgrades in the late 80s. It

can achieve vacuum levels of 10-6 torr and can also be

operated in conjunction with the LHMEL II laser. This

chamber has been used for a variety of simulated space

environment testing scenarios, including thermal simulation

testing for solar cells and arrays for the Propulsion Directorate’s

Survivable Power Systems for Satellites (SUPER) program. In

addition, the chamber has been involved in NASA-sponsored

programs that examined space debris removal. The concept

behind the program was to introduce heat and energy to one

side of a piece of space junk, changing its momentum and

knocking it out of its stable orbit, and allowing it to enter the

atmosphere and burn up. LHMEL scientists have also used this

chamber for theoretical testing of laser-based propulsion concepts, evaluating the idea that objects can be propelled through

space using lasers.

The remaining 30-inch x 30-inch chamber, which is

portable, is used for materials phenomenology experiments 

at the basic material level. Most often, this chamber is used in

conjunction with the 15 kW LHMEL I laser to test basic 

material responses of two-, four- and six-inch test coupons. 

SHUTTLE REUSABLE SOLID ROCKET MOTOR PROGRAM

The LHMEL facility has also supported space-related testing

outside of the vacuum chamber. During the last five years, 

scientists at LHMEL have used the LHMEL I and LHMEL II

lasers to simulate the thermal environment inside solid rocket

nozzles (upwards of 4,000° Kelvin) in support of the Reusable

Solid Rocket Motor (RSRM) program. 

Engineers have taken advantage of LHMEL’s unique capabilities to reduce the cost of testing composite nozzle materials

(used on solid rocket motors) by over 90 percent. Using the

150 kW laser, LHMEL provided a reproducible, controllable,

and affordable testing environment, capable of simulating the

thermal conditions inside a solid rocket motor nozzle during

firing. 

Solid-fuel rockets use nozzles made of carbon-phenolic 

composite to focus and direct hot exhaust gases. Some charring

and lifting of plies in the composite material is expected and

acceptable during firing, but too much leads directly to nozzle

failure. Nozzle thickness is critical – excess thickness adds 

Figure 3. An Outside View of LHMEL’s Cryogenically-Shrouded

Vacuum Chamber.

Figure 4. Thermal and Environmental Testing of a Space-Qualified MicroSatellite.
34 The AMPTIAC Quarterly, Volume 8, Number 1

costly weight and reduces payload; not enough thickness

destroys the nozzle and its usefulness. Traditional nozzle testing

was complicated and expensive, involving the ground firing of

a solid rocket motor equipped with data-gathering instrumentation. The cost of each firing was $2 million, and all too 

often nozzles and instrumentation were consumed, making it

difficult to quantify test results. 

The 150 kW CW carbon dioxide laser uses electric 

discharge technology to provide a large, uniform flat-top beam

profile at a wavelength of 10.6 μm. As material is irradiated

during testing, its surface temperature can be precisioncontrolled within a range of 20° Celsius. For the solid rocket

motor nozzle composite, 22 tests were conducted on eight

material configurations to determine the relationship of test

sample ply lifting and charring to that seen in actual solid

rocket motor nozzles. The carefully controlled laser exposures,

as well as the combination of the range of specimens and the

visualization techniques, allowed thorough investigation of the

modes of material response.

In simulating the performance of a solid rocket motor, the

LHMEL 150 kW laser proved its value and versatility as a tool

for simulating the thermal conditions inside a solid rocket

motor nozzle during firing. The total cost of a two-day

LHMEL test was $18,000, a 99.1 percent reduction in cost

over the traditional nozzle testing method.

SUMMARY

A national asset, LHMEL offers unique material testing and

laser processing opportunities. From its inception, LHMEL has

provided reliable yet economical testing in an environment that

allows the user to focus on testing needs rather than on 

facility-related performance issues. While other laser test 

facilities exist in the DOD, none provides the scope of capabilities, the throughput-per-unit-dollar, or the ease of use that are

available at the LHMEL facility. LHMEL is truly a “one-stop”

shop for government agencies and industry who, like early

astronomers, are fascinated by space, but who are also savvy

enough to understand the importance of “getting bang for 

their buck.”

REFERENCES

John J. Eric, Robert J. Hull, Timothy R. Anderl, “Laser

Hardened Materials Evaluation Laboratory Simulates Space

Environment for Advanced Materials and Space Systems

Testing.” Materials Technology Highlights, Fall 2003, pp 2-3

John J. Eric, “Laser Hardened Materials Evaluation Laboratory

Performs Evaluation of Advanced Materials for Aerospace

Systems.” AFRL Technology Horizons, Vol. 1, No. 4, December

2002, pp 10-12

Mr. John J. Eric is a materials engineer at the Hardened Materials Branch of the AFRL Materials and Manufacturing

Directorate at Wright-Patterson AFB, OH. He is program manager of the Laser Materials Testing program with the

Anteon Corporation for the operation and materials interaction research efforts conducted at the Laser Hardened

Materials Evaluation Laboratory (LHMEL). 

Mr. Robert J. Hull currently works for Anteon Corporation as the Program Manager for the Air Force Research

Laboratory’s Laser Hardened Materials Evaluation Laboratory. Prior to 1997, Mr. Hull was an Air Force employee

serving as the Government Project Engineer for that same laboratory. Mr. Hull holds an MS in Materials

Engineering and a BS in Mechanical Engineering, both from the University of Dayton. 

Mr. Timothy R. Anderl is a technical writer and editor with Anteon Corporation, and has been working in support

of AFRL/ML for over five years. Mr. Anderl has a BS in Journalism from the E.W. Scripps School of Journalism at

Ohio University.
The AMPTIAC Quarterly, Volume 8, Number 1 35

INTRODUCTION

Material and process technologies are extremely critical to all

future space systems. These types of technologies are also

among a system’s most limiting factors because space-launch

applications require lightweight, affordable, durable, reliable,

high temperature materials. Additionally, these materials must

be easy to inspect to facilitate quick vehicle maintenance and

turnaround.

Because the current state-of-the-art in materials and process

technology limits the ability of the Air Force and the National

Aeronautics and Space Administration

(NASA) to have cheap and routine

access to space, revolutionary material

technologies are needed to ensure that

the US has a lasting presence in the

final frontier. It is understood by most

people involved in space technology

that revolutionary material technologies take a generation or

more to achieve. In response to this prospect, the space and

defense communities and industry are involving materials

experts in forming and implementing a national plan to ensure

that material technology advancement is kept at the forefront

of national research and development efforts.

INTEGRATED HIGH PAYOFF ROCKET PROPULSION

TECHNOLOGY PROGRAM

The Integrated High-Payoff Rocket Propulsion Technology

(IHPRPT) program, which began in 1995, is a collaborative

effort initiated by the Department of Defense (DOD), NASA,

and industry. The primary objective of IHPRPT is to double

rocket propulsion capability by 2010. IHPRPT’s Steering

Committee, is co-chaired by Mr. Andrew Culbertson from

DOD, and Dr. John Rogacki from NASA.

The IHPRPT approach is to develop and satisfy a set of firm,

challenging, but attainable propulsion technology goals that are

time-phased and measurable. The program is focused on looking at the overall propulsion needs and developing common

advanced component technologies that can be tailored to meet

specific operational requirements. In the past, technology

requirements for a system were typically determined up-front

during the program planning stage so that schedule risk could

be minimized. The problem with this approach is that the costs

associated with the technology development are often borne by

a single program or a single organization. Another problem

with this process is that the potential exists for duplication of

technical effort. 

Propulsion systems, whether for manned space flight, satellites, or tactical weapon delivery systems, have a large degree of commonality between them. The components

required to construct these military,

civil, or commercial rockets are similar,

providing an opportunity for joint

research and development of technologies that will meet the technical requirements of 

multiple applications. Coordination of technology developments between government agencies and outside organizations,

as occurs in the IHPRPT program, is a valuable tool for 

technology transition and transfer to meet a broad range of

needs in the government and commercial arenas.

The IHPRPT program is supported by a series of subplans

from industry designed to cooperatively ensure that the consolidated research and development goals are met. The

IHPRPT team is responsible for using this plan to establish 

the required program budgets and schedules. All resulting technology advancements will be available for use in new or existing military, civil, and commercial propulsion systems.

IHPRPT planning has identified the technologies that have the

highest potential benefit to systems and focuses government

and industry resources on developing them. From this plan,

multi-year development programs can proceed without interruption. To maximize technology transition opportunities,

time-phased technology demonstrations are conducted in 

configurations that verify the maturity of the technology and

Brian A. Stucke

IHPRPT Initiative Leader/Project Engineer

Timothy R. Anderl 

(Anteon Corporation)

Materials and Manufacturing Directorate

Air Force Research Laboratory

Wright-Patterson AFB, OH

Technology Development for Launch and Reentry Applications Section III

The primary objective of

IHPRPT is to double rocket

propulsion capability by

2010.
36 The AMPTIAC Quarterly, Volume 8, Number 1

show that it is ready to be applied to a specific system.

While IHPRPT members have a vested interest in meeting

IHPRPT goals, each organization derives different benefits

from the program. So, a major part of the IHPRPT planning

process involves analysis

of component level

needs as they relate to

specific goals and mission needs of the organizations and industries

involved. These specific

needs are incorporated into IHPRPT goals where appropriate.

For example, while NASA’s major thrust may be affordable,

reusable space transportation, the Air Force’s thrusts may

revolve around aircraft-like operations such as quicker alerts,

increased sorties, easier maintenance, and all-weather capability. Finding commonality of requirements in differing operational philosophies is the crux of the IHPRPT program.

Current baseline materials being used for propulsion components will not allow the IHPRPT program to achieve their

anticipated goals in liquid propulsion, boost and orbit transfer,

and tactical and spacecraft propulsion. IHPRPT-developed

materials will be required to operate in liquid oxygen and 

liquid hydrogen, high-temperature and high-pressure 

environments. 

Perhaps the most significant challenge facing IHPRPT 

participants involves weight, mass, and volume improvements

in boost and orbit transfer, spacecraft, and tactical propulsion

systems. Satisfying this critical challenge will require development of affordable materials and processes, and subsequently

supporting them with an adequate industrial base. 

Moreover, the liquid-rocket engine, solid fuel motor, and

spacecraft systems research and development arenas all come

with their own unique set of challenges. Improved liquidrocket engine systems will require higher temperature, higher

strength-to-weight materials, improved and more robust turbopump materials, oxidation resistant materials and coatings,

lightweight, high temperature materials for chambers and nozzles, and reduced part count for processing and manufacturing.

Solid fuel motor system developments will rely on adapting

polymeric materials for reduced weight components, lighter,

and lower cost insulation materials, and high-strength case

materials that will decrease component mass and volume.

Durable energy conversion devices for electric propulsion and

monopropellant-compatible materials are just a few of the challenges awaiting development teams in the spacecraft arena.

IHPRPT MATERIALS WORKING GROUP 

Because materials are at the core of space systems like the

Reusable Launch Vehicle (RLV), the IHPRPT Materials

Working Group (IMWG) was chartered by the IHPRPT

Steering Committee in February 1997. The IMWG enables

focused and unified materials development and collaboration

between NASA, DOD, and industry. Together, they have

developed joint requirements that are intended to satisfy 

multiple program needs. The IMWG partners share information and resources, proceeding from a joint roadmap. In short,

IMWG is one example of government-funded technology

development designed to achieve breakthroughs despite 

budgetary constraints and other seemingly insurmountable

limitations. 

Like the national IHPRPT steering committee, IMWG is

comprised of representatives from various NASA, DOD, and

industry organizations (such as Aerojet, Boeing Rocketdyne,

ARC, TRW, Pratt and Whitney and ATK/Thiokol). However,

while IHPRPT is focused on broad system challenges, IMWG’s

membership specializes in specific material and component

concerns. Currently co-chaired by Mr. Eric Becker from the Air

Force Research Laboratory Materials and Manufacturing

Directorate, and Dr. Raymond (Corky) Clinton of NASA’s

Marshall Space Flight Center, their primary activities include

developing material plans to meet IHPRPT’s liquid, solid, and

spacecraft propulsion goals. 

MATERIALS AND MANUFACTURING DIRECTORATE’S

CRITICAL ROLE

Representatives from the Air Force Research Laboratory’s

Materials and Manufacturing Directorate, who first received

funding for their participation in IHPRPT in 2001, are 

currently conducting research and development efforts

designed to address the requirements of several of the team’s

plans. These include materials and process development for

advanced solid rocket motors, rockets used for boost and

orbit transfer, and spacecraft propulsion and liquid boost and

orbit transfer requirements. 

For example, to meet goals in the boost and orbit transfer

arena, IMWG collaborators are developing management

devices* for liquid propellant, combustion and energy conversion devices, nozzles, throats, cases and insulation for solid

propulsion components. When successful, IHPRPT expects

reduced hardware costs, reduced support costs, improved thrust

to weight ratio, increased mean time between removals,

improved mission life and reusability, and reduced stage failure

rates. 

Similarly, in the spacecraft materials and processes arena 

are projects to develop Hydroxyl Ammonium Nitrate (HAN)

catalyst materials, laminated ceramic hall thruster chambers,

and carbon-carbon ion grid screens; as well as improve the

processing of high temperature reaction materials. Several 

benefits are expected to result from these efforts, including

reduced handling costs and safety risks related to hydrazine

propellants, larger spacecraft propulsion margins for extended

mission profiles, and extended operational lifetimes for

thrusters and grids.

From these plans, a consolidated list of materials needs was

developed, which has dictated the resulting materials development programs that the directorate has undertaken. The directorate currently manages 23 programs, which stretch across

four Core Technology Areas (CTAs), as well as a handful 

of delivery order contracts. Many of the directorate’s projects

consist of applied research efforts; such as assessing the applicability of emerging materials (like metal matrix composites or

nanomaterials) for aerospace systems. Working closely with

each individual program, experts from the directorate have

established cost and performance requirements for potential

materials. One such organization is the Air Force Research

Finding commonality 

of requirements in 

differing operational

philosophies is the crux

of the IHPRPT program.
The AMPTIAC Quarterly, Volume 8, Number 1 37

Laboratory’s Propulsion Directorate at Edwards AFB, CA,

which manages IHPRPT’s technology demonstrators. They 

are closely surveying challenges related to the representative

geometries and shapes of system components, and the processing requirements of these materials.

The directorate is addressing a number of IHPRPT goals.

Among these goals are developing advanced materials for

lightweight ducts, turbopumps, chambers, and nozzles, and

developing key applications for solid fuel motors; including

nozzle throats, insulation, tools and cases, thruster chambers

and grids, and reaction chambers. These materials and applications will enhance the overall performance of rockets by

increased usage temperatures, reducing the number of processing steps, lowering erosion rates, and developing lighter weight

components.

TRANSPIRATION COOLING RESEARCH 

For example, scientists and engineers at the Materials and

Manufacturing Directorate are developing and demonstrating

transpiration-cooling concepts and materials that could lead to

lightweight, high-efficiency combustion chambers for rocket

engines. Researchers expect the new technology to reduce the

weight of current, actively-cooled thrust chambers by 50%,

along with significantly reducing system complexity, part

count, cost, and coolant volume.

In a liquid-fueled rocket engine, such as the Space Shuttle

Main Engine, a fuel (in this case liquid hydrogen) and an 

oxidizer (liquid oxygen) are injected into a thrust chamber

where they mix and react. The fuel/oxidizer reaction products

are high temperature gases, which expand through a nozzle producing thrust (see Figure 4 in our article, Ceramic Materials for

Reusable Liquid Fueled Rocket Engine Combustion Devices).

Combustion takes place at temperatures in excess of 6000°F,

which is higher than the melting point of conventional engine

materials. Therefore, the chamber materials must be cooled 

by the continuous flow of a fluid that carries heat away from

the chamber walls. In this case the hydrogen fuel also serves as

the coolant fluid. 

The directorate has partnered with Ultramet, a small business in Pacoima, CA, to demonstrate the transpiration-cooling

concept. To date, numerous small cylinders, representing thrust

chambers, have been fabricated. At this stage in development

there are two main concepts being explored: a metallic-based

concept and a ceramic-based one. 

With IHPRPT funding and input from the directorate’s

Ceramics Branch, partners from Boeing-Rocketdyne began

conducting a series of tests on the small cylindrical thrust

chamber specimens. Coolant was pumped into the cylindrical

plenum where it then transpired through the porous inner

liner. The tests demonstrated that the amount of coolant that

flowed through the specimens would be adequate to cool an

actual thrust chamber. During the next year, researchers will

prepare a sub-scale combustion chamber for hot-fire testing at

the National Aeronautics and Space Administration (NASA)

Glenn Research Center’s Cell-22 rocket test facility. One chamber of each design, using the metal and ceramic concepts, was

scheduled to be delivered to NASA for testing in February

2004. More information on this subject is available in our 

article, Ceramic Materials for Reusable Liquid Fueled Rocket

Engine Combustion Devices, also published in this issue of the

AMPTIAC Quarterly.

ROCKET MOTOR CASE RESEARCH

Yet another example of an area where the directorate is enjoying tremendous success is in reducing the steps needed to 

produce a rocket motor case. Currently, a significant number of

cost-inefficient and labor-intensive process steps are involved 

in the manufacturing of a contemporary motor case. Some

examples of these labor and cost intensive process steps include

metal component machining, hand-spraying or automated

spraying of layers with single functionality, such as primers,

adhesives, barrier coatings and linings, and the process of

choosing and incorporating appropriate insulation materials. 

Under the management of the Nonmetallic Materials

Division’s Polymer Branch, five external contractual programs

focus on developing technologies to decrease labor and costs

and improve performance of rocket motor technology in consonance with the goals set forth by the IHPRPT Program

Steering Committee. The contractors, ATK/Thiokol, Aerojet

Corporation (formerly Atlantic Research Corporation), and

Pratt &Whitney Space Propulsion-CSD, are expected to drive

future design programs for larger scale development of tactical

and boost propulsion. 

One example of an approach to improve the rocket motor

case structure is to design a case in which materials perform 

several functions. Besides serving as a structural component,

materials may enhance the energy of a propellant grain, 

provide ablative properties, and insulate the rocket motor.

Engineers from the Polymer Branch recognize that multifunctional composites may provide those attractive physical

properties, weight savings, and reduction in production

process steps need to significantly improve rocket motor case

structures. Because composite materials are traditionally 

composed of a resin and fiber reinforcement, the combination

of a material with insulation properties, such as inherently

insulative fiber reinforcement, and a high performance resin

yields a multifunctional composite with both structural and

insulating properties. Thus, a combination of the materials

choices and the processing of the composite influence the 

ultimate performance and use of the multifunctional material

in the motor case.

Research and design collaborators from industry are currently developing small-scale prototype motors and will be studying the degradation of materials following exposure to the rocket motor environment during test firings at their facilities.

Engineers from the Polymer Branch expect to have selected a

stable case structure design by 2006 and anticipate beginning

live fire testing of the advance technology demonstrator structure at China Lake in the out-years of IHPRPT. 

ADDITIONAL MATERIALS AND MANUFACTURING

DIRECTORATE EFFORTS

The directorate has achieved a wide variety of successes in other

areas as well. For liquid rockets, engineers have completed permeability testing of chemical vapor-deposited (CVD) foam

chambers. Contributing to solid rocket propulsion, the direc-
38 The AMPTIAC Quarterly, Volume 8, Number 1

torate’s experts have successfully tested Plasma-Sprayed

Tungsten and Electroplated Tungsten (1 percent Rhenium) 1”

diameter nozzles during seven-second firings, demonstrating

minimal erosion. In addition, researchers have developed insulator materials surpassing a 400°F target for the self-insulating

case program. HAN propellant catalyst materials also passed a

1650°C tube test and met sintering targets.

The examples above represent just a small sample of the

activities undertaken by the IMWG, and by the Materials and

Manufacturing Directorate. Considering the complexity of

rocket engines and the multitude of components such as fuel

lines, ducts, valves, thrust chambers, throats, and nozzles that

comprise them, it becomes easy to recognize the critical 

importance of materials and process technology research and

development efforts to the IHPRPT program.

IMWG MEETINGS AND PEER REVIEW

IMWG meetings are conducted once or twice each year so that

members of the team can stay abreast of developments and

results on all fronts. During 2002, a peer review was conducted on efforts at the Materials and Manufacturing Directorate.

Experts from NASA, the Navy, and the Air Force reviewed each

of the directorate’s

IHPRPT-related projects and offered their

feedback and technical

expertise/advice.

Though feedback was

extremely positive, the

directorate’s scientists

and engineers capitalized on the information windfall provided

by experts from each visiting organization, using it to modify a

few of the parameters of their current projects and testing

plans.

The collaborative environment created by the IMWG’s

approach represents a significant benefit to the taxpayers, realized in terms of cost-savings and technical output. Simply put,

a single, unified effort drawing from all contributors yields

greater technical dividends than multiple parallel projects. For

example, the Materials and Manufacturing Directorate is conducting a project related to thermal barrier coatings, which

are, in this case, used as linings for chambers where the combustion of liquid hydrogen and oxygen takes place. The directorate’s goal is to develop coatings that will keep the inside wall

of the chamber cool enough that melting, or other problems

associated with transferring heat from the combustion chamber, will not affect the operability of a system. During the

1980s and 1990s, NASA conducted research and gathered

data on several potential thermal barrier coatings, which they

were able to provide to scientists and engineers at the directorate. This atmosphere of shared goals and collaboration

allowed the Air Force to avoid duplication of previous research

efforts, and the additional costs associated with undertaking

the efforts. In this way, they are able to direct their time, energy and budget towards potential solutions that push the technology envelope in new directions.

CONCLUSION

The key to IHPRPT’s success will be the ambitious shared

vision of its participants and advocates. By working together,

participants have already begun to see the benefits of sharing

information, sharing funding, joint planning, and joint

research and development. The program is already yielding significant improvements to the nation’s capability to move into

full-scale development of rocket propulsion systems with

improved performance, affordability, operability, reliability, and

maintainability. Work accomplished during IHPRPT is also

expected to set the baseline for future, cutting-edge projects

such as the Orbital Space Plane.

NOTES AND REFERENCES

* Propellant management devices ensure proper fuel flow from

the propellant tanks to the engines in many satellites. The

microgravity environment of space causes liquid fuels to randomly slosh around inside propellant tanks (which hampers

fuel flow) when not controlled with these devices.

Michael A. Stropki, Dr. Raymond J. Clinton, and Daniel A.

Cleyrat, “Beyond Coordination: Joint Planning and Program

Execution, the IHPRPT Materials Working Group.” The

AMPTIAC Newsletter, Vol. 4, No. 2, 2000, pp 1-5, 18

Mr. Brian A. Stucke currently manages the Air Force Research Laboratory Materials and Manufacturing

Directorate’s efforts supporting the Integrated High Payoff Rocket Propulsion Technology (IHPRPT) program. He 

has been an employee of the Materials and Manufacturing Directorate for 14 years. Brian holds degrees in 

computer science and business management from Wright State University.

Mr. Timothy R. Anderl is a technical writer and editor with Anteon Corporation, and has been working in support

of AFRL/ML for over five years. Mr. Anderl has a BS in Journalism from the E.W. Scripps School of Journalism at

Ohio University.

The collaborative 

environment represents

a significant benefit to

the taxpayer in terms 

of cost savings and 

technical output.
The AMPTIAC Quarterly, Volume 8, Number 1 39

The Air Force Research Laboratory’s Materials and Manufacturing Directorate (AFRL/ML) and their industry partners

are developing ceramic materials and manufacturing processes

for use in reusable liquid fueled booster rocket engine combustion devices (thrust chambers and nozzles) for next-generation,

reusable launch vehicles. Part of the Integrated High Payoff

Rocket Propulsion Technology (IHPRPT) program, the efforts

of the directorate include:

1. Developing continuous fiber reinforced ceramic matrix 

composites (CMCs) for actively cooled thrust chambers and

nozzles

2. Demonstrating the feasibility of a transpiration-cooled thrust

chamber

3. Evaluating ceramic matrix composites for radiation cooled

nozzles. 

The goal is to develop and demonstrate these new technologies so that they may be incorporated into future rocket

engines. Using lightweight ceramics has the potential to reduce

the weight of the combustion devices by up to 50%.

COMBUSTION DEVICE MATERIALS- HISTORICAL

PERSPECTIVE

In a liquid-fueled rocket engine, a fuel (such as hydrogen or

kerosene) and an oxidizer (such as liquid oxygen) are injected

into a thrust chamber where they mix and react. The fuel/

oxidizer reaction products are high temperature gases, which

expand through a bell-shaped nozzle to produce thrust. Gas

temperatures in the chamber may exceed 6000°F, while gas

temperatures in the nozzle may range from 3000°F - 5000°F.

These temperatures are too extreme for any conventional aerospace material; therefore engines must employ some type of

cooling scheme. Often the walls of combustion devices are constructed of tubes or channels. During operation, coolant is

pumped through the tubes or channels to keep materials within their temperature limitations. 

To illustrate, take the state-of-the-art in reusable rocket

engines, the Space Shuttle Main Engine (SSME). The SSME

thrust chamber consists of an inner copper liner with 390

milled cooling channels that run axially the length of the liner.

The cooling channels are closed out with a layer of electrodeposited nickel and then an outer structural jacket made of a

nickel-based superalloy is welded in place. During operation,

hydrogen coolant flows through the slotted channels in the

high conductivity copper liner to keep the component cool.

The SSME nozzle consists of 1080 tapered stainless steel tubes

that are brazed together and then brazed to an outer structural

jacket made of a nickel-based superalloy. During operation,

hydrogen coolant flows through the tubes to keep the nozzle

materials from exceeding their melting points.

Table 1 lists the materials and type of construction of 

numerous combustion devices, both historical and current. As

the table shows, the materials of choice (for all the engine manufacturers) for combustion devices in large liquid fueled rocket

engines have historically been stainless steels, nickel-based

superalloys, and copper alloys. These materials are selected for

their high strength and high thermal conductivity in order 

to cope with the stresses and extreme thermal environments 

of rocket engines. Since these alloys also have high densities 

(8-9 g/cm3), widespread reliance on them has traditionally

resulted in heavy engines.

Designers would like to reduce the weight of rocket

engines. A key performance criterion for engines is thrust-toweight ratio. Lighter engines and launch vehicles would allow

heavier payloads to be placed into orbit at a lower cost. One

path to lighter weight engines is replacement of conventional

high-density engine alloys with lightweight, high specific

strength ceramic composites. Two attractive candidates for

this application are carbon fiber reinforced silicon carbide

(C/SiC) and silicon carbide fiber reinforced silicon carbide

(SiC/SiC). These materials have low densities (2.0-2.4 g/cm3)

and high strengths that they maintain to relatively high temperatures (2400-3000°F).

Captain Steven Steel

Materials And Manufacturing Directorate

Air Force Research Laboratory

Wright-Patterson AFB, OH
40 The AMPTIAC Quarterly, Volume 8, Number 1

ACTIVELY COOLED CERAMIC STRUCTURES — 

PREVIOUS DEVELOPMENT

Preliminary work on actively cooled ceramic composite structures occurred under earlier government funded programs

such as the Actively Cooled Airframe Program for the

National Aerospace Plane (NASP) and the Linear Aerospike

Engine Nozzle Ramp for the X-33 reusable launch vehicle.

The goal of these programs was to demonstrate the feasibility

of using actively cooled lightweight ceramic composites for

hot structures such as

scramjet nozzles or the

aerospike engine nozzle

ramps. These programs

included proof-of-concept

demonstrations of small

actively cooled C/SiC

panels. Numerous different concepts were fabricated and tested. Several

of the concepts consisted

of a C/SiC composite

panel structure with metallic tubes embedded within,

through which the coolant flowed. C/SiC is an inherently

porous material as fabricated, so the metallic tubes were used

to provide impermeable coolant containment. 

An alternative concept to the hybrid metallic-ceramic 

panels described above was also demonstrated in the aerospike

engine nozzle ramp program. This was an all-C/SiC panel.

The all-composite approach has even greater potential for

weight savings than the hybrid approach. However, further

development is needed to make the C/SiC less permeable so

that it can contain coolant without leaking. Figure 1 shows

one of the all-C/SiC panels. The panel was fabricated by

Rockwell Scientific Company of Thousand Oaks, CA.

One of the unique attributes of this technology is the weaving process that allows the carbon fiber preform for an entire

tube-wall panel to be integrally woven as a single piece, as

opposed to single tubes being constructed separately and then

joined together as is the case for a metallic tube-wall nozzle.

This simplifies nozzle construction and results in a strong component with fewer joints. The feasibility of using structures like

the one shown in Figure 1 for nozzles was demonstrated in

tests. Technicians pumped water through the tubes while subjecting the panels to rocket exhaust gases. The panels were

exposed to rocket nozzle-like conditions for a total duration of

several minutes. 

CURRENT DEVELOPMENT OF ACTIVELY COOLED CMCS

Given these promising past results, the AFRL/ML Ceramics

Branch wanted to investigate whether the integrally woven

C/SiC technology could be adapted to the bell-shaped rocket

nozzle component geometry. AFRL/ML and partners from

Rockwell Scientific Company and Boeing-Rocketdyne have

begun to investigate whether the structures made previously for

the aerospike engine application are adaptable to the bellshaped nozzle application. Specific challenges include: weaving

tapered tubes rather than constant diameter tubes, forming the

curved panels needed to form the bell-nozzle, joining the separate panels together to form a complete nozzle, and developing

coatings or surface treatments to reduce the composite’s permeability. An additional challenge for a reusable nozzle is to

make the ceramic composite durable enough to withstand

hours of exposure to high temperature exhaust gases without

degrading. 

Results achieved to date show that the integrally woven

C/SiC composite technology is adaptable to the bell-shaped

nozzle geometry. Figure 2 shows a carbon fiber preform that

has been woven and formed into the shape of a bell nozzle segment. The preform is a single piece with the tube structure

integrally woven into the panel. Figure 3 shows a similar preform after it has been infiltrated with the silicon carbide matrix.

Sixteen segments like the one shown will be fabricated and then

joined together to form a complete subscale bell nozzle.

Techniques for joining the segments together are currently

under development.

An additional goal of the current research is to extend the

integrally woven tube panel technology to SiC/SiC composites.

SiC/SiC composite panels will be slightly heavier than C/SiC,

but they will be more oxidation resistant and may therefore be

more durable in a nozzle application. SiC/SiC composites may

also be less permeable than C/SiC and therefore provide better

coolant containment. Research thus far has shown that the

integrally woven tube panel technology is amenable to SiC/SiC

composites. The higher modulus of SiC fibers means that they

Figure 1. C/SiC Panel with Integrally

Woven Tubes.

Table 1. Materials and Construction Details for Past and Current Rocket Engines.

Engine Year Launch Vehicle Device Construction Material

V-2 German Army 1942 V-2 Chamber Double wall Low alloy steel Cr-Mn-V

LR91 Aerojet 1960 Titan Stage II Chamber Tube wall Stainless steel later Hastelloy X

RL-10 Pratt & Whitney 1963 Centaur Chamber Tube wall AISI 347 stainless steel

F-1 Rocketdyne 1967 Saturn V Chamber Tube wall Inconel X

SSME Rocketdyne 1981 Space Shuttle Chamber Channel wall CuAgZr liner Inconel 718 jacket

SSME Rocketdyne 1981 Space Shuttle Nozzle Tube wall A248 stainless steel tubes Inconel 718 jacket

Vulcain 2 Astrium 2003 Ariane 5 Chamber Channel wall CuAgZr liner Nickel jacket

Vulcain 2 Volvo 2003 Ariane 5 Nozzle Square tube wall Inco 600
The AMPTIAC Quarterly, Volume 8, Number 1 41

are more difficult to weave into intricate preforms. However, in

this program SiC fibers have been woven into straight tube

panel preforms like the configuration shown in Figure 1. The

tapered tube panel geometry remains to be demonstrated with

SiC fibers. Early tests of pressurized SiC/SiC tubes also verified

that they are indeed less permeable than the C/SiC.

In the on-going work, researchers will continue to develop

both the C/SiC and the SiC/SiC panels and nozzle segments.

Woven panels will be fabricated and subjected to high-heat flux

testing and thermal and mechanical fatigue testing. Sub-scale

actively cooled nozzles will be fabricated from C/SiC and/or

SiC/SiC and tested in an actual rocket environment at NASA

Glenn Research Center’s Cell 22 rocket test facility. 

TRANSPIRATION-COOLED THRUST CHAMBER

A second area of research is focused on an alternate and potentially more efficient form of active cooling, transpiration cooling. To investigate the feasibility of a transpiration-cooled

thrust chamber, AFRL/ML has partnered with Ultramet, a

small business in Pacoima, CA, and Boeing-Rocketdyne. This

innovative thrust chamber concept would do away with the

traditional cooling tubes and channels that have become 

standard in liquid rocket engines. A concept for a transpirationcooled thrust chamber is shown in Figure 4.

The transpiration-cooled thrust chamber concept is comprised of three major components; a porous inner liner, an

intermediate foam core and an outer jacket for structural support and cooling containment. During operation, coolant

would be pumped into the foam layer. The coolant would flow

through the open cell foam and be dispersed throughout the

foam layer. From the foam layer the coolant would then seep

through the millions of naturally occurring micropores of the

inner liner. The micropores distribute the coolant evenly and

efficiently over the combustion facing surface to keep it cool

and enable it to withstand exposure to the high temperature

combustion gases. 

Research thus far has focused on

the inner liner, particularly

achieving the right amount of

permeability in the liner to allow

adequate coolant flow. To date,

Ultramet has fabricated numerous small cylinders, representing

thrust chambers. The optimum

materials for the application have

not yet been determined; however the ceramic demonstration

articles consist of a lightweight

silicon carbide foam core with a

mixed molybdenum disilicidesilicon carbide inner liner. The

microporous inner liner is integrally bonded to the macroporous

foam support structure. The

cylinders were fabricated with 

different inner liner thicknesses to

determine the effect of liner

thickness on coolant flow. The

use of high temperature ceramics is appealing because they will

minimize the amount of coolant that must be provided to the

liner surface. The temperature limits of the materials are about

3000°F, so the optimum amount of coolant would keep the

inner surface of the liner to just below this temperature. 

Boeing-Rocketdyne tested the small cylindrical thrust chamber specimens in the fixture shown in Figure 5. Gases were

pumped into the cylindrical foam plenum where they then

transpired through the porous inner liner. The coolant flow was

visualized by applying a soap film to the inner liner and observing the bubble pattern as shown in Figure 6. The amount of

coolant that flowed through each cylinder was measured and

compared to computer models which were used to predict the

amount of coolant flow needed for a transpiration-cooled

thrust chamber. The tests demonstrated that the microporous

inner liner would allow an adequate amount of coolant flow to

cool the thrust chamber under predicted operating conditions.

The thickness of the liner can be varied to provide different

coolant flow rates. The cylindrical test specimens were also subjected to high pressures to verify the integrity of the foam-liner

Figure 4. Transpiration-Cooled Thrust Chamber Concept.

Figure 2. Carbon Fiber Preform for Bell Nozzle Segment.

Figure 3. C/SiC Bell Nozzle Segment.
bond. These “burst” tests verified that the foam-liner bond was

strong enough to withstand the pressures found in a rocket

engine.

On-going research includes further optimization of the inner

liner to provide the optimal amount of coolant flow for peak

engine efficiency. In the tests of the cylindrical specimens it was

noted that occasional large pores in the inner liner would cause

an excessive amount of cooling flow. In an actual engine this

would mean wasted fuel and reduced engine efficiency.

Additionally, the geometric complexity of the fabricated chambers will be increased from the current cylindrical shape to 

the “hour-glass” shape of an actual thrust chamber. A silicon

carbide foam core for a thruster is shown in Figure 7. The subscale combustion chamber will be hot-fire tested in an actual

rocket environment at the NASA Glenn Research Center’s

Cell-22 rocket test facility. The hot-fire test will determine if

the chamber can withstand the thermal stresses associated with

the rocket environment.

CERAMIC MATRIX COMPOSITES FOR

RADIATION–COOLED NOZZLES

An alternative approach to using actively cooled ceramics as

described above is to construct a component from a high temperature ceramic matrix composite that does not require any

active cooling. The component is cooled only by the radiation

of heat from its surface, thus eliminating the need for cooling

tubes or channels. Rocket engine construction would be simplified by the reduced cooling requirements and the elimination of tubes and manifolds. This approach is not feasible in the

hottest regions of combustion devices where temperatures

exceed the capability of prospective materials. However, one

could envision a rocket engine that consists of an actively

cooled thrust chamber and forward nozzle section that transition to an uncooled aft nozzle section at a downstream location

where gas temperatures are less extreme. 

It should be noted that uncooled nozzles have been used 

frequently in the past in the form of ablative nozzles. Ablative

nozzles have been used since the 1950s and are still a viable

option today. For example, Boeing-Rocketdyne selected an ablative nozzle, for its simplicity and technology readiness, for the

new RS-68 engine. Ablative nozzles are constructed of phenolic

plastics with various reinforcements that have included asbestos,

silica and carbon fibers. However, ablative nozzles must be constructed with thick walls to account for the charring and vaporization of the phenolics during use. They are therefore very

Figure 5. Permeability Testing of Cylindrical Specimens.

42 The AMPTIAC Quarterly, Volume 8, Number 1

Figure 6. Soap Bubbles Show Gas Flow Pattern.

Figure 7. Silicon Carbide Foam Thrust Chamber Core Shown on

Right; A Molybdenum Foam Core is Shown on Left.
The AMPTIAC Quarterly, Volume 8, Number 1 43

heavy. Furthermore, due to the loss of material during operation, ablative nozzles are not suitable for reusable applications. 

To identify and evaluate potentially suitable high temperature composites for the uncooled nozzle application,

AFRL/ML has partnered with Boeing-Rocketdyne and

Rockwell Scientific Company. For this application materials

must be able to withstand temperatures of 3000°F-4000°F.

They must also resist degradation when exposed to a rocket

engine’s combustion environment; including temperature,

environment, and stress; all for durations on the order of 10

hours (For a booster engine 8.5 minutes per launch with a

desired life of 40-100 launches).

In this research effort a variety of high temperature materials are being evaluated. Candidates materials include C/SiC,

SiC/SiC, C/C (carbon fiber reinforced carbon matrix) and

C/ZrC (carbon fiber reinforced zirconium carbide matrix).

Left unprotected, these materials would react with oxygen

and/or steam present in rocket exhaust gases and would quickly degrade. Therefore the materials must incorporate a protective coating, oxidation inhibitors in the matrix, or some other

form of protection. The materials screened in this effort were

selected for their innovative oxidation protection schemes.

Candidate materials were fabricated by seven different 

material suppliers and are currently being subjected to a battery

of screening tests. Key among these are stress tests, both static

and cyclic, on small specimens in an air atmosphere in a high

temperature furnace. So far, the best performing material can

support a load of about 20 ksi for 8 hours at 3000°F. Material

degradation mechanisms have been analyzed using microscopy

of post-test specimens. Ways to improve the materials have

been identified and will guide the next round of material 

selection and synthesis. The goal is to develop an improved oxidation resistant material and then demonstrate it in an actual

rocket nozzle environment.

The application of ceramics and ceramic matrix composites

to rocket engines is still in its infancy. This article has

described three on-going research efforts to apply ceramics to

combustion devices of liquid fueled rocket engines. Replacing

metallic materials that have been in use for over 50 years in

rocket engines will not happen overnight. This research is laying the groundwork for future application of these materials to

help create the next generation of lightweight rocket engines.

Captain Steven Steel received a BS in Mechanical Engineering from the Ohio State University and an MS in

Material Science from the Air Force Institute of Technology. He recently retired from the Air Force after 20 years of

service. The last four years of his Air Force career were spent working in the Ceramics Branch of the Air Force

Research Laboratory’s Materials and Manufacturing Directorate at Wright-Patterson Air Force Base.

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44 The AMPTIAC Quarterly, Volume 8, Number 1

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The AMPTIAC Quarterly, Volume 8, Number 1 45

BACKGROUND

Thermal protection systems (TPS) are materials that are used to

protect launch vehicles, reentry vehicles, and spacecraft from

exposure to high temperatures experienced during operation.

The tiles on the Space Shuttle are an example of TPS but there

are numerous other kinds as well. Some of these materials ablate

(degrade away) during operation while other classes are either

insulative or actively cooled by some sort of fluid flowing

through interior passageways. A final type of TPS, known as

“hot structure” is made from materials that can withstand the

direct exposure to high temperatures without experiencing any

degradation. These materials possess the strength and temperature resistance to function in high-temperature applications but

they do nothing to insulate and protect underlying components.

TPS materials enable high temperature operation of launch,

reentry, and spacecraft systems. They are either single-use or

reusable materials that require significant maintenance following use. The current state-of-the-art in TPS technologies does

not provide the characteristics and properties required to enable

advanced space systems. Planners for both the Air Force and

NASA envision future reusable launch vehicles (RLVs) that are

far easier to maintain than the Space Shuttle. Advanced TPS is

not needed just for RLVs. Expendable launch vehicles and

hypersonic weapon systems also require TPS to protect critical

systems during use.

One of the critical factors limiting the availability of

advanced TPS today is the difficulty in maturing new technologies. For example, the fundamental science of designing a

material, and the related sciences of processing the material to

an application and understanding its reaction to an environment provides a major challenge to TPS system development.

Breakthroughs cannot be scheduled, and they may take a generation to achieve; however, new technologies must be available, or at least their availability must be assured, during the

early stages of engineering planning and design for new systems

or capabilities. Development times can take very long and there

are many fundamental limitations imposed by the basic laws of

physics. The only way to effectively mitigate these factors and

to accelerate development is to begin work on advanced TPS

during the earliest stages of the development cycle when future

systems are being planned.

A TEAM APPROACH TO TECHNOLOGY DEVELOPMENT

The Air Force and NASA recognize both the limitations of 

current TPS technologies and the difficulties associated with

developing replacement materials. To strategically address

future needs, a joint DOD/NASA/industry team has been

working for nearly five years to develop requirements, technology forecasts, and program roadmaps to facilitate development

planning for the technologies needed for our future systems.

Modeled after the Integrated High Payoff Rocket Propulsion

Technology (IHPRPT) materials working group [IMWG,

reported in Volume 4, Number 2 (2000) of the AMPTIAC

Quarterly] the TPS team has structured an approach whereby

the DOD, NASA, and industry will cooperatively develop a

plan addressing and prioritizing TPS technology approaches

necessary for future space systems and military weaponry.

In 1999 the first TPS materials workshop was held in

Dayton, OH. This was a government-only meeting intended to

develop a better understanding of the TPS materials research

efforts sponsored by the Air Force and NASA. An additional

goal was to use the information gained at this meeting to help

develop the Air Force Research Laboratory’s Materials and

Manufacturing Directorate’s investment strategy for future 

TPS materials development. Attendees agreed that a partnership between various government agencies and industry could

provide the focus to develop the needed technologies.

A number of recommendations were made at the first meeting. One of the most important was to hold a continuing series

of workshops to help ensure coordination between DOD and

NASA research efforts. Also recommended was the development of a national TPS plan to help ensure coordination and

advocacy for research efforts, thus maximizing cooperation and

minimizing the potential for duplicative programs. The final

recommendation was to use the planning process developed

under the IHPRPT program.

A government and industry supported TPS working group

has been formed to provide the guidance needed to carefully

plan future research and development efforts. Led by a small

steering committee of Air Force and NASA representatives, 

this group contains a good cross-section of the users and developers of TPS technologies. Table 1 displays the organizations

supporting this group.

David H. Rose, Director

AMPTIAC

Rome, NY 

Daniel A. Cleyrat

(Anteon Corporation)

Materials and Manufacturing Directorate

Air Force Research Laboratory

Wright-Patterson AFB, OH
Following the initial meeting, a series of workshops have been

used to develop and refine a National TPS plan and strategy.

Figure 1 illustrates the purpose and outcomes from each of

these workshops. The single most important outcome has

been the development of an integrated plan and strategy that

the Air Force and NASA are now using to prioritize their TPS

research investments. Because of the coordination resulting

from the working groups, the Air Force and NASA can now

maximize their use of resources, provide opportunities for

cooperative programs, and maintain world-class leadership in

TPS technology.

The challenge in developing an integrated plan that both

the Air Force and NASA could live with was in determining

common requirements. The DOD is interested in “operationally responsive” space lift and hypersonic weapons. To be

an effective tool for promoting and defending national political objectives, the Air Force requires space systems that perform similar to aircraft in that they can operate in all weather

conditions and also be rapidly brought back to flight-ready

status after returning from space. While reduction of TPS

maintenance is a goal for NASA, the agency doesn’t have the

goal to develop aircraft-like operations. NASA’s primary

emphasis concerns low-cost and crew safety for manned spaceflight. Likewise, since commercial spacecraft typically rely

upon expendable rockets, they don’t have the requirement for

reusability. DOD also requires expendable systems so there is

a commonality of need here. The seemingly small differences

between DOD and NASA RLV requirements do have an

impact on the technologies that must be developed for future

systems. The TPS materials working group has been successful

in finding enough commonality of requirements to assign 

priorities to future RLV and expendable launch system technology development efforts.

The National TPS materials plan addresses the strategy 

needed to maintain the momentum to advance the state-of-theart in TPS technologies. One of the most important attributes

46 The AMPTIAC Quarterly, Volume 8, Number 1

Table 1. Primary Organizations Supporting 

the TPS Working Group.

Government Industry

AFRL/Materials and Manufacturing 

Directorate* Boeing

AFRL/Space Vehicles Directorate Lockheed Martin

NASA/Ames Research Center* Goodrich

NASA/Goddard Space Flight Center Northrop Grumman

NASA/Johnson Space Center Orbital Sciences Inc.

NASA/Langley Research Center* Oceaneering

NASA/Kennedy Space Center Triton Systems

NASA/Marshall Space Flight Center Hitco

NASA/Dryden Flight Research Center Ultramet

Carderock Division, Naval Surface Fiber Materials Inc. (FMI)

Warfare Center

Department of Energy/Oak Ridge Composite Optics Inc.

National Laboratory

Army/Redstone Alliant Technologies (ATK)

*co-chairs

Figure 1. History of TPS Workshops.

TPS Workshops History

Purpose Outcome

• Obtain initial government input on materials requirements

• Develop understanding on current programs and future direction

• Need for industry involvement

• Desire for forum to continue

• Need for “National” coordination

• Obtain initial industry input for materials requirements 

& planning need

• Visit shuttle TPS processing facility

• Recommendation to continue as organized forum with 

scheduled meetings

• Recommendation to develop “National” TPS Material Plan

• Develop an initial recommendation for Air Force TPS 

materials investment

• Coordinate team strategy for “first iteration” of National 

TPS Materials Plan: “Tool for the Future”

• Initial AF recommendation

• 1st draft of TPS Materials Plan

• Formalization of working group

• Review future plans and requirements for AF and NASA

hypersonic systems

• Work the National TPS Plan

• 2nd draft of TPS Materials Plan

• Established Co-Chair Leads: Eric Becker, AFRL/ML &

Sylvia Johnson, NASA ARC

• Review NASA 2nd Gen & AF Projects

• Continue to work the National TPS Materials Plan

• 3rd draft of TPS Plan

• Durability Panel

• Validation and update of current TPS Plan

• Development of a recommended TPS program “portfolio” & Roadmaps for DOD and NASA NAI/NGLT

• Review of ground and flight test opportunities

• Expanded scope of Working Group & Plan for TPS

• Developed coordinated roadmaps, level four technology charts and portfolio

• Identified/coordinated flight testy opportunities

TPS Sidebar meetings at 02 NSMMS & NASA/AF Airframe Conference

Workshop V

DOD, NASA, 

DOE, Industry

Workshop VI

DOD, NASA, 

DOE, Industry

Jul 99

Jan 00

Aug 00

Apr 01

Oct 01

Mar 03

Workshop IV

DOD, NASA, 

DOE, Industry

Workshop III

DOD, NASA, 

DOE, Industry

Workshop II

DOD, NASA, 

DOE, Industry

Workshop I

DOD, NASA, DOE 

(gov only)
of the plan is that it addresses the critical path (highest 

priority) required to develop the needed technologies. It is

instrumental in helping to determine technology areas that need

investment and it is certainly helping to avoid duplication of

efforts. It also is providing an opportunity for each participating

organization to leverage the efforts of the other participants,

thus accelerating the development of new technologies.

Initially, the working group was primarily focused on 

materials technologies, but over the past several years the scope

has broadened. Following numerous discussions and meetings

with TPS developers and with the joint DOD/NASA National

Aerospace Initiative (NAI) team and NASA’s Next Generation

Launch Technology (NGLT) team, the TPS Materials Working

Group’s steering committee recommended the expansion of the

scope of the working group and current materials plan to incorporate structures, integration, and testing, thus becoming a more

complete “TPS Working Group.” There is also a recommendation to include TPS “operations” at a point in the near future.

APPROACH FOR DEVELOPING PLAN

The first thing that had to be done was to look at current and

projected programs and from this analysis determine individual

performance requirements by system. During the timeframe

that the plan was being developed, the Air Force had a number

of proposed systems requiring advanced TPS; including the

Space Operations Vehicle (SOV), Space Maneuvering Vehicle

(SMV), Common Aero Vehicle (CAV), hypersonic vehicles,

and engines. Similarly NASA projected a number of systems

requiring advanced TPS; including Shuttle upgrades, RLVs,

Orbital Space Plane (OSP), hypersonic vehicles, and engines.

Performance requirements for these systems were analyzed and

compared so that common TPS solutions could be identified.

In normal circumstances when materials engineers are developing new materials or evolving existing ones, there is no need

to consider the full spectrum of how the materials will be

applied and maintained. In most cases the new materials are

similar enough to existing materials that expanding the research

to incorporate the additional requirements is not needed.

Anyone who has followed NASA’s experience with the Space

Shuttle knows that the TPS currently being used, while effective in operation, is fragile and requires costly maintenance

between flights.

A building block approach to developing new TPS is being

used. Included is a time-phased strategy for incrementally

advancing TPS performance to meet the needs of near-term,

mid-term, and far-term projected systems. For each phase of

the plan a number of elements must be considered. Table 2 outlines those issues that must be addressed in order to ensure that

the new system can operate with an acceptable level of risk.

Specific performance goals at the materials level are dictated

by system level performance requirements. The National

Aerospace Initiative (NAI) is a joint DOD/NASA effort to

improve access to space. NAI has three objectives:

• Space Access: Develop and demonstrate technologies that

enable responsive, safe, reliable, and affordable access to

and from space.

• High Speed/Hypersonics: Develop and demonstrate 

technologies that enable air-breathing hypersonic flight.

• Space Technology: Develop and demonstrate technologies

that enable transformational and responsive in-space 

capabilities.

NAI has established goals and a phased timeline for new

technology development. Figure 2 outlines these goals.

TPS MATERIALS TECHNOLOGY REQUIREMENTS

If we are to achieve more responsive, reliable, and affordable

space transportation vehicles then major advances in TPS

materials technology is required for each of the critical path

components. Included in these components are leading edges,

control surfaces, acreage TPS, and seals. Figure 3 illustrates

these critical components. Each of the critical path items will

be discussed more fully later but the one thing to note is that

The AMPTIAC Quarterly, Volume 8, Number 1 47

Table 2. Plan Elements.

Identify vehicle components requiring TPS

Identify performance requirements associated with 

component objectives

Identify candidate material options for the TPS component

Identify candidate materials’ current properties

Identify issues/risks associated with candidate materials

Identify current materials programs associated with component

Identify recommended materials initiatives associated 

with component

Figure 2. National Aerospace Initiative Goals.

Figure 3. TPS Components.

Phase III – 2020

• Crew Safety – 0.99999

• Vehicle Loss Rate – 1/5000

• Cost/Mission – $15M

• 95% System Availability due

to Weather

• Call-Up Time – 6 Hours

• Alert Hold – 10 Days

• Launch Rate – 3 Launches in

24 Hours for 7 Consecutive

Days

Far Term

Phase II – 2015

• Crew Safety – 0.9998

• Vehicle Loss Rate – 1/2000

• Cost/Mission – $30M

• 85% System Availability due

to Weather

• Call-Up Time – 2 Days

• Alert Hold – 48 Hours

• Launch Rate – 1 Launch in 24

Hours for 7 Consecutive Days

Mid Term

Phase I – 2009

• Crew Safety – 0.998

• Vehicle Loss Rate – 1/1000

• Cost/Mission – $75M

• 70% System Availability

due to Weather

• Call-Up Time – 7 Days

• Alert Hold – 12 Hours

• Launch Rate – 1 Launch in

7 Days Sustained

Mid Term

Baseline

Shuttle, EELV

2009 2015 2020

• Leading Edges

Blunt

Sharp

• Control Surfaces

Hot Structure

Insulated Structure

• Acreage Surfaces

Windward

Leeward

• Seals

Control Surfaces/Penetrations

TPS Panels

Environmental Pressure Seals
since each of the critical path components experience different

maximum sustained temperatures and heat flux, then a variety

of TPS materials is required.

Leading Edges: Concepts for advanced leading edge materials

included passive and active systems. Passive systems are

designed to withstand the rigors of operation with no need for

removing the excess heat. Cooled leading edges have either heat

pipes buried within them or involve some other mechanism

that allows for the heat flux to be absorbed and carried away.

There are three approaches being considered for passive 

leading edge materials: refractory composites, functionally

graded/hybrid materials, and ultra-high temperaturecapable materials. Refractory composites including coated 

carbon-carbon (C/C) and carbon fiber reinforced silicon 

carbide (C/SiC) show promise for multi-cycle use in lightly

loaded structures exposed to maximum temperatures in the

range of 3000-3200°F. Functionally graded and hybrid concepts utilizing novel applications of existing materials such as

carbon and/or ceramic foams, tiles, and other high temperature materials are also being considered. The “holy grail” is

ultra-high temperature materials that can withstand the same

conditions as refractory composites with the exception that the

maximum temperature capability would be in excess of

3600°F. Materials being considered include monolithic ceramics and fiber reinforced composites. Actively cooled leading

edges comprised of composite, ceramic, or metallic materials

are also being considered.

Control Surfaces: Materials used for constructing control 

surfaces will experience different conditions than leading edges

so a different approach is needed. Control surfaces are loadbearing and this major difference alone requires additional

considerations during design. Also, control surfaces cannot 

be cooled, so the materials used to construct them must be

capable of handling the heat flux experienced during flight.

There are two types of control surfaces: hot structures and

insulated structures. Both of these approaches must be able to

withstand multi-cycle use in an oxidizing environment. The

materials used to construct hot structures are designed to withstand the heat flux during operation as well as the mechanical

loads. For hot structures, three types of materials; metals,

C/SiC and C/C, are being considered. The major differences

between these are the maximum temperatures that they can

withstand during operation. Metallic materials can be used up

to approximately 1500°F. C/SiC can experience a maximum

temperature in the range of 2800°F while C/C is expected to be

good through approximately 3200°F. Conversely, insulated

structures consist of an insulative material over the top of the

load bearing structure. Different insulation materials could be

selectively used to address the thermal/mechanical requirements at various control surface locations.

Acreage: Acreage TPS refers to the materials used to cover the

large portions of the external structure of the vehicle. The most

recognizable example of this type of material is the tiles used on

the Space Shuttle. In an RLV like the Space Shuttle, or in the

advanced concepts being considered, there are different environmental conditions imposed on the aeroshell depending

upon the orientation of the vehicle during reentry. For example, since the bottom of the craft (the “windward” side) is more

directly exposed to airflow during reentry, it experiences

increased heating as compared to the “leeward” surfaces found

on the top of the craft. Again using the Shuttle as an example,

the tiles on the windward side are black while the leeward side

tiles are white. The differences in these two different tiles relates

to their ability to withstand the heat flux during reentry.

High temperature metal and intermetallic acreage TPS are

being considered as a means to facilitate multi-use, low maintenance, and rapid turn around times for future applications.

Specific types of materials being considered include oxide 

dispersion-stabilized (ODS) alloys, superalloys and aluminides.

These materials have an expected temperature range of 1000-

2500°F (depending upon the exact material used). Ceramic

matrix composites than can withstand temperatures on the

order of 2000-2600°F are also being considered. Likewise,

C/SiC shows promise for demonstrating long-life multi-use

capability in the 2600-3000°F range.

48 The AMPTIAC Quarterly, Volume 8, Number 1

Table 3. Technology Issues to Consider during TPS Development. 

Fabrication and Processing

Cost

Durability

All weather

Strength at temperature

Thermal shock

Cycling resistance

Cycles to failure/repair/replacement

Handling

No shape change

Low to no water absorption

Abrasion/erosion

Impact Resistance

Low speed

Hypervelocity

Health Monitoring

Operability 

Repair, rework, inspection

Field or depot repair

Accessibility: Installation/removal

Minimal processing for turnaround

Cross range max L/D

Abort capable

Cost

Performance

Temperature (single and multi-use)

Pressure

Time duration

Structural loads

Heat flux

Vibro-acoustic

Material design allowables

Density

Life

Refurbishment

Thermal Properties

Conductivity

Emissivity

Less Variability on Material Properties
The AMPTIAC Quarterly, Volume 8, Number 1 49

Seals: Seal materials are used to protect the area where two 

different types of structures are joined. An example of a seal

location is where the flight control surfaces are joined to the

wing. This is a very important location because there is movement between adjacent structures. Seals must be resistant to a

variety of different temperatures depending upon the usage and

location on the craft. Pressure seals have much lower requirements and must only withstand temperatures around 800°F.

Light-weight, all-weather, multi-cycle life materials are required

for advanced control surface and TPS seals. Control surface

seals that abut hot control surfaces must withstand a temperature in the range of 2000-2500°F. Interpanel seals for highly

reusable TPS are expected to experience even greater temperatures in the range of 2500°F or higher.

MUCH WORK REMAINS TO BE DONE

A tremendous amount of work has to be done to bring a new

material to a technology readiness level sufficient for safe operation. This is especially true for thermal protection systems

since they have to operate in such harsh environmental conditions. Issues such as joining requirements and anti-oxidative

coatings by themselves can be showstoppers. There are numerous other technical issues that must be addressed before a new

TPS technology can be considered mature. Table 3 outlines

some of the issues that must be addressed during a technology

development program.

CONCLUSIONS

Successfully developing advanced spacecraft and hypersonic

systems can only be achieved if new thermal protection systems

are developed. Existing TPS have fundamental limitations that

preclude them from being effectively used in many of the proposed applications. (It is for this reason that these materials are

viewed as enabling technologies.)

The Air Force and NASA has built upon the successful collaboration that started with the IHPRPT program’s materials

working group and assembled a parallel group consisting of

government agencies and industry to focus upon TPS. This

group has worked with development teams planning for future

systems within the two agencies, and from the knowledge

gained, have assembled a consolidated list that prioritizes

needed technologies. This list has subsequently been used to

develop programmatic roadmaps to help plan for the needed

TPS research efforts.

The existing TPS plan has been through a number of iterations, starting with the first draft in August of 2000. Planners

involved with the National Aerospace Initiative have used the

plan extensively. With the focus now provided, these planners

can have some confidence that the development of the required

TPS technologies will be accelerated and new materials will be

available at the time they are needed for future systems.

Advanced TPS are an essential group of technologies that

must be available if the Air Force and NASA are to develop the

systems envisioned. Over the years there have been numerous

instances where materials technology was developed concurrently with a new system. However, when considering that TPS

must survive high temperature and often ionizing environments, developing new materials to survive even more extreme

conditions will be very difficult. Only if the technology development is separated from the system development and the

proper resources, focus, and planning are applied can the development of advanced thermal protection system technologies be

accelerated and available when needed for future systems. The

TPS working group has made excellent progress in this area. By

the consolidated efforts of the Air Force and NASA, we can

now develop the enabling technologies that will allow us to

reduce the cost and increase the rate at which our country can

gain access to space for military and commercial purposes.

Mr. David H. Rose has been with AMPTIAC since its beginning in 1996, and has been its Director since June of

1998. Prior to AMPTIAC, he was an officer in the US Air Force, retiring in 1996. During his Air Force career he

worked in a variety of technical and acquisition roles including five years at the Materials and Manufacturing

Directorate where his research interests included micromechanical analyses of composite materials. He has a BS in

Mechanical Engineering from the University of Washington, an MS in Mechanical Engineering and has completed his PhD coursework in Materials Engineering from the University of Dayton.

Mr. Daniel A. Cleyrat had a 24-year distinguished military career as an Air Force officer, where he served in Missile

and Space operations and other capacities. He is currently with Anteon Corporation in Dayton, OH and is

assigned as the principal engineer to the Space Office of AFRL/ML. Mr. Cleyrat also serves as the Technical Chair

for the National Space and Missile Materials Symposium.
IMAPS International High Temperature

Electronics Conference (HiTEC 2004)

05/17/04 - 05/20/04

Sante Fe, NM

Contact: Intl. Microelectronics Pkg. Society

611 2nd St, NE

Washington, DC 20002

Phone: 202.548.4001

Fax: 202.548.6115

Email: imaps@imaps.org

Web Link: www.imaps.org/flipchip

SAMPE 2004

05/18/04 - 05/20/04

Long Beach, CA

Contact: SAMPE IBO

PO Box 2459

Covina, CA 91722

Phone: 800.562.7360 x 601

Fax: 626.332.8751

Web Link: www.sampe.org

10th DoD ElectroMagnetic 

Windows Symposium

05/18/04 - 05/20/04

Norfolk, VA

Contact: Curtis Martin, Code 645 

Naval Surface Warfare Ctr., Carderock Div.

9500 MacArthur Blvd

West Bethesda, MD 20817

Phone: 301.227.4501 

Fax: 301.227.4732

Email: martinca@nswccd.navy.mil

EASTEC APEX

05/25/04 - 05/27/04

West Springfield, MA

Contact: Society of Manufacturing Engineers

One SME Dr 

PO Box 930

Dearborn, MI 48121-0930 

Phone: 800.733.3976 

Fax: 313.425.3407

Web Link: www.sme.org

Great Lakes Photonics Symposium

06/07/04 - 06/11/04

Cleveland, OH

Contact: Anteon Corporation

5100 Springfield St, Suite 509

Dayton, OH 45431

Phone: 937.254.7950

Fax: 937.253.2296

Web Link: www.usasymposium.com

2004 Nat’l Space & Missile Matls Symp.

06/21/04 - 06/25/04

Seattle, WA

Contact: M. Kubal, Anteon Corporation

5100 Springfield St, Suite 509

Dayton, OH 45431

Phone: 937.254.7950

Fax: 937.253.2296

Email: mkubal@anteon.com

Web Link: www.usasymposium.com

2004 IEEE Nuclear & Space Radiation

Effects Conf. (NSREC 2004)

07/19/04 - 07/23/04

Atlanta, GA

Contact: D. M. Fleetwood, Vanderbilt Univ.

PO Box 92, Station B

Nashville, TN 37235

Phone: 615.322.1507

Fax: 615.343.6702

Email: dan.fleetword@vanderbilt.edu

Web Link: http://www.nsrec.com/

11th International Conf. on

Composites/Nano Eng’ing (ICCE-11)

08/08/04 - 08/14/04

Hilton-Head, SC

Contact: Dr. David Hui, U. of New Orleans

Dept. of Mech. Engng.

New Orleans, LA 70148

Phone: 504.280.6652

Fax: 504.280.5539

Email: dxhme@uno.edu

Web: www.uno.edu/~engr/composites/html

10th Int’l Symp. on Superalloys

09/19/04 - 09/23/04

Champion, PA

Contact: TMS Meeting Services

184 Thorn Hill Rd

Warrendale, PA 15086

Phone: 724.776.9000 ext 253

Fax: 724.776.3770

Email: mtgserv@tms.org

Web Link: http://doc.tms.org

MS&T ‘04 (Matls. Science & Tech. 2004)

09/26/04 - 09/30/04

New Orleans, LA

Contact: TMS Meeting Services

184 Thorn Hill Rd

Warrendale, PA 15086

Phone: 724.776.9000 ext 243

Fax: 724.776.3770

Email: mtgserv@tms.org

Web Link: cms.tms.org

SPACE 2004

09/28/04 - 09/30/04

San Diego, CA

Contact: Customer Service, AIAA

1801 Alexander Bell Dr.

Reston, VA 20191

Phone: 703.264.7500

Fax: 703.264.7657

Web Link: www.aiaa.org

FABTECH Int’l Forming & Fabricating -

Stamping - Tube & Pipe - Welding

10/26/04 - 10/28/04

Cleveland, OH

Contact: Society of Manufacturing Engineers

One SME Dr, PO Box 930

Dearborn, MI 48121-0930 

Phone: 800.733.3976 

Fax: 313.425.3407

Web Link: www.sme.com

CANEUS 2004–Conference 

on Micro-Nano Technologies

11/01/04 - 11/05/04

Monterey, CA

1801 Alexander Bell Dr

Reston, VA 20191

Phone: 703.264.7500

Fax: 703.264.7657

Web Link: www.aiaa.org

IMAPS 2004 Conference

11/14/04 - 11/18/04

Long Beach, CA

Contact: Intl. Microelectronics Pkg. Soc.

611 2nd St, NE

Washington, DC 20002

Phone: 202.548.4001

Fax: 202.548.6115

Email: imaps@imaps.org

Web Link: www.imaps.org/imaps2004

2004 Insensitive Munitions &

Energetic Materials Technology Symp.

11/15/04 - 11/17/04

San Francisco, CA

Contact: Christina Buck, NDIA

2111 Wilson Blvd, Suite 400 

Arlington, VA 22201

Phone: 703.247.2587

Fax: 703.522.1885

Email: cbuck@ndia.org

2004 AIAA Missile Sciences Conference

11/16/04 - 11/18/04

Monterey, CA

1801 Alexander Bell Dr

Reston, VA 20191

Phone: 703.264.7500

Fax: 703.264.7657

Web Link: www.aiaa.org

Aging Aircraft 2005

01/31/05 - 02/03/05

Palm Springs, CA

Contact: R. Loeslein

NAVAIR, Aging Aircraft Program

Bldg 2185, Suite 2100 C4

22347 Cedar Point Rd, Unit 6

Patuxent River, MD 20670-1161

Phone: 301.342.2179

Fax: 301.342.2248

Email: loesleinGF@navair.navy.mil

Web Link: www.agingaircraft.utcdayton.com

2005 TMS Annual Meeting &

Exhibition

02/13/05 - 02/17/05

San Francisco, CA

Contact: TMS Meeting Services

184 Thorn Hill Rd

Warrendale, PA 15086

Phone: 724.776.9000 ext 243

Fax: 724.776.3770

Email: mtgserv@tms.org

Web Link: http://doc.tms.org

50 The AMPTIAC Quarterly, Volume 8, Number 1

Mark Your Calendar
The AMPTIAC Quarterly, Volume 8, Number 1 51

OVERVIEW

The cost of placing objects into low earth orbit has been 

estimated to be roughly $10,000 per pound. Because of this

fact weight is closely managed during system design. Given that

a launch vehicle’s lift capability is fixed, designers have historically paid close attention to the weight of satellite subsystems.

Reducing the weight of support subsystems allows for additional payload mass that translates into more transponders 

on communication satellites or perhaps larger aperture optics

and sensors for surveillance satellites. 

The backbone of the satellite, referred to as its primary

structure, accounts for 10% to 20% of the total dry mass. This 

estimate does not account for station keeping propellant or 

secondary structures such as antennas, solar array supports, or

component boxes. Thermal management components account

for an additional 2% to 5% of dry mass.[1] As mentioned

above, reducing the weight of these systems provides opportunities for increasing satellite performance. For over 40 years

the Nonmetallic Materials Division’s Structural Materials

Branch (AFRL/MLBC) has led the way by developing composite materials to reduce the weight of satellite structures and 

subsystems. 

WHY COMPOSITES

Composite materials are one of eleven technologies that the

Materials and Manufacturing Directorate currently identifies as

a core technology area. However, emphasis on composites by

the Air Force is not a recent development. In 1963 the USAF

initiated Project Forecast with the purpose of identifying the

most promising breakthrough technologies to enable future air

and space systems. The Air Force Materials Laboratory (now

AFRL’S Materials and Manufacturing Directorate) alerted the

Project Forecast team to the weight savings that could result

from using low density polymer matrix composites made from

high modulus, high strength fibers. The laboratory staff pointed out that boron fiber-reinforced polymer composites had the

strength of steel and a stiffness greater than aluminum, while at

the same time possessing a density less than beryllium. As a

result, Project Forecast identified the development of advanced

composites as one of the technologies that showed tremendous

potential for enabling increased systems performance. 

After composites were identified as a critical technology, the

Materials Laboratory (ML) began work on a long-range plan

for developing this new class of materials. The result was a

white paper that laid out a $360 million program to bring the

promising technology from coupon-level specimens to full aircraft production. In 1965, this white paper was submitted to

Air Force officials for consideration of funding. ML was soon

thereafter provided with $6 million and a mandate to demonstrate success before further funding would be provided. The

Advanced Filaments and Composites Division of the Materials

Laboratory was established to begin this work and research into

composites has been ongoing ever since.

Since that time, composites have been at the forefront of aircraft and spacecraft research and development. The attributes

of composites are well known and include high specific

strength (strength per unit weight) and stiffness, corrosion and

fatigue resistance, tailorable conductivities, controlled thermal

expansion, and the ability to be processed into complex shapes.

As structural materials, composites offer numerous system level

benefits over their predecessors. They are lightweight, which is

critical for any space platform since less structural weight allows

for more fuel or payload. The physical, mechanical, electrical,

and thermal properties of composites are highly tailorable,

which can also afford them multifunctionality (materials that

perform more than one critical function simultaneously, such

as bear loads, serve as an antenna, and be a thermal radiator).

Moreover, composites may be designed with directional properties, enabling the creation of novel structures not possible

with conventional materials. 

Over the past 40 years, there has been a natural progression

of composite technology development and applications, and

today composites are viewed as an accepted structural material. Composites have become the first widely used engineered 

Jeffrey D. Guthrie

AMPTIAC

Rome, NY

Tia Benson Tolle, Chief

Structural Materials Branch, Materials and Manufacturing Directorate

Air Force Research Laboratory

Wright-Patterson AFB, OH

David H. Rose, Director

AMPTIAC

Rome, NY

Technology Development for Satellite Applications Section IV
52 The AMPTIAC Quarterly, Volume 8, Number 1

material and are paving the way for other designed materials,

including those based upon nanotechnology and computational materials science. Most in the field agree that for many

aerospace applications there’s a lot of unexplored design space

and undiscovered benefits that composites may yet realize. In

fact, the National Research Council’s 2003 study, Materials

Research to meet 21st Century Defense Needs, concluded,

“Composites offer the greatest opportunities for significant

advances in material design and function”.[2]

DESIGN ISSUES WITH COMPOSITES FOR SPACE

APPLICATIONS

Launch Environment

One of the functions of a satellite’s primary structure is to first

help the spacecraft survive the launch environment. Engine

thrust during launch and ascent imposes a steady state acceleration to the spacecraft. This acceleration persists until a stage or

booster burns out. When the next stage ignites, the acceleration

resumes. Together, both the steady state and the transient acceleration events impose inertial loads on the spacecraft. These

loads are expressed in multiples of the weight at sea level or Gs.

For the Shuttle, this is limited to a relatively light 3.5G load,

while the major expendable launch vehicles (Titan, Atlas,

Delta) experience a max 4G to 6G load. Acoustic vibration

from the launch vehicle’s engines coupled with aerodynamically induced vibration provides an additional critical source of

loading on satellite structures. During ascent into orbit, several

explosive events occur to separate the stages, boosters, and payloads. These explosive events generate shock loading into structures. While shock loads attenuate quickly, their peak value can

be extremely high. The combined effect of all of the loading

mechanisms discussed above must be considered during the

design process.

The satellite must also be rigid enough during launch to

avoid interacting with the payload fairing on the launch vehicle (The payload fairing protects the satellite from wind 

loading and contamination). Structures that use materials with

a high modulus of elasticity have a high natural frequency, 

and thus small deflections. Thus the gap between the payload

fairing and the satellite can be made smaller when “stiffer”

structures are used.[1]

Orbital Environment

The major function of a satellite’s structure in the orbital environment is to keep the sensors and payload components in the

same relative position to each other and to protect internal

components from the space environment. Consequently, structures must be very stiff to resist dimensional changes. Higher

conductivity materials also reduce thermal distortion by

isothermalizing the structure from both shadows the vehicle

casts upon itself and from cycling in and out of the Earth’s

shadow. A low (near zero) coefficient of thermal expansion

(CTE) can help reduce thermal distortion as well. 

A material’s strength is not as important once the satellite is

in orbit since the loads are minimal. The most common loads

are generated by thermal cycling as the spacecraft moves in and

out of the Earth’s shadow. Turning the satellite (slewing) and

moving it to change its orbit can also generate small loads.

However, stiffness is again more critical for these events. A

slewing or moving maneuver causes the satellite to vibrate. The

size of the displacement and the natural frequency of vibration

are proportional to the structural material’s modulus of elasticity. Thus, after a slewing or moving event, pointing accuracy

can be established more quickly with high modulus materials

than with lower modulus materials. 

The design of composite satellite structures must account for

the effect of the space environment. Many polymers outgas

when subjected to the vacuum of space. Outgassing can result

from desorption of water bound to the external surface of a

polymer or contained within its interior. It is also caused when

other volatile materials contained within the polymer escape or

when the polymer decomposes due to its interaction with the

space environment. A portion of the outgassed species will

deposit on the satellite and form a film that darkens its surface.

This can seriously degrade the performance of optical, sensor,

solar cell, and thermal control materials. Materials that have a

total mass loss of less than 1% and a collected volatile condensable material of less than 0.1% are typically required. Some

satellite systems may have even more stringent requirements.[3]

Desorption of water from some polymers can also reduce their

dimensional stability. In these cases distortions due to desorption can be forty times greater than the thermal distortions of

a high modulus composite. Analogous to CTE, polymer composites have a characteristic coefficient of moisture expansion

(CME) that is used as a design parameter for satellite structures.

Composites are also susceptible to dimensional changes from

thermal cycling, which can result in matrix microcracking to

relieve internal stresses. This phenomenon can occur when the

CTE of the matrix and the fiber are substantially different. The

problem can be compounded by particle radiation that can

embrittle the matrix in some composites. Further compounding this issue is that higher modulus fibers desired for their

dimensional stability have higher CTE mismatches with the

polymer matrix. Since outgassing is a diffusion-controlled

process, microcracks can increase the rate of moisture desorption – further affecting dimensional stability.

For low Earth orbits, such as experienced by the Shuttle or

the International Space Station, atomic oxygen (AO) can also

degrade the performance of composite materials over time. The

carbon bonds of many organic polymer materials are susceptible to the impact energies of AO. Interaction of the polymer

with AO can in some cases severely erode the material.

The Structural Materials Branch strives to exploit the 

benefits of composite materials while addressing the issues of

operating in the space environment. The focus of this work

has been to provide designers with materials that enable

dimensionally stable structures, and other multifunctional

applications including structures that also help manage heat

generated by mission electronics.

RECENT EFFORTS

During the late 1970’s, the Advanced Composite Satellite

Equipment Support Module Study sponsored by the Air Force

demonstrated that composites could provide a 15-20% weight

reduction over satellite construction practices used at the time.

Since then AFRL/MLBC has made tremendous strides in
The AMPTIAC Quarterly, Volume 8, Number 1 53

advancing the development of composites for both structural

and thermal applications. Some of the recent advances in polymer matrix and carbon-carbon composites are highlighted in

the following sections.

Dimensionally Stable Structures

Polymer Matrix Composites (PMCs): During the early 1990s

AFRL/MLBC teamed with Oak Ridge National Laboratory to

explore the potential of thermoplastic composites for Strategic

Defense Initiative (SDI) applications. Thermoplastics were

investigated because they exhibit low moisture pick-up (low

CME), low outgassing, and ease of formability as compared to

state-of-the-art graphite/epoxy composites. Composites made

from very high modulus fibers and both Polyetheretherketone

(PEEK) and Polysulfone polymers were analyzed to determine

how they would perform in the space environment. The laminates tested were designed and fabricated so that they had near

zero CTEs. Even though thermoplastic composites possessed

the performance characteristics desired by satellite designers,

the higher processing temperature requirements and increased

fabrication costs limited their acceptance.

Despite the fact that thermoplastic matrix materials weren’t

embraced by the design community, the use of high modulus

fibers for dimensionally stable structures was well received. This

posed a problem since Dupont had discontinued production of

their E series pitch fibers, leaving offshore sources such as the

Japanese as the only producers of these materials. A Defense

Production Act Title III program was initiated to develop a new

domestic source of pitch-based carbon fibers that possessed the

high modulus and high thermal conductivity needed for satellite applications. P100 and P120 fibers were the first fibers to

be developed under this program. Later efforts improved upon

fiber processing methods to yield a higher strength pitch fiber

(P100X HTS) that had better handleability without sacrificing

the higher modulus. Large quantities of pitch fiber were purchased and made widely available to DOD users with the 

condition that any data generated through its usage would be

reported back to the Air Force.

In 1993 a workshop was held at the Institute for Defense

Analysis (IDA) to examine the development needs for a new

class of composites based upon cyanate ester resins.

AFRL/MLBC along with other government organizations and

industry participated in this workshop. Cyanate ester-based

composites are processed like epoxies, which makes them

affordable, yet they possess the attractive performance characteristics of thermoplastics, including low moisture outgassing

and low microcracking. The critical factor identified at this

workshop was the need for a statistical properties database. In

response, the Structural Materials Branch generated a designoriented database containing fully characterized mechanical

properties pertaining to ultra thin cyanate ester composite laminates made from P100X HTS fibers. The database helped

accelerate the acceptance of this new class of materials, which

are now the material of choice for many space applications. 

Carbon-Carbon Composites: Many of the space systems being

developed for SDI during the early 1990’s were very large and

required extreme pointing accuracies. Coupled with this

requirement was the need to withstand directed energy and

nuclear threats. AFRL/MLBC initiated five programs to 

develop carbon-carbon (C-C) composites for survivable and

dimensionally stable structures. AFRL/MLBC provided the

technologies needed by the Propulsion Directorate

(AFRL/PR) for a system that was being developed. In the

course of these efforts, thin wall high modulus 2-D C-C structures were demonstrated. Thermal control coatings for space

structures were also demonstrated at this time.

C-C composites are desirable for spacecraft usage since their

dimensional stabilities are better than other structural materials

such as graphite/epoxy composites. This is because the carbon

matrix will not absorb moisture, hence the composite doesn’t

possess a CME. Initially C-C structures were not economically

feasible because of their high cost of fabrication. However, the

aforementioned programs developed several approaches that

dramatically reduced processing costs. C-C also possesses

tremendous survivability attributes that were demonstrated in

both underground nuclear and directed energy tests. Even

when the threat of directed energy or nuclear attack on space

structures declined, C-C development continued based upon

interest in its extreme dimensional stability. 

Using the success of the early SDI programs, AFRL/MLBC

initiated a long-range initiative, known as the Lightweight

Dimensionally Stable Structures (LDSS) Program, to develop

the support technologies needed to transition C-C technology

into structural applications. This program employed a building

block approach and examined the joining technologies needed

to integrate C-C materials into prototype subassemblies. To

facilitate design, an extensive material property database was

developed for C-C composites. This database addressed C-C

containing many of the same leading fiber systems used for

space structures made from polymer matrix composites.

Applications for several dimensionally stable components were

investigated including a metering truss structure and an optical

bench. 

Thermal Management

Many space structures perform both structural and thermal

management roles. For instance, a spacecraft’s bus structure

usually supports its electronic boxes. These boxes generate heat

that must eventually make its way through the structure to a

radiator surface. Additionally, with the trend towards higher

performance payloads, denser packaging of electronics, and

longer mission lives, lightweight materials are needed to dissipate heat energy more effectively than materials used in the

past. The development of both high modulus/high thermal

conductivity PMCs and C-Cs led to realization that these

materials offered substantial weight savings over aluminum

thermal management materials. Working with industry as a

team, AFRL/MLBC ensured many successful transitions of

PMCs and C-Cs for thermal management applications.

Polymer Matrix Composites: In response to major advances in

high thermal conductivity fiber and matrix systems seen in the

mid 1990’s, AFRL/MLBC initiated a pair of programs that

examined whether thermal and structural functions of components could be combined to dramatically reduce weight. These
efforts used high thermal conductivity fibers (K13C and

K1100) in a cyanate ester matrix to develop and demonstrate

affordable, thermally efficient lightweight polymeric composite

materials. A prototype radiator panel that provided the

required thermal management functions while withstanding

launch loads was developed under the first program. This 

multifunctional panel directly supported nine electronic boxes

and weighed one third less than a comparable aluminum

design. The second program also utilized thermostructural

composites to fabricate a number of flight components for

three different spacecraft built at Lockheed-Martin’s

Astronautics Operation. The components demonstrated

include K1100/954-3 radiator fins, a composite battery panel

(see Figure 1), and a thermally isolated radiator. Each of the

components successfully survived the launch environments,

and their in-space performance was consistent with predicted

performance.[4, 5] 

Carbon-Carbon Composites: Like the PMC efforts described

above, C-C composites are excellent materials for thermal

applications. PMC and C-C composites can possess equivalent

thermal conductivity in the two dimensional plane of the laminate since this property is driven by the thermal conductivity

of graphite fibers used. An advantage of C-C is they typically have through-the-thickness thermal conductivities

that are a couple of orders of magnitude larger than

most PMCs. This is because through-the-thickness conductivity is a property that is highly dependent on the

matrix material and carbon conducts heat much better

than polymers. A disadvantage of C-C is they are more

costly and have mechanical properties that are inferior

to PMCs. Thus, C-C technology is traditionally applied

in areas where higher conductance is required. 

The previously discussed LDSS program examined

joining technologies needed to transition C-C into thermal

management applications. Thermal components examined

under LDSS include thermally conductive battery cases and

high temperature radiator fins. As a result of this effort, the

materials developed under this program transitioned to components on various flight vehicles. A low modulus, high thermal conductivity C-C thermal doubler flew on the Mars

Global Surveyor along with a C-C engine shield (see Figure 2).

High modulus, high thermal conductivity C-C composites

were used as heat sinks on a Titan instrument subsystem. In

addition, the Deep Space 1 spacecraft used a P120/carbon thermal doubler on an optical bench and under multichip modules.

The remarkable thing to consider is that high modulus, high

conductivity C-C materials were successfully transitioned to

flight in just over 10 years from the start of development.[6]

The Carbon-Carbon Spacecraft Radiator Program

(CSRP) was a collaborative program between the Air

Force, Navy, NASA, and industry that successfully demonstrated a revolutionary radiator panel

launched on Earth Orbiter 1 (EO1). This

radiator was built to demonstrate that CC could significantly reduce thermal control costs associated with satellites

54 The AMPTIAC Quarterly, Volume 8, Number 1

Figure 1. Schematic of Thermal/Structural Battery Panel.

Figure 2. Carbon-Carbon Thermal Doubler on Mar Global

Surveyor.

Figure 3. C-C Radiator on EO1.
The AMPTIAC Quarterly, Volume 8, Number 1 55

and possibly extend their operational lives (see Figure 3). For

additional information on CSRP, please refer to The AMPTIAC Quarterly Volume 2, Number 3.[7] 

AFRL/MLBC also developed and demonstrated C-C for

heat sinks used in electronic packaging applications and as

materials for battery case applications. The effort to develop

electronic heat sinks demonstrated an 18% improvement in

thermal performance with a 12% decrease in weight over traditional aluminum components. The battery development initiative demonstrated a carbon-carbon multifunctional structure

for sodium sulfur (NaS) batteries. Sodium sulfur batteries offer

the promise of much higher energy densities than conventional batteries; however, they operate at higher temperatures

(roughly 400°C). To further improve the energy density of the

batteries being developed, high thermal conductivity 2-D C-C

materials for battery cases were developed. These materials are

desirable because they are lightweight and chemically inert to

the electrolytes. C-C composites not only substantially reduced

the weight of the battery case, but their high thermal conductivity eliminated the need for a thermal sleeve that helped

isothermalize the battery. 

CURRENT PROGRAMS

Dimensionally Stable Structures

Graphitic Foams: Most previous efforts at developing dimensionally stable space structures focused upon tailoring the properties of 2-D laminated composites. One of the earliest

attempts at developing a novel material to replace composites

for these applications occurred during the early 1990’s when an

SDI program produced carbon fiber-reinforced carbon foam.

While this material exhibited interesting properties it could not

be reproduced in usable quantities. Even though the program

wasn’t entirely successful, it did demonstrate the potential

advantages of using a controlled 3-D open cellular graphitic

structure with tailored properties in the cell ligaments. Carbon

foam is attractive as a structural material because it provides the

potential for near net shape parts with near isotropic properties.

The initial efforts described above ultimately led to a new class

of engineered microcellular materials for structural and thermal

applications. Several manufacturers now produce graphitic

foam products. 

If the mechanical properties can be improved, ultra-lightweight, ultra-stiff carbon foam may be the perfect material for

large space mirror structures requiring extreme dimensional 

stability. The Structural Materials Branch is developing refined

foam processing methods with the goal of modifying the

microstructure and thus enabling improved compressive, tensile, and shear strengths for mirror and other applications.

Additionally, an effort is ongoing to develop the ideal mirror

surface. For more information please see the article addressing

mirror technologies contained in this issue of the AMPTIAC

Quarterly.

Thermal Management

Thermal management in the space environment is more critical than in all other environments for two reasons: material

density is much more important in space (due to launch costs)

and convective cooling is not available. High conductivity

materials and active cooling mechanisms employing radiators

are used to ‘wick’ heat away from where it’s produced, say for

example an electronic box, and allow it to be expelled into

space. The Structural Materials Branch is continuing to develop new technologies such as composite heat pipes and highly

crystalline graphite that will provide improved thermal management capabilities as compared to current state-of-the-art

approaches. 

Composite Heat Pipes: Previous AFRL/MLBC efforts have

been highly successful at developing composite materials and

joining techniques for thermal management applications.

However, one of the highest payoff applications for composite

materials, radiators, remains unresolved. As satellite power

increases, the size of radiators required to emit heat into space

through radiative cooling must also increase. To operate at

maximum efficiency, heat must be uniformly distributed across

a radiator surface. Heat pipes are attached to a radiator panel’s

surface using a compliant room temperature vulcanizing

(RTV) silicone adhesive. The physical contact between the

panel and the pipes provides the mechanism for heat to be dispersed into the panel. A critical factor to consider is that the

radiator’s thermal conductivity must be designed so that the

heat will flow away from the heat pipes, thus ensuring a uniform temperature distribution across the radiator.

There are several design factors that must be considered

when developing satellite cooling systems. The advantage of

using higher conductivity materials is that fewer heat pipes are

needed to distribute the heat, which dramatically reduces system weight and complexity. Radiators are normally fabricated

by attaching heat pipes to one of its surfaces, which imposes the

need for a good thermal joint. However, since two different

materials, a composite panel and metallic pipes are used, any

change in temperature can cause a distortion of the joint due to

differences in the CTE of the two materials. If the distortion

causes the heat pipe to pull away from the radiator panel, then

heat transfer in that region of the radiator becomes impaired.

An additional factor to consider is that a cooling fluid, normally ammonia, is used to transfer heat from where it is generated

to the radiator. This requires the pipe to be impermeable and

resistant to attack by ammonia.

To further reduce the weight and improve the efficiency of

satellite thermal control systems, the Structural Materials

Branch has three programs underway that are developing the

technologies that will enable advanced heat pipes. Two of these

programs are examining a variety of composite heat pipe solutions, including metal/composite hybrids. The third program is

investigating heat pipes that are embedded within the radiator

panel as well as low CTE metallic pipes. 

Some of the composite heat pipe technologies that are being

examined include braided carbon/cyanate ester heat pipes,

small diameter impermeable glassy carbon in novel radiator

designs, pultruded and carbonized carbon/phenolic tubes,

both copper and aluminum plated C-C tubes, and a hybrid CC infiltrated with molten aluminum. Some of these approaches employ grooves machined into the tube wall, known as wick

structures, that help return condensed ammonia to an evaporator. Figures 4 and 5 display some of the technologies being
56 The AMPTIAC Quarterly, Volume 8, Number 1

considered. As mentioned previously, permeation of the

ammonia through the tube is a serious consideration.

Nanocarbon particles in the matrix are being investigated as a

mechanism that may help decrease permeability. 

A partnership between ML and the major spacecraft prime

contractors has been established to help resolve key shop floor

issues that are common across the industry. This partnership is

also examining the issue of the heat pipe/radiator thermal joint.

In this program, heat pipes directly embedded into the radiator

panel itself are being examined. Other alternatives to resolving

the thermal distortion problem, including the use of a low CTE

alloy, Invar, for heat pipes are being considered.

Highly Crystalline Graphite: The thermal conductivities of

graphitic carbon fibers as measured along the axis of the fiber

can exceed that of copper by over three fold. In a composite

laminate the amount of high conductivity fiber in any one

direction is limited by the composite’s construction details and

also its fiber volume. As a result, the in-plane thermal conductivity will not approach the axial thermal conductivity of the

fiber itself. Planar sheets of graphite offer the extremely high

thermal conductivities in two directions that graphitic fibers

offer in one direction. However, most of the mechanical properties of graphite are poor in comparison to composites. 

The Air Force is striving to exploit the thermal conductivity

of planar graphite by using hybrid approaches to improve

mechanical robustness. The two approaches currently being

considered both involve

encapsulating graphite in

either C-C, polymer matrix

composites, or aluminum.

The Space Vehicles

Directorate (AFRL/VS) is

examining encapsulated

annealed pyrolytic graphite

(APG), which is manufactured using a chemical vapor

deposition (CVD) process. Similarly the Structural Materials

Branch is examining encapsulated ThermalGraph®. This product is made using a unique process where raw pitch fibers

(before graphitization) are compacted and then graphitized.

The process creates a homogeneous graphitic material that

depending upon construction details can be highly conductive

in one dimension (if the precursor fibers are unidirectionally

oriented before compaction) or has good in-plane thermal conductivity (if chopped precursor fibers are randomly oriented).

Both AFRL/VS and AFRL/MLBC are examining application

of these materials in lightly loaded applications such a thermal

doublers or thermal planes (heat sinks for electronics). Figure 6

shows a heat sink made from ThermalGraph® encapsulated by

aluminum. 

Graphitic Foams: Graphitic foams promise tremendous weight

savings over conventional materials

used in thermal

management applications. These

materials exhibit

thermal conductivities similar to

many aluminum

alloys at 1/5th the

density. Current

research efforts are

investigating the

use of graphitic

foam for various

thermal management components

such as high performance heat sinks used in electronics and as

a core material to replace aluminum honeycomb in sandwich

panels for structural radiators. The thermal and weight advantages of graphitic foams are so immense that the Air Force is

also investigating the use of graphitic foams in heat exchangers

for aircraft applications. Figure 7 shows the cellular microstructure of high-density foams.

Nanotailored Composites

Advanced capabilities needed to enable the next generation of

satellites will come from multifunctional composites.

Nanotechnology offers the potential to combine functions without sacrificing structural performance that designers expect. For

satellite applications, AFRL/MLBC is developing nanotailored

materials with both thermal and electrical conductivity, which

can be processed using traditional composite fabrication techFigure 4. Hybrid C-C Heat Pipe. Figure 5. Al Plated C-C with Wick.

Figure 6. ThermalGraph Heat sink.

Figure 7: High Density Graphitic Foam.
The AMPTIAC Quarterly, Volume 8, Number 1 57

niques. The following sections describe current efforts.

Nanolayered Silicates: Nanotechnology may provide the ultimate means of tailoring a material’s microstructure and properties to yield dimensionally stable structures. Conventional

polymer matrix composites have many strong attributes but

there are some issues that limit their use for dimensionally stable structures. Matrix microcracking and CTE mismatch

between the fiber and matrix are two factors that can reduce a

material’s dimensional stability when subjected to temperature

cycling while in orbit. Incorporating nanoscale reinforcements

into the matrix resins used to fabricate PMCs may improve

these tendencies and subsequently enable wider use of PMCs

for critical applications.

AFRL/MLBC is investigating the potential for nanotailoring

of a composite matrix to preclude or reduce microcracking and

CTE mismatch

issues. This is not a

trivial endeavor and

as such, some basic

research must be

undertaken to investigate critical issues.

Current efforts are

examining the effect

of the nanoconstituents on polymer

resin viscosity. Nanoscale reinforcement

and interface properties are also being

characterized. In a

related effort, AFRL/

MLBC is developing modeling techniques to address the enhancement of composite properties resulting from nanoscale

reinforcements. The ultimate goal is to apply this basic knowledge obtained under these programs to both dimensionally 

stable structures and possibly also to thermal management

materials. 

The nanoscale reinforcements currently being considered are

hydrated aluminosilicate platelets roughly one nanometer

thick. AFRL has been investigating nanocomposites based

upon these materials for quite some time[8, 9, 10]. This new

initiative builds upon the earlier work and is concentrating

upon using smaller quantities of nanoscale particles as a complement to a composite’s traditional fiber reinforcement. The

challenge of this research is to uniformly disperse the particles

and prevent them from agglomerating. While other organizations have overcome these problems in thermoplastic resins, it

appears that AFRL/MLBC will be the first to achieve a uniform

distribution in thermoset materials key to aerospace applications. Figure 8 shows a photomicrograph of nanolayered silicates (dark phase) distributed in epoxy. 

Nanofibers: The Structural Materials Branch is currently investigating the use of vapor grown carbon nanofibers as a means

of improving the thermal and electrical conductivity of polymers. These efforts revolve around the use of Pyrograph-III®, a

carbon nanofiber produced by Applied Sciences Corporation,

which has an axial thermal conductivity roughly five times

greater than copper. Pyrograph-III®is a multi-walled nanotube

(or nanofiber) that are produced with random diameters that

range from 40 to 200 nanometers. The aspect ratios of these

fibers are quite large (see Figure 9).

Pyrograph-III® nanofibers are commercially available in

large quantities. When dispersed in a polymer matrix, the

small diameters and large aspect

ratios of these fibers enable them to

contact each other and create a network at a relatively small volume

fractions of fibers (<5%). This network greatly increases the thermal

and electrical conductivity of the

base polymer. Current research has

focused upon making the nanocomposites discussed above. However,

similar to efforts involving the hydrated aluminosilicate

platelets discussed earlier, Pyrograph-III® dispersed in small

quantities within matrix resins may enable an order of 

magnitude increase in the through-the-thickness thermal 

conductivity of polymer matrix composites.[11] Similarly,

small additions of these nanofibers to adhesives may greatly

increase thermal conductivity.

SUMMARY

The AFRL Materials Directorate’s Structural Materials Branch

has a long history of developing thermal and structural materials for use in satellites. Following this legacy the current diverse

and robust research program is helping develop technologies

that will enable the next generation of satellite structures.

Although not discussed in this article, AFRL/MLBC is similarly focused on developing technologies for space launch applications. Materials being developed include composites for cryogenic tanks and C-C thermal protection materials. An example

of this work is the ongoing examination of nanolayered silicates

to reduce permeability of cryogenic fuel tanks.

Building upon the successes mentioned above,

AFRL/MLBC plans to continue development of hybrid 

materials, multifunctional structures, and nanocomposites. A

new program starting Fall 2004 will examine advanced multifunctional structures. To-date most multifunctional materials

have combined thermal and structural functions. This new

program will examine whether other functionalities such as

electro-static discharge (ESD) protection or electromagnetic

interference (EMI) shielding can be built into the structure.

Adding additional functionalities creates opportunities for 

further reducing the weight of space structures, thus enabling

increased mission capabilities through the use of additional

sensors, transponders, and other systems.

REFERENCES

[1] James R. Wertz, Wiley J. Larsen, “Space Mission Analysis

and Design”, Microcosm Press, 1999

[2] UTC, “History of the Air Force Material Laboratory

Volume II, Part II Landmark Technology Programs 1960-

1969”, March 2000

Figure 8. Nanolayered Silicates in Epoxy.

Figure 9. Multi-walled

Carbon Nanotubes.
Mr. Jeffrey D. Guthrie has been a senior materials engineer for AMPTIAC since 1998. Prior to joining AMPTIAC,

Mr. Guthrie served as a liaison between the Air Force Philips Laboratory (space laboratory, now AFRL/VS and

AFRL/DE) and the Materials Directorate. Mr. Guthrie also spent six years working on Carbon-Carbon Composites

for space applications for the Structural Materials Branch (AFRL/MLBC). Mr. Guthrie holds BS and MS degrees in

Metallurgical Engineering from Montana College of Mineral Science and Technology. His MS thesis focused on

processing ceramics.

Ms. Tia Benson Tolle is the Acting Chief of the Materials and Manufacturing Directorate’s Structural Materials

Branch, and the Core Technology Area Leader for Organic Matrix Composites. She is also an International

Executive Cabinet Officer for the Society for the Advancement of Material and Process Engineering (SAMPE). Her

experience spans nearly two decades and includes 15 years at the Air Force Research Laboratory (AFRL) working

with space structural materials. Prior to joining AFRL, she worked as an instructor at the NASA Johnson Space

Center’s Space Shuttle Flight Training Division. She holds a BS in Mechanical Engineering from the University of

Washington. She also holds an MS in Materials Engineering and is a PhD Candidate in Materials Engineering at

the University of Dayton.

Mr. David H. Rose has been with AMPTIAC since its beginning in 1996, and has been its Director since June of

1998. Prior to AMPTIAC, he was an officer in the US Air Force, retiring in 1996. During his Air Force career he

worked in a variety of technical and acquisition roles including five years at the Materials and Manufacturing

Directorate where his research interests included micromechanical analyses of composite materials. He has a 

BS in Mechanical Engineering from the University of Washington, an MS in Mechanical Engineering and has 

completed his PhD coursework in Materials Engineering from the University of Dayton.

58 The AMPTIAC Quarterly, Volume 8, Number 1

[3] Edward M. Silverman, “Space Environmental Effects on

Spacecraft: LEO Materials Selection Guide”, NASA Contractor

Report 4661, August 1995

[4] Stephen S. Tan, Robert D. Torczyner, Kenneth A. Barczi,

“Thermal Structural Materials Solutions for Spacecraft”, Final

Report, September 1998

[5] Suraj Rawal, David Martin, Jerry Draper, “Thermal

Structural Materials Solutions for Spacecraft”, Final Report,

May 2001

[6] Suraj Rawal, Frank Kustas, “Lightweight Dimensionally

Stable Structures”, Final Report July 1999

[7] Elizabeth Shinn, Howard Maahs, “Carbon-Carbon

Radiator Panel Improves Thermal Control on Satellites”, 

AMPTIAC Quarterly Volume 2, Number 3, Fall 1998

[8] Vaia, Richard, “Polymer Nanocomposites Open a New

Dimension for Plastics and Composites”, AMPTIAC Quarterly

Volume 6, Number 1 

[9] C. Chen, D. Curlis, “Organoclay-Aerospace Epoxy Nanocomposites”, SAMPE Journal Volume 37, Number 5, 2001

[10] Chenggang Chen, “Curing Behavior and Processing/

Property Relationship of Epoxy Nanocomposites”, Proceedings

of the 35th International SAMPE Technical Conference, 2003

[11] William Tarasen, “Pyrograph Carbon Fibers”, AMPTIAC

Quarterly Volume 2, Number 4, Winter 1998
The AMPTIAC Quarterly, Volume 8, Number 1 59

INTRODUCTION

From planet Earth, the boundlessness of space may seem 

tranquil enough, yet this same environment is filled with a

potentially deadly array of problems for both manned and

unmanned spacecraft. Even before mankind soared into the

heavens, scientists and engineers knew space was a harsh environment, prone to remarkable fluctuations in radiation levels

and temperature, either of which by themselves, or in combination, could cripple or disable a spacecraft over a period of

time, or even instantaneously. One of the crucial challenges

was finding efficient and cost effective ways to protect a spacecraft’s critical subsystems (e.g., power supplies, life support 

systems, sensitive electronics, communications equipment 

and sensors), as well as the spacecraft structure, from an unforgiving and dynamic environment. 

Protecting spacecraft and their working components from

the harmful effects of extreme temperatures (a prime concern

of design engineers) is no simple task. This is the main function

of the spacecraft’s thermal control system (TCS). A welldesigned and properly operating TCS is essential to any spacecraft, given the varying amounts and forms of thermal energy

bombarding it at any given time. The principal forms of energy include solar radiation, albedo (electromagnetic radiation

reflected off of spacecraft surfaces), heat generated by onboard

equipment, and earth-emitted infrared (IR) rays [1].

The challenges imposed by the harshness of space become

even more pronounced when one considers the extreme shifts

in temperature endured as orbiting spacecraft repeatedly pass

through day and night, a scenario referred to as “thermal

cycling.” A satellite in low-earth orbit, for example, must survive temperature swings from –80 to +80° Celsius (a range of

about 288° Fahrenheit) approximately every 90 minutes; not

once, but tens of thousands of times during its operational 

lifetime, all while continuing to perform its mission!

A spacecraft’s TCS can be categorized as passive, semi-passive,

or active. The most prevalent of the three forms of thermal

control for space environments, passive subsystems were 

initially developed in the 1960s, relying heavily on reflective

coatings such as paints and mirrors. Passive subsystems may

also employ multilayer insulation (MLI), phase-change

devices, and radiators. Semi-passive subsystems go one step 

further, incorporating simple temperature-activated controls

to open or close conductive paths. Heat pipes and louvers are

sometimes used in semi-passive subsystems [2]. Active subsystems, by comparison, use heaters and mechanical refrigerators, while advanced active subsystems would have the ability

to change a spacecraft’s heat rejection capability as required in

real time. This article limits its discussion to the work of the

Air Force Research Laboratory’s Materials and Manufacturing

Directorate (AFRL/ML) in passive and active thermal control

systems.

TCS can vary from one spacecraft to another, depending on

the spacecraft’s mission and other key factors. However, all

share the basic objective of maintaining equipment temperature in a specified range throughout a spacecraft’s mission life,

thus ensuring optimum performance when equipment is

operating, and avoiding damage when not in use. The 

thermal control function ensures stable temperature levels for

delicate electronics and optical components, minimizes temperature gradients between units and along structural 

elements, and maintains boundary temperatures at interfaces

between subsystems[3] (please refer to the sidebar in this article for more information about heat transfer fundamentals).

MATERIALS AND MANUFACTURING DIRECTORATE’S 

CRITICAL ROLE

Thermal control technologies are developed through careful

design, analysis, and testing, and by determining and managing the factors that influence mission success. All of this must

be accomplished using available resources within the constraints of the space environment [4].

Scientists and engineers at the Air Force Research

Laboratory’s Materials and Manufacturing Directorate

(AFRL/ML) at Wright-Patterson AFB, OH are currently

Elizabeth S. Berman

Space Materials Direction Manager

Joel A. Johnson, Robert A. Mantz, Peter S. Meltzer, Jr.

(Anteon Corporation)

Materials and Manufacturing Directorate

Air Force Research Laboratory

Wright-Patterson AFB, OH
60 The AMPTIAC Quarterly, Volume 8, Number 1

involved in multiple highly innovative research and development programs aimed at improving thermal control technologies. Through in-house efforts and contracted work with

industry, they hope to increase the operational life for 

spacecraft in geosynchronous earth orbit (GEO) threefold.

This translates into a service life increase from an average 

of five to seven years to a minimum of 15 years. At the same

time, they are trying to realize these improvements, while 

significantly upgrading system performance to meet the needs

of smaller, higher-powered satellites and other advanced spacebased systems envisioned for the 21st century.

THERMAL CONTROL COATINGS RESEARCH 

AND DEVELOPMENT

The Directorate’s in-house research and development effort in

passive TCS, “Improved Passive Thermal Control Coatings,” is

being funded by the National Aeronautics and Space

Administration (NASA) and executed by the Directorate’s

Coatings Research Group. The Directorate’s external research

and development efforts in active TCS include: “Conducting

Polymer Electrochromic Cell Thermal Control Surface” and

“Selectable Wide-Range Epsilon and Alpha Thermostate via

Electronically-Controlled Reflectance (SWEATER) for

Spacecraft,” currently underway at Physical Sciences

Incorporated (PSI), and the “Conformal Appliqué for Thermal

Control in Space” research and development effort at

Sensortex, Inc. The last two are Air Force Small Business

Innovation Research (SBIR) projects. 

Passive Thermal Control 

Current state-of-the-art thermal control systems for spacecraft

use passive technology. The absorptance (i.e., amount of solar

energy absorbed and turned into heat) and emittance (i.e.,

capacity to emit IR energy for cooling) of

these materials is fixed at the time of application and are a function of the raw materials

used and the processing techniques employed

in applying them to the space hardware. The

primary change that occurs to these materials

in-service is degradation of the physical and

optical properties due to the harsh space environment. For passive thermal control paints,

the change in solar absorptance goes from initial values of 0.15-0.20, to final values in

excess of 0.4 after five to seven years of service.

For a long in-service life, the TCS must be

designed for the higher absorptivity value (i.e.

0.4), which means a larger, heavier, more

expensive TCS. 

Researchers in ML’s Nonmetallic Materials

Division (Nonstructural Materials Branch)

are actively engaged in an in-house research

project, co-funded by NASA, to improve 

the performance of passive thermal control

coatings (TCC) through weight reduction

and improved end-of-life properties*. The

researchers are working to reduce film thickness and lower the specific gravity (i.e., density) of the coating. They are also working to improve end-oflife solar reflectance properties by using materials with lower

overall solar absorptance, and which do not degrade significantly when exposed to the space environment.

Improvements in end-of-life properties will lead to additional weight reductions, given the smaller thermal radiator 

surface area required.

Since it is very difficult to qualify new materials for space

applications, the in-house research team has begun by optimizing passive TCCs through formulation changes of existing

material combinations. A diffuse reflective coating, for example, requires a binder (typically made of potassium silicate or

methyl silicone) and a pigment with a high refractive index

(such as zinc oxide) to scatter incident radiation back out of the

film, so that it does not pass through and possibly become

absorbed by the substrate. These coatings are noticeably bright

white in appearance. For a given combination of materials, the

scattering efficiency (i.e., the percentage of incident light that

gets reflected per unit film thickness) is dependent upon two

major factors: the pigment volume concentration (PVC) and

the pigment particle size distribution. Intuitively, as the PVC

increases, one would expect the scattering efficiency to increase.

However, after a certain PVC is reached, there is a point at

which the pigment particles begin to crowd together and the

scattering efficiency is actually reduced. 

Continuing to even higher PVC’s, there is a point at which

there is just enough volume of binder present to fill-in the

interstitial volume between the pigment particles, referred to as

the critical PVC or CPVC [5]. At PVCs above the CPVC,

voids are now present in the film due to the insufficient volume

of binder present. These voids assist in the scattering of incident light and the efficiency suddenly increases once again;

however, this is often at the price of poor physical properties

Figure 1. Scattering Efficiency is a Function of Pigment Volume Concentration.

Increasing PVC

CPVC

Voids Present

Binder =

Pigment=

Scattering Efficiency
The AMPTIAC Quarterly, Volume 8, Number 1 61

due to the very small amount of binder present. Figure 1

depicts this relationship.

The current passive TCC in widespread use is very porous

because it is formulated above the CPVC. These porous 

surfaces are easily contaminated by dust and oils that are

extremely difficult to remove and can significantly increase

solar absorptance. ML researchers are investigating coatings

formulated below the CPVC to reduce porosity, while maintaining the ability to use minimal film thickness. An added

benefit is that the weight of films with identical thickness is

lower because there is a lower volume concentration of the

high-density zinc oxide (ZnO) pigment.

The particle size of the pigment also plays an important role

in light scattering efficiency. The researchers have used Mie

scattering theory[6] to determine the optimal particle size of

ZnO pigment to be used in a thermal control coating (TCC)

(Figure 2). The results of the modeling indicate the optimal

particle size of ZnO is between 0.25 and 1.50 microns. Particle

size analysis of the ZnO pigment currently in use showed the

presence of ZnO particles and clusters between three and five

microns. Since then, ML researchers have investigated the performance of a ZnO pigment whose size falls within the optimal

range and have obtained coatings with equivalent solar

reflectance values using thinner, lighter weight films. These

coatings are being evaluated for accelerated exposure durability.

One of the disadvantages of using a high refractive index

material like ZnO as a pigment is the associated strong UV

absorption characteristic. A significant amount of solar energy

resides in the UV region (200-380 nanometer (nm) wavelength) and coatings that use ZnO absorb nearly all of this

energy, which can be transformed into heat or ultimately lead

to the chemical degradation of the TCC. An ideal TCC would

reflect all incident light ranging from UV to visual (VIS) to

near infrared (NIR). Unfortunately, there is a very limited

selection of space-stable coating materials that are UV-transparent; thus innovative methods are needed to obtain such

highly reflective coatings.

One method of achieving this is to introduce voids (in a controlled manner) into the binder. In this case, scattering is the

result of the binder-void refractive index difference rather 

than the binder-pigment difference. ML researchers are 

investigating synthesis of hollow silica (SiO2) spheres of

appropriate size for this purpose. When incorporated into a

UV-transparent binder such as potassium silicate, the internal

voids of the hollow silica spheres promote scattering without

introducing film porosity. Additionally, because silica is also

very UV-transparent, this coating design would be capable of

extended reflection into the UV region (Figure 3). Since

almost no incident solar energy would be absorbed, there

should be little driving force for exposure degradation either.

While coatings using this technique may require a larger film

thickness, the overall weight would be lower due to the use of

voids in place of a dense pigment. The result of these research

efforts will be a TCC that has lower overall solar absorptance

and weight, while lasting longer than the current state of the

art coating.

Active Thermal Control

Due to passive thermal control’s limitations, there is a strong

need for active thermal control in some spacecraft. Active thermal control manages the effects of sudden changes to incident

radiation experienced in the space environment. Active spacecraft TCS responds in real-time to these changes by adjusting

the level of power-assisted cooling. Active thermal control can

offer several advantages over passive thermal control. First of

all, large temperature swings in a spacecraft, resulting from sudden changes in the space environment, can be reduced to just a

few degrees. While there are no power requirements for passive

TCCs, the power requirements for active devices (currently

under development) can be reduced by at least one order of

magnitude over current state-of-the-art systems, (e.g. louvers).

ML presently has three contractual programs under way to

develop these devices.

Figure 2. Scattering Coefficient is a Function of Both Particle Size

and Incident Wavelength.

Figure 3. Schematic Cross-Section of a Thermal Control Coating

that Uses Hollow Silica Particles to Backscatter Incident Solar

Energy via the Presence of Controlled Voids.

Scattering

Coefficient

(1/ μm)

0

0.5

1.0

1.5

2.0

2.5

3.0

3.5

4.0

4.5

5.0

5.5

6.0

6.5

7.0

7.5

8.0

0.0 0.5 1.0 1.5 2.0 2.5 3.0 3.5 4.0 4.5 5.0

ZnO Particle Diameter (μm)

Wavelength (nm)

1000

960

920

880

840

800

760

720

680

640

600

560

520

480

440

400

UV/VIS/NIR

IR

Potassium Silicate

Silica Shell

Hollow Core
62 The AMPTIAC Quarterly, Volume 8, Number 1

Variable Emittance Materials Using Conducting Polymers

The first contractual effort involves variable emittance materials. The program is entitled “Conducting Polymer

Electrochromic Cell Thermal Control Surface” and focuses 

on conducting polymer materials for the electrochromic stack

supporting the variable emittance function. Under this concept, changes in emittance are performed in real time using 

the electrochemistry of the selected materials through an

applied voltage. To date, this research and development effort

has successfully demonstrated the proof-of-concept, and has

also completed the cell design. Development of solid-state 

electrochromic devices with electronic leads and connectors 

is under way, including the development of the device’s outer

window.

Voltage cycling stability testing is also under way. Space 

environmental effects materials qualification testing has yet 

to be completed. Passive space flight testing for a minimum of

three years in duration is in progress on the International Space

Station as part of the Materials on the International Space

Station Experiment program (MISSE) and active space flight

testing opportunities are currently being considered.

Selectable Wide-Range Epsilon and Alpha Thermostat via 

Electronically-Controlled Reflectance (SWEATER)

SWEATER is the Directorate’s second contracted active 

thermal control device R&D effort. The objective of this

Space-Based Infrared (SBIR) program is to develop spacestable thin-films, and to build and optimize prototype active

thermal control devices for both variable absorptance, and

variable emittance. Phase I of the SBIR was implemented to

identify potential candidate materials. Phase II, recently initiated, will further develop these materials into viable devices

and will include simulated space flight testing.

The SWEATER program could lead to significant improvements in spacecraft thermal management by reducing 

temperature variation, heater power, weight, and cost with

the additional application of IR signature control. To date,

this program has successfully demonstrated the feasibility of

the variable solar absorbtance device concept, including the

reversible electrodeposition of reflective metal ions. The team

has demonstrated that the lower and upper limit-controlled

solar absorptance is dependent upon the quality of metal ion

deposition. The research team has analyzed the feasibility of

the variable emittance device concept using computer modeling and further analysis is also planned.

Conformal Appliqué for Thermal Control in Space

The objective of the third effort, the “Conformal Appliqué for

Thermal Control in Space” SBIR program is to develop an

electrostatically switched radiator (ESR) appliqué for tunable

emittance (between 0.2 and 0.8). The appliqué system is

expected to provide a minimum of 15 to 20 years in-service

thermal management within the low-earth to geosynchronous

environments, superior heat transfer between the spacecraft

and the external space environment, and improved costeffectiveness. Phase I of this program identified potential 

candidate materials. The goal of Phase II (recently initiated) is

to further develop these materials into viable devices, complete

simulated space flight testing and, if possible, undertake and

complete actual space flight testing in conjunction with NASA’s

ST-5 space flight experiment.†

The ESR operates by actuating a thin compliant membrane,

typically a thin polymeric film, in and out of contact with the

surface of the spacecraft using an applied voltage (see Figure

Figure 5. Results on Conductive Membrane with Insulating Base. This is a Plot of Radiated Power versus Time, with Voltage being

Alternately Applied to the Sample (Higher Power) and Removed (Lower Power).

Figure 4. Micro-ESR Structure[7].

ESR Membrane

Spacer

Conductive Coating

Outer Layer

of Spacecraft

6

5.5

5

4.5

4

3.5

0 500 1000 1500 2000 2500 3000 3500

t[sec]

P [W]
4). When in contact, heat is conducted from the spacecraft

skin to the radiation membrane, where it is radiated out to

space. Without the applied voltage, the membrane loses 

contact, stopping heat conduction, and thereby reducing the

emitted radiation. There are very stringent requirements for

this membrane. It must be very thin and flexible to ensure low

operating voltage requirements. At the same time, it must 

have high thermal absorption (emissivity) with low optical

absorption, and must also have an outer conductive layer. 

A larger ESR device, previously developed on another 

program, uses a conductive paint to apply voltage to the membrane, which is satisfactory for those devices but is too thick

for the desired micro-ESR that this program is centered

around; hence, other options are being investigated. The most

promising alternative is the use of a transparent conductor,

such as indium-tin-oxide (ITO), over-coated with an anti-reflection coating. This

formulation results in a visually transparent film with high emissivity. A second

approach uses a polyimide layer filled with

conductive carbon black. Along with a

high emissivity, this structure unfortunately also has a high solar absorption, making this approach

unsuitable for use in actual devices but suitable for demonstrating proof-of-concept. Thus far, the experimental results

for this approach have been encouraging.

The ESR system requires a space environment to operate;

therefore, it is necessary to simulate these conditions for testing.

A contractor has developed an experimental chamber, which

employs an “Emissivity Measurement System” for this purpose.

This is a calorimetric system, whereby lost heat is measured by

the amount of electrical heat required to maintain the sample

at a constant temperature. 

Measurements on a carbon black filled membrane, using this

space-like environment test, are shown in Figure 5. Since the

carbon filler makes the sample conductive, no additional electrode is required. Voltage is alternately applied to the sample

(high radiated power) and removed (low radiated power). This

shows a change in radiated power of 1.22 watts with electrical

switching, with an effective change in emissivity of 0.5. These

test results are actually quite good in comparison to existing

technologies. However, changes in the film thickness and composition, planned during Phase II of this SBIR program, should

improve performance even more – perhaps with emissivity

changes of up to 0.8, resulting in effective emissivities ranging

from 0.1 to 0.9.

SPACE ENVIRONMENTAL EFFECTS CHARACTERIZATION

The Directorate’s testing facilities consist of three components: (1) the space environmental effects test facility (Figure

6), (2) the Optical Measurements Facility (Figure 7); and 

(3) the Directorate’s Materials Characterization Laboratory

equipment (Figure 8). The space environmental effects 

test facility simulates the GEO environment, approximately

22,000 miles above the Earth’s surface, complete with ultraviolet (UV) radiation and electron exposures. Pre- and postThe AMPTIAC Quarterly, Volume 8, Number 1 63

Figure 7. Optical Measurements Facility with 

Front View of the Facility’s Spectrophotometer.

Figure 8. Materials Characterization Laboratory 

Showing X-Ray Photoelectron Spectroscopy (XPS)

Instrument.

Figure 6. Materials

and Manufacturing

Directorate’s Space

Environmental Effects

Test Facility.

Figure 9. Location of Two Passive Experiment Carriers 

(PECs) on International Space Station.

Figure 10. PEC

Showing Specimens.
64 The AMPTIAC Quarterly, Volume 8, Number 1

measurements of the simulated space environment are 

performed in the Optical Measurements Facility and the

Materials Characterization Laboratory. Actual onboard space

environmental testing of new materials is conducted as space

flight opportunities are available. Many new thermal control

materials are currently being tested on MISSE, managed by

ML in conjunction with NASA and industry (Figure 9 and

Figure 10) [8].

CONCLUSION

The Air Force is much better positioned to meet its major

objectives for the 21st century through enabling ML programs

in thermal control coatings and space environmental testing.

ML is actively supporting the Air Force’s migration to space,

and innovative research and development programs are in place

to meet both near and long-term Air Force spacecraft needs for

thermal management. Both programs have the potential to dramatically improve the performance and durability of spacebased systems supporting the Air Force mission.

Continuing research and development in these technologies is

critical to qualification and implementation of improved passive and active thermal control systems. Innovations in thermal

control technology for spacecraft will have positive implications

for the private sector as well, directly impacting the expected

lifetime of commercial satellites.

NOTES AND REFERENCES

* Thermal control systems are designed to fully protect spacecraft for the entirety of its service life. To do this, engineers

must account for how much the TCS’ protective capabilities

degrade over that time, and size the system based on the predicted properties at the end-of-life. This means that most TCS

are over-designed to insure thermal protection through end-oflife (degrading from a performance level of ‘extremely good’ to

‘adequate’). If future TCS are made to degrade less, than it

won’t be necessary to overdesign their systems as much, saving

Figure A. Conductive Heat Transfer.

Thermal control of spacecraft is all about managing

heat transfer into and out of systems and vehicles. For

the uninitiated, we offer this brief primer on the

major modes of heat transfer and what role each plays

in thermal control management. 

There are three major modes of heat transfer: conduction, convection, and radiation. Conduction is the

primary mechanism by which heat propagates

through solids. The second law of thermodynamics

dictates that heat travels from a hotter point to a colder point. With few exceptions, the temperature gradient through a solid is linear in profile (Figure A).

Convection is the primary mechanism by which

heat propagates from a solid via moving fluids (both

liquids and gases). Convective heat transfer rates tend

to be several orders of magnitude higher than conduction for similar temperature gradients (Figure B).

Radiation is the most complex (and least understood) of the three mechanisms. Unlike conduction

and convection, radiation does not require matter as

a transfer medium (it’s actually more efficient in a

vacuum). Moreover, conductive and convective heat

transfer are exclusively driven by the size of the temperature gradient across the transfer area – radiant

heat transfer depends not only on the gradient, but

on the absolute temperature of the source material as

well. While any electromagnetic transmissions, from

cosmic rays to radio waves, may be classified as radiation, only those in the wavelength range of 100 nm to

100,000 nm (from near-UV through visible light to

IR) may be considered thermal radiation. For an

exposed spacecraft component, the incoming thermal

Heat Transfer Mechanisms

Thickness of Solid

Direction of

Heat Transfer

(T2 > T1)

T1

T2
The AMPTIAC Quarterly, Volume 8, Number 1 65

Figure B. Convective Heat Transfer.

Figure C. Radiant Heat Transfer.

radiation (mostly from the sun) is either absorbed by

the constituent material (which heats it up), reflected

back off into space, or is transmitted through the

medium. These three characteristics are known as

reflectivity, absorptivity, and transmissivity, respectively

(Figure C).

Heat Transfer in the Space Environment

For Earth-orbiting space systems, thermal radiation is

the primary heat transfer mechanism (conduction is

an important secondary one). Convection is negligible in the vacuum of space. As an object in space

absorbs radiant energy, the object’s temperature

begins to rise. Simultaneously, the object emits some

of this energy back out into space as thermal radiation

– this property is known as emissivity (ε). The amount

of radiation emitted is a direct function of the object’s

temperature. As the object’s temperature increases, so

does the emittance. The object’s temperature will rise

until eventually the absorptance rate equals the emittance rate – at this point, the object has reached an

equilibrium temperature.

Each TCS uses these principles to serve their own

application: some systems are designed to cool components, others to heat them, and still others to maintain temperatures in some critical range. AFRL/ML,

NASA, and industry must examine each application

carefully and devise a strategy that best protects the

vehicles and ensures the continued nominal performance of their subsystems.

money and improving reliability.

† Part of NASA’s New Millennium Program to develop and

test advanced technology in space flight, Space Technology 5

(ST-5) is an effort to fabricate microsatellites. The mission,

scheduled sometime in 2005, will launch multiple miniature

spacecraft (known as small-sats) to demonstrate and flight-qualify several leading edge technologies which measure solar and

magnetic effects in the atmosphere.

[1] M. Fischer, H. Alvidres, and D. Hoetger, “10 Thermal

Control,” http://www.tsgc.utexas.edu/archive/subsystems/

thermal.pdf, February 1995

[2] “7.0 Thermal Control Subsystem,” http://students.cec.

wustl.edu/~sapphire/design/ log/thermal.html, January 1995

[3] “Spacecraft Thermal Control”, prepared by the European

Space Agency, http://www.estec.esa.nl/thermal/default.html,

August 1, 2003

[4] “Spacecraft Thermal Control”, prepared by the European

Space Agency, http://www.estec.esa.nl/thermal/default.html,

August 1, 2003

[5] J.C. Weaver, (1992) CPVC, Critical Pigment Volume

Concentration - An Overview. Journal of Coatings Technology

64, 45-46

[6] C.F. Bohren, and D.R. Huffman (1983) Absorption and

Scattering of Light by Small Particles. John Wiley & Sons, Inc.:

New York, NY. (Note: The Mie scattering theory is a comprehensive mathematical-physical theory pertaining to the scattering of electromagnetic radiation by spherical particles, developed by Gustav Mie in 1908)

[7] W. Biter, “Thermal Control Using Multi-Spectral

Switching” (Micro-ESR structure drawing), http://www.

winbmdo.com/scripts/sbir/abstract.asp?log=0159&Phase=1&

Ph1Yr=96&firm_id=2096&int=96

[8] “Overview: Spacecraft Materials Development Programs for

Thermal Control Coatings & Space Environmental Testing”

(briefing photos), Air Force Research Laboratory Materials and

Manufacturing Directorate, March 14, 2003

Direction of

Heat Transfer

(T2 > T1)

T1

T2

Solid Surface

Fluid Flow near a

Solid Object

Typical Temperature Profile through a Fluid

Reflected Radiation (ρ)

Incident

Radiation

Absorbed

Radiation (α)

Transmitted

Radiation (τ)

For any given unit of incident radiation, the following relationship is true:

ρ + τ + α = 1

Emitted Radiation (ε)

(Flux α T4)
Dr. Elizabeth S. Berman is a senior materials research engineer and the space materials direction leader in the

Nonstructural Materials Branch within the Materials and Manufacturing Directorate (AFRL/MLBT) at WrightPatterson AFB, OH. She is in charge of developing both thermal control coatings and tribological materials for

space. She is the government lead for the only Air Force owned facility capable of simulating the space environment for testing of space materials. She also is the AFRL/ML Program Manager for the AFRL MISSE (Materials on

the International Space Station Experiment). Dr. Berman holds a BS in Chemical Engineering from the University of

Minnesota, an MS in Chemical Engineering from Purdue University, and a PhD in Polymer Science from the

Pennsylvania State University.

Dr. Joel A. Johnson is the coatings direction leader for the Materials and Manufacturing Directorate’s Nonstructural

Materials Branch (AFRL/MLBT) at Wright-Patterson AFB, OH. Dr. Johnson conducts research on a variety of coatings projects, including those related to thermal control, electrical conductivity, environmental compliance, and nontoxic corrosion protection. He holds a BS in Chemistry from the University of Wisconsin – River Falls, and a PhD in

Polymers And Coatings from North Dakota State University. Dr. Johnson is an active member of both the American

Chemical Society and the Federation of Societies for Coatings Technology.

Dr. Robert A. Mantz is the Research Group Leader of the Paints and Coatings Group of the Materials and

Manufacturing Directorate’s Nonstructural Materials Branch (AFRL/MLBT) at Wright-Patterson AFB, OH. He is

actively engaged in developing new chromate-free coatings systems for aircraft, as well as new space environmental protective coatings for polymeric materials and improved thermal control spacecraft coatings. Dr. Mantz

also continues to investigate novel applications utilizing ionic liquids. Dr. Mantz earned his BS from the United States

Air Force Academy, his MS from California State University – Northridge; and his PhD from North Carolina State

University.

Dr. Peter S. Meltzer, Jr. is senior technical writer and editor at AFRL/ML’s Technology Information Center (operated

by Anteon Corporation), Wright-Patterson AFB, OH. Dr. Meltzer received his PhD in Public Administration from

Kennedy-Western University. He also holds an MA in Political Science (with a concentration in National Security

Studies) from Georgetown University, an MA in Communications, and a BA in Journalism from the University of

Massachusetts.

66 The AMPTIAC Quarterly, Volume 8, Number 1

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The use of monolithic glass to produce large, rigid segmented members for lightweight space-based mirror systems has reached its limit due

to the long lead times, high processing costs, and launch load/weight requirements. New material solutions and manufacturing processes

are required to meet the Air Force’s directed energy weapons, reconnaissance/surveillance, and secured communication needs. Mirror structural substrates made out of advanced composite materials (metal, ceramic, and polymer), foams, or microsphere arrays should allow for

CTE and modulus tailorability, low-density, high strength, stiffness, and toughness. Due to the multi-phase complexities of these new systems, mechanical polishing will be difficult. This article introduces some of the issues surrounding the materials and processes traditionally used to fabricate mirrors for space applications, and also covers the evolution from conventional materials and processes to newer concepts for use in space-based mirrors. Selected material and process approaches are detailed as well.

The AMPTIAC Quarterly, Volume 8, Number 1 67

THE NEED FOR ADVANCED MATERIALS AND METHODS

Large, lightweight, high precision mirrors are critical for

enhanced surveillance/reconnaissance missions, directed energy

weapons and communication systems, laser radar systems, 

X-ray and UV telescopes, as well as

large astronomical telescopes. The

Department of Defense (DOD), the

National Aeronautics and Space

Administration (NASA), and other

US Government agencies have 

developed concepts for airborne and

space-based systems that will require

mirrors ranging in size from 0.1 to

100 meters in diameter. In all of these

applications, mass, size, natural frequency, and reliability are major issues

due to launch and flight constraints as

well as anticipated life times.

Cost, production infrastructure,

and scheduling have also become

important technology drivers in some

missions where large surface area mirrors are needed. For example, a primary mirror for an airborne

directed energy system might consist of a monolithic design

with a diameter up to 1.5 meters (allowing it to fit inside existing airframes). In such a case, the mirror is shielded or encased

within the airframe, thus the mirror’s fracture toughness would

not be a major issue. In such circumstances, existing lightweight

glass mirror technology might suffice[1]. In contrast, a spacebased directed energy system would require a constellation of

very large mirrors, 10 to 15 meters in diameter. They would use

a segmented design with segments no larger than roughly 3.25

meters (so they could be stowed in the Space Shuttle cargo bay

or a launch vehicle shroud). Once launched to proper orbit,

these segments would need to be assembled or deployed. To

meet the payload weight constraints

of launch, the mirror would also have

to be very lightweight (with an areal

density (mass per unit surface area) of

less than 15 kg/m2). The reliability

and robustness of these mirror sections are major concerns because they

would be unshielded (open to the

space environment) and thus subject

to high launch loads, in-space assembly damage, and micrometeor

impacts. Most likely, monolithic glass

mirrors would not be able to meet

these requirements; hence the need to

develop advanced mirror materials

and design practices.

To meet the anticipated demand for

space-based applications, the US

industrial base of aerospace contractors would require the capacity to manufacture approximately five to seven 10-meter diameter primary mirrors (segmented) per year for a period of 15 or

20 years. Unfortunately, the infrastructure needed to meet such

a large production quantity does not currently exist in the US.

Therefore, substantial government and corporate investment

would be required to meet this goal. Some of this infrastructure

investment could be shared by NASA, as long as common goals

could be achieved (e.g. the mirrors could be cooled to cryogenic

Lawrence E. Matson, David H. Mollenhauer

Materials and Manufacturing Directorate

Air Force Research Laboratory 

Wright-Patterson AFB, OH

Image courtesy of NASA

Figure 1. James Webb Space Telescope – Being

Constructed to Replace the Hubble Space

Telescope.
68 The AMPTIAC Quarterly, Volume 8, Number 1

temperatures without increasing their figure or finish distortions). They could be employed in systems such as the James

Webb Space Telescope (Figure 1), which is envisioned to replace

the Hubble Space Telescope.

TRADITIONAL MATERIALS AND PROCESSES

Monolithic Glass Mirrors

Silica-based glasses have been used for hundreds of years as

both the optical and structural substrate for mirrors. Glass

makes a good optical substrate due to its low softening point,

low hardness, and its absence of a grain structure (amorphous).

This allows it to be formed, ground, and polished into complex

shapes with precise figure (≅ λ/20 RMS*,†) and very smooth

surface finishes (as low as ≅ 3Å RMS). Once the figure and 

finish requirements are met for a particular application, the

mirror is then coated with a very thin metal film which 

provides its reflective characteristics.

Another benefit for using glass in mirror designs is that its

amorphous structure allows for a wide variation in chemistries.

Varying the chemical composition of the elements contained in

the glass is a method used to adjust the coefficient of thermal

expansion (CTE) to values approaching zero at various use

temperatures. These mirror glasses provide dimensional stability (of the mirror figure) with respect to the thermal variations

of the environment. A brief review of glass chemistry shows

how this was made possible.

Silica-based glass is an inorganic material with a metastable

amorphous structure (no long range order). Most glasses are

produced by supercooling their liquids to a rigid condition

without crystallization. Some glasses are prepared without 

cooling from the liquid state such as vapor grown or solution

grown glasses. In either case, the atoms of a silicate glass form a

continuous framework of silicon and oxygen atoms tetrahedronally bonded together at their corners. The bonds are

strong, but they have large variations in the Si-O-Si bond angle.

Such a random network is not necessarily uniform, and local

variations in density and structure are to be expected. This 

variability in local density and bond angle allows the absorption

of thermally inducted vibrational energy through transverse

modes of vibration and the adjustment of bond angles, resulting in a low CTE.

Corning Glass uses a vapor growth process and a chemical

composition employing fused silica (SiO2) and 7% (by weight)

Titania (TiO2) to produce its Ultra Low Expansion (ULE) glass

[2]. One of the contractors under Air Force (AF) sponsorship

recently produced this composition by a sol-gel approach [3].

The CTE of a glass can also be tailored by the homogeneous

precipitation of negative CTE crystalline phases throughout

the glass matrices. This is the case with Zerodur glass [4], a

glass–ceramic system containing Li2O, Al2O3, and SiO2 (LAS)

with a small amount of TiO2. ULE and Zerodur both suffer

from very low thermal conductivity, strength, modulus, and

fracture toughness. Since they can be polished to a complicated figure with a very smooth surface finish, glasses such as ULE

and Zerodur are good optical surfaces. However, since they

both possess poor mechanical properties they may not be the

best material choice for a space-based mirror.

To surmount the aforementioned shortcomings in mechanical properties, thick and heavy mirror designs have typically

been used along with good handling and mounting practices.

Reliability and robustness were assured in space-based systems

by encasing mirror systems into metallic tube structures.

However, in future airborne and space-based mirror systems,

where weight and size are constrained by payload requirements,

this shielded approach, and high areal density design may not

be feasible. Non-shielded, low areal density designs may be

required to meet system goals; and robustness must be provided by either new structural substrate materials (with high 

modulus, strength, and fracture toughness) or with new hybrid

designs.

Metallic Mirrors 

Because they are easily formed, machined, polished, handled,

and are relatively inexpensive, metals like aluminum and 

nickel have been used as mirror materials in terrestrial applications for years. Additionally, most metals are inherently highly

reflective, so they do not require coatings to yield high broadband reflectivity. Unfortunately, they generally have high CTEs

at ambient temperatures, which limit their applications.

However at cryogenic temperatures below 70°K, beryllium has

a near zero CTE. This along with its low density, high stiffness,

and cryogenic-thermal conductivity make it a desirable mirror

material for highresolution infrared

imaging. Beryllium’s

utility is compromised by its high

cost, toxicity, slow

machineability, and

anisotropic behavior

due to its hexagonal

crystal structure.

Processing by standard metallic hot

working methods

such as forging and

rolling results in texturing that cannot

be homogenized through recrystallization. This textured grain

structure results in an anisotropic modulus and thermal expansion behavior of the mirror substrate. Only powder processes

such as hot isostatic pressing, vacuum hot pressing, sintering, or

vacuum plasma spray will yield a randomly oriented grain

structure that shows isotropic behavior (Figure 2).[5, 6, 7]

Metals can also be used to create the optical substrate, which

is the polished material directly between the reflective surface

and the structural substrate. Several manufacturers have used

polycrystalline silicon as an optical substrate (also known as a

mirror cladding). Its isotropic behavior (diamond cubic structure), low density and CTE, high thermal conductivity and

excellent polishability make it a good candidate for an optical

substrate [8]. Polycrystalline silicon can be fabricated using

melt/solidification, sintering, plasma spraying, physical vapor

deposition (PVD), and chemical vapor deposition (CVD).

Highly polishable single crystal silicon from the electronics

industry has also been used as an optical substrate material on

Courtesy of NASA

Figure 2. A 1.4 Meter Beryllium Mirror.
The AMPTIAC Quarterly, Volume 8, Number 1 69

a flat mirror, but its size is limited to approximately 12 inches

by single crystal growth methods. A detriment to the use of 

silicon is that it has a low modulus and an affinity for atomic

oxygen, which may limit its applicability for certain missions.

NEW MATERIALS AND PROCESSES

Increasing the size of mirrors for space applications requires

that new materials and processes be developed to reduce the

areal density as compared to monolithic glass and metallic 

mirrors. One approach that has been under development for

quite some time involves the use of monolithic glass sheets that

are bonded to a lightweight honeycomb core. Since an internal

honeycombed structure is used, this approach allows for the

same performance as a monolithic glass mirror but at a greatly

reduced weight. Another approach that has seen quite a bit of

development uses monolithic silicon carbide fashioned using

various fabrication techniques. Both of these methods are 

discussed further in the following sections.

Lightweight Monolithic Glass Mirrors

Significant effort has been made over the past twenty years to

reduce the weight of airborne and space-based glass mirrors.

These efforts were driven by needs to reduce launch costs,

decrease thermal mass (to equilibrate temperatures in shorter

times), and to minimize the weight and complexity of the 

support and handling equipment.

The Hubble Space Telescope mirror was a lightweight

design. It was produced using a frit bonding process to join

individual ULE glass plates to create an open-faced honeycombed core structure[5]. The core was then frit bonded

between two ULE glass face-sheets to produce a lightweight

mirror blank. The mirror blank was ground, polished, and

vapor coated with aluminum to create a visible light reflective

surface. This mirror is 2.4 m (8 ft) in diameter and weighs 828

kg (1820 lbs). It has an areal density of approximately 200

kg/m2. The major problems encountered in building this mirror were thermal distortions and uneven stresses in the plates

and joints. Additionally, the processing method was extremely

labor intensive, time-consuming, and expensive.

Many programs have attempted to overcome the processing

difficulties experienced while building the Hubble mirror and

to develop methods/processes to substantially reduce the areal

density to less than that of the Hubble mirror. Most recently,

the Advanced Mirror Systems Demonstrator (AMSD) program

(run by NASA’s Marshall Space Flight Center), a jointly funded effort by the Air Force, NASA, and the National

Reconnaissance Office (NRO), attempted to advance the stateof-the-art in large, lightweight mirrors[5, 9].

In the AMSD program, an abrasive water jet (AWJ) milling

technique was used to fabricate thin-walled, open-ended, 

honeycomb core structure out of bulk ULE glass (Figure 3).

The front and back faceplates were attached using a low 

temperature fusion bonding process; producing a honeycomb

sandwich-type design. A 1.4-meter wide hexagonal mirror 

segment is shown in Figure 4. This lightweight glass mirror segment was then attached to a carbon fiber-reinforced composite

reaction structure with 16 actuators. This total mirror system

has an areal density of about 12 kg/m2. However, the contractors appear to have reached the lower limit in areal density that

one can expect for a monolithic glass mirror. Any additional

reductions would have to come though hybrid material

approaches, new materials, or composites. The AMSD 

Courtesy of NASA

Figure 3. ULE Honeycomb

Core Produced by AWJ

Milling in the AMSD

Program.

Courtesy of NASA

Figure 4. Kodak’s 1.4 Meter Lightweight Glass Mirror Segment.
70 The AMPTIAC Quarterly, Volume 8, Number 1

contractors were also very successful in reducing the cost and

fabrication time of a glass mirror to half that of a comparable

Hubble mirror.

Silicon Carbide-Based Mirrors

Silicon Carbide (SiC)-based materials are promising for use as

structural substrates in large mirrors because of their high specific stiffness (high modulus and low density) and their low

thermal distortion susceptibility (low thermal expansion and

high thermal conductivity) (Figure 5). However, there are

many different types of SiC-based materials, and each will 

polish differently producing variability in surface figure and

finish. This variability comes from the differences in hardness

and modulus between the phases present, their crystallography,

and morphology.

Pure SiC can be grouped into two crystalline forms: 1) alpha

(α-SiC), which has a rhombohedral or hexagonal structure and

is produced when the formation or processing temperature

exceeds 2000°C; and 2) beta (β-SiC), a face-centered cubic

phase that is formed by processes that occur below 2000°C.

The best surface finish for crystalline SiC would be expected

from a dense, fine equiaxed grain, single-phase, β-cubic structure because its hardness, strength, modulus, and thermal

expansion behavior should be isotropic and the grain shape

equiaxed. 

The highest density, smallest grain size, single-phase β-SiC is

produced by controlled re-nucleation using a CVD process.

Since the deposition rates are rather slow, β-SiC is mainly used

for depositing thick coatings or claddings (the optical substrate)

onto less expensive, less dense, sintered, or siliconized (excess

silicon) SiC based materials (the structural substrate). The polishability of CVD β-SiC cladding material is limited by its high

hardness. Too high a pressure on the polishing tool causes quilting-type print-through effects on lightweight honeycomb core

structures, while too low a pressure increases the polishing time

making it uneconomical.

However, β-SiC mirrors can be fashioned through other

techniques. Lightweight, near net-shaped β-SiC mirrors are

produced by taking a special grade of graphite and machining

it to the desired structure followed by conversion of the

graphite to silicon carbide through a gas phase reaction. This

reaction process has a very small but predictable dimensional

change. The as-converted β-SiC structure has approximately

20% porosity. This mirror substrate is then cladded with CVD

β-SiC to make a 100% β-SiC mirror. It is then polished using

conventional optical polishing.

Sintering of pure alpha (α-SiC) particles to full density is difficult unless both sub-micron powders and sintering aids are

used. Powders and binders are usually mixed and isostatically

pressed at room temperature into a green body[‡] blank. The

green body can then be machined into an oversized geometry

that is predicted by sintering shrinkage models. The oversized,

machined green body part is then sintered above 2100°C,

resulting in approximately 17% linear shrinkage. The sintered

SiC component is approximately 97% dense and contains 98

to 99% α-SiC, depending upon the amount of sintering aids

used. Sintered SiC produced by Boostec in France[10] has a

morphology consisting of 5 μm grain size with 3% porosity

showing up as 1 to 2 mm voids along particle boundaries. A 

1.3 m sintered SiC demonstration mirror has been produced

for the first near-IR telescope [11].

For optical applications at shorter wavelengths, sintered SiC

is coated with CVD SiC and the blank is ground, polished, and

etched by an ion plasma beam. This Boostec-sintered (α-SiC)

technology has recently been acquired in the USA by Coors Tek

in Boulder, Colorado [12].

Two-phased, siliconized-SiC is produced by several methods. The first method uses slip cast (α-SiC) powders followed

by sintering until they are approximately 60 to 80% dense.

This preform is then infiltrated with silicon metal by either gas

Figure 5. A Comparison of Parameters for a Variety of Mirror Substrate Candidate Materials.

Fracture Toughness

Al

Be

Si

C/Mg

C/Al

Si-foam

Sinter-SiC

C/SiC

Si-SiC/SiC

ULE

Zerodur

C/Pyrex

HM-C/Epoxy

HD-C Foam

Specific Stiffness Thermal Stability

1000

100

10

1

0.1

Fracture Toughness is in MPa-m0.5, Specific

Stiffness is in Units of GPa-cc/gm, and

Thermal Stability[§] (k/α) in W/m-ppm.
The AMPTIAC Quarterly, Volume 8, Number 1 71

phase deposition or liquid infiltration to fill the open porosity.

The second method involves slip casting SiC with carbon 

particulates in a polymer binder that decomposes to a carbon

network upon vacuum calcination. The calcined SiC/C body

is then exposed to either Si vapor or molten Si to convert the

carbon phase to SiC through a reaction bonding process. Any

remaining porosity is then filled with silicon metal. This

method is used by several manufacturers to produce mirror

substrates and optical support components. Both methods

produce a two-phased (SiC+Si) continuous structure that is

intertwined. The latter process is used to produce siliconized

α-SiC mirrors [13, 14].

NEW RESEARCH DIRECTIONS

There are many other approaches that are being considered for

developing space-based mirrors. Some of these approaches are

similar to the lightweight monolithic glass mirror discussed

above, in that a lightweight core is used to reduce the total mass

of the mirror. Included in these emerging technology approaches are glass mirrors with a lightweight core fashioned from glass

microballoons bonded together with glass deposited using solgel techniques. Other methods involve using various types of

foam materials or composites. The following sections discuss

these approaches.

Lightweight Glass Mirrors 

Sol Gel and Microballoon Arrays: The Air Force Research

Laboratory’s Materials and Manufacturing Directorate

(AFRL/ML) is sponsoring research to create near zero CTE

glass sol-gels for use as bonding

agents, surface coatings, and

claddings. Additionally, they

are fabricating near zero CTE

glass microballoons that can be

bonded by sol-gel deposited

glass, thus forming a microballoon array (Figure 6).

Lightweight glass mirrors will

be fabricated by bonding ULE

glass face-sheets to cores made

from microballoon arrays. This design should result in low areal

density mirrors that are more robust than current honeycomb

web designs. The sol-gel chemistry approaches could also allow

for the incorporation of nano-particles or nano-fibers to

increase the tensile strength, modulus, and fracture toughness

of the glass. Sol-gels could also be used to build up glass

claddings on glass ceramic composite mirrors or mirrors made

of other materials. Near zero CTE glass micro-balloons could

also be used for CTE control and strengtheners for polymerbased composites. 

Mirrors Constructed from Foams 

An open cell foam structure can provide structural support to a

dense facesheet without sagging, hence eliminating distortions

as seen in web-based mirror architectures. Foams also allow for

a substantial decrease in mirror weight. Foams have been produced out of a variety of materials including aluminum, silicon,

silicon carbide, and carbon. Foams can also be formed by the

intermixing of hollow spheres of glass with a polymer matrix,

thus making a substance referred to as syntactic foam. Below

you will see examples of these types of materials.

Silicon Foams: Silicon foam can be produced by several methods, including a displacement reaction between carbon foam

and SiO vapor.

SiOg + Cfoam → Sifoam + COg

It can also be produced by the diffusion of carbon through a

chemical vapor infiltrated silicon surface coating that has been

deposited onto the ligaments of open cell carbon foam. The latter process has been used to convert machined carbon foam

preforms into net-shaped silicon foam preforms[7, 15]. The

surface of the silicon foam is then coated with silicon particulate slurry or by CVD silicon until the surface pores are closed

off. A fully dense optical silicon substrate is then applied using

CVD, and the surface is polished to the appropriate figure and

finish.

Glass/Ceramic Syntactic Foams: Composite materials made

from glass, ceramic, or polymer microballoons with polymer

resin matrices, broadly known as syntactic foams, are showing

great promise for space-based mirror systems. Essentially, they

are highly filled polymer systems where the fill material is made

up of very small hollow spheres (2-100 μm diameter). Ceramic

and glass microballoons are of special interest for use in mirror

systems. These types of microballoons can have very low CTE

and are relatively stiff compared to their bulk densities.

Syntactic foams typically have low densities. Most advertised

syntactic foams, however, are essentially nonporous since all

interstitial spaces between microballoons are filled with resin. A

significant portion of the density of the composite is made up

of resin. For a low density, lightly loaded mirror, the resin’s only

purpose in the composite is to serve as a binder of sufficient

structural integrity for the expected mechanical and thermal

load. Any more resin would increase the density and CTE, as

well as the cure shrinkage of the whole composite. Recent

research at AFRL/ML has developed a methodology to produce

syntactic foam composites with very low resin content, hence

minimum density[16]. 

Carbon foams: An exciting new material with potential applications to space-based mirrors is carbon foam. There are a variety

of manufacturers that produce a wide range of carbon foam

products. These foams can be tailored (density and morphology) to some degree during manufacturing to display moderate

stiffness, moderate compressive strength, low density, low CTE,

low or high thermal conductivity, and high or low electrical

Courtesy of NASA

Figure 6. Near Zero CTE Glass

Spheres.

Figure 7. Carbon Foam Examples.

a) Carbon foam. 

b) Carbon foam sandwich with 

ordinary composite laminate facesheets top-coated with a layer of 

multi-walled nanofibers embedded 

in a polymer matrix.
72 The AMPTIAC Quarterly, Volume 8, Number 1

resistivity. They typically show in-plane isotropy on a plane perpendicular to the foaming direction [17]. Research is currently

underway to engineer a better microstructure to improve the

compressive, tensile, and shear strengths of these foams.

Additionally, work is ongoing to identify methods for closing off

the porous surfaces, adding a polishable substrate, strengthen

attachment locations, and protecting the surfaces from atomic

oxygen. Possible solutions to this problem include functionally

grading the materials, adding a facesheet to cover up the pores,

or adding a cladding material to fill up the surface pores. Figure

7 shows two examples of carbon foam, including one with

facesheets of ordinary composite laminate with a topcoat of

multi-walled nanofibers bonded by an epoxy matrix. Hybrid

mirror designs using low CTE glass sols as the close-off and

optical substrate materials are also being considered.

Composite Mirrors

Organic Matrix Composite Mirrors: Organic matrix composites made with conventional carbon fibers have been considered for use in space-based mirrors for quite some time. They

exhibit many properties that are extraordinarily good for mirror systems. They have a high specific stiffness and depending

on fiber choice, they can

have good in-plane thermal

conductivity. Most importantly, carbon fiber composites can also be designed

with near zero CTE for a

space-based mirror. Many

problems exist, however,

with using these materials

for mirrors. Polymer resin

systems can be sensitive to

a variety of environmental

factors that put into question their long-term stability. Highly uniform and

consistent raw materials

(composite prepreg –

uncured fiber/matrix

sheets) can be difficult to achieve. Finally, the multiphase

nature of the composite causes a fiber-print-through phenomenon, which deleteriously affects the surface finish of the

optic**. It is important to note that this problem affects all

classes of composite materials, not just organic matrix composites. A large PMC mirror (with an areal density of 10

kg/m2) produced for NASA Jet Propulsion Laboratory (JPL) is

shown in Figure 8.

Ceramic Matrix Composite Mirrors: The desirable features of

CMC mirrors are that their CTEs and areal densities are low

while their strength, modulus, and fracture toughness are high.

High strength and fracture toughness will be needed for durability and robustness in unshielded space-based mirror systems,

such as the James Webb Telescope and the Space-Based Laser.

Survivability and longevity requirements will demand fielded

mirrors be made of highly durable and stable materials.

Therefore, as a community we need to continue to invest in

scaling up these composite 

mirrors to very large sizes while

assuring thermal-mechanical

stability.

Many DOD and NASA

organizations have sponsored

numerous businesses to develop

carbon and SiC fiber-reinforced

SiC and siliconized SiC matrix

composites. These ceramic

matrix composites (CMCs)

have been fabricated in both

continuous and discontinuous

fashion by numerous companies using a variety of processing

routes. As in glass composites, carbon fiber reinforcements

allow for CTE and strength tailorability, but polishability problems arise when a weak carbon interfacial coating is deposited

on the fibers to promote crack deflection and improve fracture

toughness. Polishability is improved by adding a thick cladding

of dense CVD β-SiC to form an optical substrate. 

Many processes can be used to fabricate structural CMC 

substrates using chopped carbon fibers coated with a carbon

interfacial layer and bonded together in a SiC matrix. For

example, AIBG and ECM from Germany, molds a mixture of

SiC and C particles, polymer binder, and chopped carbon

fibers (with and without interfacial coatings) into a blank. The

mirror blanks are then heat treated under vacuum to form a

porous graphitized green body. The green body is easily

machined into a lightweight shape using standard CNC

milling equipment. It is then infiltrated under vacuum with

liquid silicon to form a C/SiC composite structure. [6,7],

(Figure 9). It is claimed that there is no noticeable shrinkage in

the infiltration/conversion process with virtually no residual

stresses left in the mirror. The composite is then rough ground

and cladded with slurry to form a SiC+Si surface layer that is

polished to a surface roughness of 2 nm. Other fabrication

methods use chemical vapor infiltration to densify a preform of

carbon or silicon carbide fibers. These composites are then clad

with either CVD β-SiC, PVD Si, or melt replication with Si to

form the optical substrate.

Metal Matrix Composite Mirrors: Continuous and discontinuous carbon fiber-reinforced magnesium and aluminum

matrix composites are currently being studied as possible 

mirror structural substrates[7,18,19]. As previously mentioned, the volume

fraction of carbon fibers can

be varied to tailor the net

CTE in a composite material. It can also increase the

strength, modulus, thermal

conductivity, and fracture

toughness of the base material. Carbon-reinforced aluminum and magnesium matrix composites are highly

machineable to a lightweight configuration without warpage.

These structural substrates are then coated with a CVD-Si

Figure 8. A 2-Meter Polymer

Matrix Composite Mirror Produced

for NASA JPL.

Courtesy of NASA

Figure 9. The Back Side of the

500 mm C/SiC Mirror with an

Areal Density of 8 kg/m2.

Courtesy of NASA

Figure 10. Three-Inch Diameter

C/Al and C/Mg Mirrors.
The AMPTIAC Quarterly, Volume 8, Number 1 73

cladding and polished to a surface roughness of 10 angstroms

RMS (Figure 10).

CONCLUSIONS

For many situations, monolithic glass remains the material of

choice for mirrors, due to its ease of fabrication, high precision

on curvature, roughness amplitudes much less than the wavelength range of light, and very low distortions arising from

thermal fluctuations due to low thermal expansion. The

AMSD program has shown that monolithic glass mirror

designs can be lightened to around 15 kg/m2 and their cost

reduced substantially compared to Hubble-based technology.

However, many believe that current glass technology is fast

approaching its theoretical limit, and that any further substantial reduction in areal density would require hybrid mirror

designs. It is anticipated that an order of magnitude increase in

fracture toughness, compared to monolithic glass, would be

required in large segmented mirror designs for unshielded

space-based applications. The high fracture toughness is desired

to avoid catastrophic mirror failure due to assembly accidents

and micrometeoroid impact damage. 

The future needs of the Air Force and the Department of

Defense will also require reductions in lead times, processing

costs, and launch load/weights. New materials and processing

methods for providing mirrors are needed. Mirror structural

substrates made out of advanced composite materials (metal,

ceramic, and polymer), foams, and microsphere arrays do allow

for CTE and modulus tailorability, low-density, high strength,

stiffness, and toughness. Small-size mirror structural substrates

have been fabricated from these new emerging materials, but

scaling to 2- or 3-meter segments will require new resources,

both in time and funding. Producing the surface finish and figure requirements needed for visible quality optics from these

multi-phase complex systems will be difficult. New methods of

polishing, replication, and sol-gel or polymer spinning will be

required to produce quality optical substrates. Finally, research

will be required to produce uniform, stress free reflective coatings and dielectric stacks on such large mirror systems. 

In this article we have addressed the majority of existing mirror technologies as well as the current research areas and their

related technologies. While not all-inclusive, our discussion

highlights the important issues relating to constructing large

mirrors for use in space. For additional information on specific programs and materials approaches we refer the reader to the

references.

NOTES AND REFERENCES

For further information on this topic, please forward inquiries

to the authors at either Lawerence.Matson@wpafb.af.mil or

David.Mollenhauer@wpafb.af.mil.

* λ denotes the wavelength of incident electromagnetic radiation (mainly visible light in the case of optical mirrors)

† Root Mean Square (RMS) – a statistical averaging function

that characterizes the relative roughness of a surface.

§ Thermal Stability (k/α) is defined as the thermal conductivity (k) divided by the coefficient of thermal expansion (α). In

SI units, k is expressed in W/m-K, α in ppm/K.

‡ Powder ceramics (and powder metals) are frequently

processed by pressing them into a preformed shape known as a

blank, or a green body. Green bodies are subsequently processed

to final form by densifying them (typically by sintering or vitrification).

**Fiber print-through is caused by a mismatch in properties

between the fibers and the resin matrix binding them together.

The resin will shrink upon cure while the fibers do not change

shape. The resin shrinkage and CTE/CME mismatch combine

to form valleys in the resin-rich zones between adjacent fibers,

resulting in significant surface roughness. This fiber printthrough phenomenon affects a composite mirror whether it

was polished or replicated except in the case where the use temperature is exactly the same as the manufacturing/or polishing

temperature.

[1] Private communication with Dr. Robert Jungquist of

Brashear LP; 615 Epsilon Drive, Pittsburgh, PA 15238

[2] Optical Materials, Vol. 1; Edited by S. Musikant; Marcel

Dekker Inc.; New York, 1990

[3] Air Force Research Laboratory, Materials Directorate

Contract F33615-01-C-5800 with MSNW Inc., San Marcos,

CA

[4] Schott Glass Technologies, Inc. Zerodur Glass Ceramics,

Doryea, PA (1982)

[5] NASA Tech Days website at http://optics.nasa.gov/

tech_days/tech_days_2002/index.html

[6] Private communications with Steve Kendric of Ball

Aerospace and Technologies. ‘skendric@ball.com’

[7] “AFRL Workshop on Advanced Mirrors for DOD

Applications” Sept. 2002, Kirtland AFB, NM

[8] R.B. McIntosh, Jr. and R.A. Paquin, “Chemical-mechanical

polishing of low-scatter optical surfaces”. Applied Optics 19, p.

2329 (1980)

[9] Stahl P., Presentation at the Optical Society of America

Conference, Optical Fabrication and Testing Session, Tucson

AZ, 2002

[10] “Silicon Carbide Technology for Large Submillimeter

Space Based Telescopes”, by F. Safa, F. Levallois, M. Bougoin,

and D. Castel, International Conference of Space Optics.

ICSO97, Toulouse, Dec. 1997

[11] “Silicon Carbide Technology for Submillimeter Space

Based Telescope”, by F. Safa, F. Levallois, M. Bougoin, and D.

Castel; 48th International Astronautical Congress, Turin,

October 1997

[12] Private communications with Eric Davenport of Coors

Tek, Boulder Colorado. [edavenport@CoorsTek.com]

[13] Private communications with Mark Earley of Xinetics Inc.

Devens MA

[14] M.A. Earley and S.M. Daigneault, “Silicon Deformable

Mirror Technology,” Proceedings of the Topical Meeting on

High Power Laser Optical Components, Oct 1989. Report No.

NWC TP 7080 Dec. 1990

[15] Private communications with William Goodman of

Schafer Corporation, Albuquerque NM. ‘wgoodman@schafer

alb.com’
74 The AMPTIAC Quarterly, Volume 8, Number 1

Dr. Lawrence E. Matson is the Direction Leader for both the Materials and Manufacturing Directorate’s Materials

Characterization Facility and the ML Mirror Team. His current mirror research is focused on producing materials for

a replicated, hybrid/composite mirror system. This includes fabricating both replicated nano-laminate/foil

facesheets from low coefficient of thermal expansion (CTE) materials and nano-sized, negative CTE dispersiods to

be used as CTE tailoring agents for mirror structural substrates and bonding agents. He holds a BS from Wright

State University, an MS from the Ohio State University, and a PhD from the University of Dayton. 

Dr. David H. Mollenhauer’s technical career has largely focused in the area of experimental mechanics. He

received a BS in Aerospace Engineering in 1990 from Texas A&M University. As part of the Air Force Palace Knight

program, he received his MS in Engineering Mechanics from Virginia Tech in 1992, and subsequently earned his

PhD there in 1997. During his tenure at AFRL, he developed a state-of-the-art photomechanics laboratory at the

Materials and Manufacturing Directorate. Using moiré interferometry, he has extensively examined the surface

behavior of drilled and molded holes in laminated composites; and quantitatively measured the micromechanical

behavior of model composite systems. The results of his work spawned great interest in a newly developed composite analysis method, known as BSAM. Contact information: david.mollenhauer@wpafb.af.mil.

[16] D.H. Mollenhauer and W.R. Ragland, “Analytical/

Experimental Examination of Syntactic Foams for Replicated

Mirrors,” Proceedings of the American Society for Composites

18th Technical Conference, Gainesville FL, Oct. 19-22, 2003

[17] S. Sihn, and B.P. Rice, “Characterization of Compressive

and Shear Properties of Carbon Foam and Its Core Application

in Sandwich Construction,” Proceedings of the American

Society for Composites 17th Technical Conference, 2002

[18] Private communications with James Cornie of Metal

Matrix Cast Composite of MA. ‘jcornie@mmccinc.com’

[19] J. Cornie, “A Review of Recent Metal Matrix Composite

Technology Spin-offs from MDA Projects”. Defense

Manufacturing Conference DMC-2002, Dallas TX, Dec. 2002

Recent US Patents

The following is a list of patents issued by the United States Patent and Trademark Office in the areas of materials and space. They

are organized into subject areas. Interested readers can obtain further information by accessing the Patent Office’s website:

http://www.uspto.gov.

SPACECRAFT STRUCTURES

6,679,456 Spacecraft protected by a coating including 

pyroelectric/ferroelectric particles, and the 

coating material

6,695,261 Shock isolation system for spacecraft fairing

POWER 

6,689,949 Concentrating photovoltaic cavity converters for

extreme solar-to-electric conversion efficiencies 

6,706,962 Hybrid solar cells with thermal deposited 

semiconductive oxide layer 

6,717,045 Photovoltaic array module design for solar electric

power generation systems

SENSORS

6,717,228 Infrared image sensor with temperature compensation

element 

6,678,048 Information-efficient spectral imaging sensor with TDI 

6,682,638 Film type solid polymer ionomer sensor and sensor cell 

6,714,345 Semiconductor optical amplifier providing high gain,

high power and low noise figure

6,704,138 Low-noise, high-power optical amplifier

6,717,544 Radar sensor

LAUNCH

6,695,256 Cryogenic propellant depletion system for a 

launch vehicle 

6,685,141 X-33 aeroshell and bell nozzle rocket 

engine launch vehicle 

THERMAL MANAGEMENT

6,689,471 Thermal management device and method of making

such a device 

6,711,904 Active thermal management of semiconductor devices 

THERMAL PROTECTION

6,716,539 Dual microstructure thermal barrier coating 

6,689,470 Thermal protection system

6,691,505 Fiber-reinforced rocket motor insulation

ROCKET

6,711,901 Rocket motor nozzle assemblies having vacuum 

plasma-sprayed refractory metal shell throat inserts,

methods of making, and rocket motors including same 

6,705,076 Rocket thrust chamber 

6,701,705 Gas-walled rocket nozzle 

6,679,965 Low density composite rocket nozzle components 

and process for making the same from standard 

density phenolic matrix, fiber reinforced materials 

6,673,449 Net molded tantalum carbide rocket nozzle

throat and method of making 

CRYOGENICS

6,681,589 Space suit backpack using solid adsorbents for 

cryogenic oxygen storage, freeze out of carbon 

dioxide and moisture, and ice heat sink 
The AMPTIAC Quarterly, Volume 8, Number 1 75

Although small, high-quality reflectors are available, adaptive

optics technology is not yet at a significantly mature stage to

resolve the optical aberrations usually observed in large inflated

membrane structures. Smaller reflectors on the order of one

meter or less in diameter have been produced in an affordable

tooling approach and the capability to produce precision tooling from conventional metallic materials has existed for some

time. However, as the size of the reflector increases, the cost of

preparing a stable metallic tool for manufacturing becomes cost

prohibitive. So from the programmatic viewpoint, the relatively high cost of conventional materials and processes has been a

limiting factor, especially for large aperture tooling. As an

example, one estimate for fabrication of a single four-meter

aperture tool made from conventional metallic materials and

finishing processes exceeded $3 million.

There are also technical challenges surrounding the use of

conventional metallic tooling materials. Materials compatibility is a typical issue when advanced polymers are produced upon

precision tools. A mismatch in the coefficient of thermal expansion (CTE) between the tool and the polymer thin film can

cause the film to pre-release during the thermal cure, tear during thermal cycling, or tear when releasing the film from the

metal tool; even when release agents have been properly

applied.

Another technical challenge involves tool dynamics. SRS and

AFRL/ML (referred to hereafter as ML) digital photogrammetry shape measurement tests of conventional tooling have

shown that significant tool shape changes can occur as a result

of thermal cycling. Some of this movement can be classified as

thermal stress relief, but some movement is clearly dynamic.

These tool shape changes ultimately produce off-dimension

thin films that degrade the performance of the application as

well as the future utility of the tool. In some cases, the tool surface coating (over conventional materials) degrades (cracks) as

a result of thermal cycling and CTE mismatches. The innovative tooling research and development in progress addresses the

programmatic and technical issues through the selection of

materials and fabrication processes that are CTE compatible,

have demonstrated the ability to retain their precision shape

and surface quality through repeated thermal cycles, and are

cost effective.

FABRICATION OF PRECISION

TOOLING

Over the last 15 years, SRS has partnered with ML and NASA to develop

thin film polymer concentrator technology, including the respective 

precision tooling. In recent years,

AFRL has focused an increasing

James C. Pearson

SRS Technologies

Huntsville, AL

Marilyn R. Unroe

Materials and Manufacturing Directorate

Air Force Research Laboratory

Wright-Patterson AFB, OH

Any discussion about placing new systems into space is at its heart, a discussion of the respective enabling technologies which facilitate or

make possible their deployment. Among these enablers are materials, manufacturing methods, and fabrication facilities. State-of-the-art

tooling is part of any up-to-date facility. This article describes the innovative design approach, fabrication processes, and materials that

SRS Technologies and the Air Force Research Laboratory’s Materials and Manufacturing Directorate (AFRL/ML) are developing for 

precision tooling directed toward the production of polymer-based thin film (membrane) reflectors for use as inflatable light weight 

antennas, solar power concentrators, and imaging optics. The designs and tooling will ultimately support production of polymer thin films

for a proposed Picosat Space Inflatable Reflector Experiment (PSIREX) space flight experiment (FE) and ongoing terrestrial radio frequency

(RF) antenna development and testing. The precision tools are also being investigated for use as substrates for the lay-up of composite 

structures.

Figure 1. One-Meter by Two-Meter 

Off-Axis Mandrel and Thin Film

Polymer Concentrator.
amount of attention upon the development of precision tooling technology. In 2001, SRS and ML successfully developed

a prototype 1-m x 2-m off-axis* tool that was utilized to produce precise, tear-resistant polymer films. Figure 1 shows the

prototype tool and the thin film concentrator that was manufactured. An outgrowth of that technology has been the

ongoing fabrication of a 4-m x 6-m off-axis tool. Figure 2

shows the tool along with prototype test hardware. The latest

effort is a 1-m on-axis* tool for the proposed PSIREX FE.

The PSIREX FE mandrel and test hardware are shown in

Figure 3. Fabrication of these tools should be completed in

CY03.

DESIGN AND MATERIALS

The design of the PSIREX FE tool began with a general

description of the FE and an initial set of tooling specifications,

as shown in Figure 4. The fundamental objective of the

PSIREX FE is to verify deployment of a thin film concentrator

and then characterize the shape of the reflective surface. One 

of the explicit specifications was a 1m-class aperture. A grid

projection optical measurement apparatus (similar to a Ronchigram or Hartmann) was selected for the FE shape characterization technique, which led to the selection of a spherical-shaped

reflective surface. A focal number† target value of 0.5 was

selected based on geometric ray trace analysis of the grid projection apparatus. By using the FE specifications, an approach

and set of tooling requirements were spawned.

As a tooling design objective, non-conventional materials

were selected that met the tooling requirements and addressed

the programmatic and technical issues. An aluminum skin honeycomb was chosen as the base because of its inherent flatness

and strength and relatively high temperature rating, and a high

temperature honeycomb material (Nomex) was chosen for the

tool core. This material has a combination of high compressive

stress, high temperature rating, and good dimensional stability.

A specially formulated anhydride was chosen as the substrate

material. This thermosetting polymer material possesses good

properties for machining and surface finishing. A specially formulated high temperature phenolic resin was chosen for the

surface coating material. The performance of each of these

materials (individually and assembled) was proven through

prototyping and extensive testing. Finally, a tooling design with

material and surface finish specifications was created.

Depending on the application, SRS sometimes employs a

tooling shape optimization algorithm that calculates the anticipated deformations of the deployed thin film concentrator

that result from thermal and pressure loads. These deformations are then compensated for in the shape of the tool. The

optimization process is further described in Figure 5. The

PSIREX tool was not optimized in order to support correlation

of the film shape characterization data with the mandrel shape.

FABRICATION PROCESS

Following design and materials selection, the fabrication

process for the PSIREX tool was defined and implemented

with only minor deviations from the original process. The

PSIREX tool process itself is a deviation from the similar

process currently being used in fabrication of the large aperture

76 The AMPTIAC Quarterly, Volume 8, Number 1

Figure 2. Four-Meter by Six-Meter Off-Axis Tool and

Prototype Test Hardware.

Figure 3. One-Meter

On-Axis Mandrel and

Proposed PSIREX FE

Test Hardware.

Figure 4. PSIREX FE Specifications.

PSIREX FE Success Criteria

Probability of

Technical Success

Probability of

Follow-on

Research

Technical

Difficulty

nominal

value

added advanced

Qualitative & Quantitative Data

Imaging Test

Qualitative Data Non-imaging Test

Ronchi & Hartmann

Grid

Projection

Assess

Mirror

Verify

Deployment Power On &

Verify Link Deploy PICOSAT

from LV
(4-m x 6-m) off-axis tool being fabricated for AFRL (shown in

Figure 2). The differences between the PSIREX tool process

and the 4-m x 6-m tool process are primarily due to scaling

issues. The mass, outside dimensions and handling requirements of the 4-m x 6-m tool require a more robust base; plus,

the PSIREX has an objective to eliminate all hand finishing of

the surface, whereas the 4-m x 6-m tool seeks to minimize hand

finishing. The 4-m x 6-m tool is currently in the surface 

finishing phase and the PSIREX tool is operational, as shown

in Figure 2.

DESIGN VERIFICATION

During the PSIREX tool fabrication process, the design was

verified via shape and surface finish measurements at three 

critical junctures: a rough verification following machining of

the core, a coordinate measurement machine (CMM) measurement following substrate machining, and a second CMM

measurement, following surface coating and two thermal

cycles. Surface measurements, via a profilometer, were specified

following substrate machining and surface coating. SRS also

uses digital photogrammetry to perform shape measurement of

both tools and concentrators.

ISSUES

The cost effectiveness of the precision tooling that is manufactured using non-conventional materials and respective 

processes will be assessed versus historical data for conventional tooling and a market assessment to see what customers are

willing to pay for the precision tooling. Certainly, the design

and quantity of material for each unit is unique and multiple

orders per design are not anticipated. However, for this given

design approach, the material selection, and fabrication

process would be identical for each tool. This would also be

true for conventional tooling approaches; hence the anticipated reductions in material cost, fabrication labor, and machine

time should translate into lower tooling costs.

The ultimate qualifier for these tools is the polymer thin film

product that is produced. Based on the prototype results, the

research and development team expects the polymer thin films

to meet the design requirements. To be sure, a correlation will

be developed between the quality of polymer thin film product

and the tool to verify the film fabrication process.

FUTURE TOOLING

The advent of new materials will undoubtedly benefit the precision tooling market. Advanced polymers will require increasingly precise tooling with longer life and a broader temperature

range to support thermal cycling. Coated polyimide foams are

currently being developed as core/substrate materials that are

very lightweight and perform at high temperatures. New

“designer” polymers will be tougher and more tear resistant,

while also allowing more control of the CTE. Thermal cycling

requirements may be reduced through the development of

chemical curing technology. Finally, the whole tooling concept

may be redefined, if the manufacture of precision pressureformed thin films proves feasible.

SUMMARY / CONCLUSIONS

Based on the prototype results and the fabrication cost of the

PSIREX tool, the innovative materials and processes that are

being developed are cost effective, programmatically capable,

and technically compatible. The research and development team

anticipates that the tool will maintain precision shape through

multiple thermal cycles. They further anticipate that the surface

finish will not significantly degrade through a typical duty cycle

of 25 thermal cycles and that polymer films will not tear during

thermal cycling and will release properly. The greater than

300°F maximum soak temperature is sufficient for many applications, but higher soak temperatures are still needed.

NOTES AND REFERENCES

* Concentrators, reflectors, and other precision components

with circular geometries are fabricated on rotating fixtures

known as mandrels. Until recently, most of these components

were affixed to their tools via the mandrel’s centrally rotating

shaft – these are known as on-axis tools. The major benefit of

on-axis tooling is that it ensures radial symmetry for components manufactured in this manner. However, the presence of a

central axle may interfere with the manufacture of some precision components, so an alternate fixturing method of mounting circular components (concentrically around the perimeter)

was developed where there was no axial interference. These fixtures are known as off-axis tools, which while mounted coaxially relative to the fixture’s rotation, are not affixed to an axle.

Note in Figure 2 there are several examples of off-axis mountings – either by riggings or rollers. Conversely, Figure 3 shows

a more traditional on-axis tool.

† The focal number is the ratio of the focal length (distance

from the lens surface to the focal point) to the aperture 

diameter. It is also known as the focal ratio, or the focal stop

(commonly known in photography as the f-stop).

J.C. Pearson, D.M. Lester, Michael R. Holmes, W.A. Wong,

“Solar Thermal Vacuum Testing of an Integrated Membrane

Concentrator System at the NASA GRC Tank 6 Facility”, presented at the STAIF02 conference, Albuquerque, NM,

February, 2002

The AMPTIAC Quarterly, Volume 8, Number 1 77

Figure 5. Optimization Process.

• 3-D Brick FEA Model including Thermal and Pressure

Loads was created in order to

determine if Thermal Strains

are significant enough that 

the FSC Design must be 

optimized for Thermal as 

well as Pressure Loads

• Thermal FSC Model includes

Canopy and Reflector Films

because on-orbit temperatures

vary between films

• Current Thermal FSC 

Model analyzed at current

on-orbit temperature data

(Canopy at -14.44°C,

Reflector at 49.44°C) • Thermal Loads result in 

significant deflections

• Concentrator must be 

optimized for Thermal 

Loads as well as 

Pressure Loads

Baseline Model Parameters

3588 Type 25 elements

Pressure: 0.013” H2O

Film Thickness: 0.001”

Create Shell

Model and verify

model through

test article 

photogrammetry

Optimize model for pressure

effects and verify method through

shape and slope error analysis

Create

FSC Shell

Model

Use optimization techniques to

optimize model for pressure effects

Generate 3D

Canopy/Reflector

Brick Model

including both

pressure and 

thermal loads

Slope Error Comparison 

along Major Axis Slope Error Comparison 

along Major Axis

Slope Error

Slope Error

X Position X Position

0.003 rad Reference Design

Optimized

Design

Baseline

Design

60 40 20 0 -20 -40 -60 -80

0.015

0.010

0.005

0.003

0

0.003 rad Reference Design

Optimized

Design

Baseline

Design

150 100 50 0 -50 -100 -150

0.0040

0.0035

0.0030

0.0025

0.0020

0.0015

0.0010

0.0005

0
AMPTIAC SENDS ITS THANKS

78 The AMPTIAC Quarterly, Volume 8, Number 1

Mr. James C. Pearson, Jr. is an Aerospace Engineer with SRS Technologies, Systems Technology Group, Huntsville

AL. He is the Principal Investigator for USAF Contract F33615-02-C-5011 under which SRS is developing affordable tools and processes for fabricating large aperture thin film deployable structures. He has over 20 years 

experience designing, fabricating, and testing large thin film polymer structures for ground and space based 

applications.

Ms. Marilyn R. Unroe is a Program Manager and Research Chemist with the Polymer Branch of the Materials 

and Manufacturing Directorate (AFRL/MLBP), Wright-Patterson AFB, OH. She has over 20 years experience in

basic polymer synthesis research and applied development of polymeric materials for aircraft and aerospace 

applications, all within the Polymer Branch of AFRL.

D.M. Cannon, B.B. McWhorter, and D.M. Lester, “Materials

for Space Formed, Rigidized Structures Used in Solar Thermal

Propulsion Concepts: National Space and Materials

Symposium, San Diego, CA, 27 February 2000 - 2 March

2000

P.A. Gierow, D.L. Lester, and J.C. Pearson, “Near Term InSpace Demonstration of an Inflatable Concentrator,” AIAA99-1073, AIAA, 37th Aerospace Sciences Meeting and Exhibit,

Reno, NV, January 10-14, 1999

M.R. Holmes, D.L. Lester, J.C. Pearson, and S.R. Wassom,

“Solar Thermal Propulsion IHPRPT Demonstration

Program,” AIAA-2000-5109, AIAA Space 2000 Conference

and Exposition, Long Beach, CA, September 19-21, 2000

M.R. Holmes, R.E. Partch, and J.C. Pearson, “Inflatable

Concentrator Performance Characterization for Space Flight

Applications,” RAES99-7701, ASME, Lahaina, Maui, HI,

1999

Don’t let your work become part of a

landfill – you’ve no doubt seen this message in past issues of the AMPTIAC

Quarterly. It is part of our continuous

campaign to preserve and maintain the

invaluable and irreplaceable material

research data generated by the DOD,

contractors, and universities over the previous decades. We regularly receive

donations from principals in the defense

materials communities, which augment

and enhance AMPTIAC ’s library of now

more than 220,000 volumes. As a semiregular feature of the newsletter, we

would like to acknowledge some of the

exceptional donations made by our readers. By donating their work to us, it now

also becomes available to you – the users

of the community!

Mr. Henry Johnson donated over 630 reference books, government

technical reports and other documents. Mr. Johnson is an alumnus of

the Air Force Research Laboratory’s Materials and Manufacturing

Directorate, who last served as the Chief of the Metals Branch in the

1980’s (AFRL/MLLM). After he left the Air Force, Mr. Johnson continued to advocate metallic materials as the President of Materials &

Manufacturing Tech Inc. Throughout his career, he collected many 

documents on metallic materials. Many of the reports he donated date

back to the early days of the space race. These documents will help the

government save both time and money by preserving valuable data. 

Dr. Stuart Schwab has quite generously donated over 70 reports from

his personal technical library, accrued over the past 20 years as an active 

participant in the development of defense technology. The reports 

Dr. Schwab forwarded focus on high temperature composite materials,

both ceramic matrix and carbon-carbon composites, for advanced engine

and thermal protection applications.
The AMPTIAC Quarterly, Volume 8, Number 1 79

AMPTIAC Directory

Government Personnel

TECHNICAL MANAGER/COTR

Dr. Lewis E. Sloter II

Staff Specialist, Materials & Structures

ODUSD(S&T)/Weapons Systems

1777 North Kent St, Suite 9030

Arlington, VA 22209-2110

(703) 588-7418, Fax: (703) 588-7560

Email: lewis.sloter@osd.mil

DEFENSE TECHNICAL INFORMATION CENTER

(DTIC) POC

Melinda Rozga, DTIC-AI

8725 John J. Kingman Rd, Suite 0944

Ft. Belvoir, VA 22060-6218

(703) 767-9120, Fax: (703) 767-9119

Email: mrozga@dtic.mil

ASSOCIATE COTRS

ORGANIC STRUCTURAL MATERIALS & 

ORGANIC MATRIX COMPOSITES

Roger Griswold

Division Chief

US Air Force

AFRL/MLS

2179 Twelfth St, Bldg 652

Wright-Patterson AFB, OH 45433-7702

(937) 656-6052, Fax: (937) 255-2945

Email: roger.griswold@wpafb.af.mil

ENVIRONMENTAL PROTECTION

& SPECIAL FUNCTION MATERIALS

Dr. James Murday

Naval Research Laboratory

4555 Overlook Ave, S.W. Code 6100

Washington, DC 20375-5320

(202) 767-3026, Fax: (202) 404-7139

Email: murday@nrl.navy.mil

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