National Aeronautics and Space Administration

A Researcher’s Guide to:

Space Environmental Effects
2

This International Space Station (ISS) Researcher’s Guide is published by 

the NASA ISS Program Science Office.

Authors: 

Miria M. Finckenor

Kim K. de Groh

Executive Editor: Amelia Rai

Technical Editor: Neesha Hosein

Designer: Cory Duke

Cover and back cover: 

a. Astronaut Stephen Robinson retrieves Materials International Space Station Experiment-1 (MISSE-1) 

during the STS-114 mission, July 2005. Photo was taken during Expedition 6 on the ISS. 

b. Scanning electron microscopy image of the back-surface of a thin film layer of Kapton H after four 

years of low-Earth orbit (LEO) ram atomic oxygen erosion while on the exterior of the ISS as part of 

the MISSE-2 Polymer Erosion and Contamination Experiment (PEACE). 

c. Scanning electron microscopy image showing the microscopic cone texture that developed after 

four years of LEO ram atomic oxygen erosion of a pyrolytic graphite sample while on the exterior of 

the ISS as part of the MISSE-2 PEACE Polymers experiment. 
Close-up view of the Materials International Space Station Experiment (MISSE) 6A and 6B Passive Experiment Containers 

(PECs) on the European Laboratory/Columbus. Photo was taken during a flyaround of STS-123 Space Shuttle Endeavor.

Soaring 250 miles above Earth, the ISS is a modern wonder of the world, 

combining the efforts of 15 countries and thousands of scientists, engineers 

and technicians. The ISS is a magnificent platform for all kinds of research 

to improve life on Earth, enable future space exploration and understand the 

universe. This researcher’s guide is intended to help potential researchers 

plan experiments that would be exposed to the space environment, while 

externally attached to or deployed from the ISS. It covers all the pertinent 

aspects of the space environment, how to best translate ground research to 

flight results and lessons learned from previous experiments. It also details 

what power and data are available on the ISS in various external locations.

The Lab is Open
4
1. Microgravity, or weightlessness, alters many observable phenomena 

within the physical and life sciences. Systems and processes affected by 

microgravity include surface wetting and interfacial tension, multiphase 

flow and heat transfer, multiphase system dynamics, solidification, and 

fire phenomena and combustion. Microgravity induces a vast array of 

changes in organisms ranging from bacteria to humans, including global 

alterations in gene expression and 3-D aggregation of cells into tissue-like 

architecture.

2. Extreme conditions in the ISS environment include exposure to extreme 

heat and cold cycling, ultra-vacuum, atomic oxygen, and high energy 

radiation. Testing and qualification of materials exposed to these extreme 

conditions have provided data to enable the manufacturing of longlife reliable components used on Earth as well as in the world’s most 

sophisticated satellite and spacecraft components.

3. Low Earth orbit at 51degrees inclination and at a 90-minute orbit 

affords ISS a unique vantage point with an altitude of approximately 240 

miles (400 kilometers) and an orbital path over 90 percent of the Earth’s 

population. 

This can provide improved spatial resolution and variable lighting 

conditions compared to the sun-synchronous orbits of typical Earth 

remote-sensing satellites.

Unique Features of the ISS 

Research Environment

1. Microgravity, or weightlessness, alters many observable phenomena 

within the physical and life sciences. Systems and processes affected by 

microgravity include surface wetting and interfacial tension, multiphase 

flow and heat transfer, multiphase system dynamics, solidification, and 

fire phenomena and combustion. Microgravity induces a vast array of 

changes in organisms ranging from bacteria to humans, including 

global alterations in gene expression and 3-D aggregation of cells into 

tissue-like architecture.

2. Extreme conditions in the ISS space environment include exposure to 

extreme heat and cold cycling, ultra-vacuum, atomic oxygen, and high 

energy radiation. Testing and qualification of materials exposed to these 

extreme conditions have provided data to enable the manufacturing of 

long-life reliable components used on Earth as well as in the world’s most 

sophisticated satellite and spacecraft components.

3. Low-Earth orbit at 51 degrees inclination and at a 90-minute orbit 

affords ISS a unique vantage point with an altitude of approximately 240 

miles (400 kilometers) and an orbital path over 90 percent of the Earth’s 

population. This can provide improved spatial resolution and variable 

lighting conditions compared to the sun-synchronous orbits of typical 

Earth remote-sensing satellites.

Unique Features of the ISS 

Research Environment
6

Unique Features of the ISS Research Environment 5

Research Priorities for Space Environmental Effects on the ISS 7

Aspects of the Space Environment 10

Vacuum 10

Atomic Oxygen 10

Ultraviolet Radiation 12

Particulate or Ionizing Radiation 13

Plasma 14

Temperature Extremes and Thermal Cycling (Coefficients of Thermal 

Expansion [CTE] Mismatch) 14

Micrometeoroid/Orbital Debris Impact 15

Orientation and Location on the ISS 17

Approaches to Mitigate Contamination 18

Lessons Learned 20

Plan for Flight Recovery Contingencies 20

Control Samples and Preflight Testing 20

Understand Sample Geometry 21

Other Lessons Learned 22

Developing and Flying Research to the ISS 24

ISS External Accomodations 26

Japanese Experiment Module Exposed Facility [Japan Aerospace 

Exploration Agency (JAXA)] 26

Multi-Purpose Experiment Platform (MPEP [JAXA]) 27

JEM Small Satellite Orbital Deployer (J-SSOD [JAXA]) 28

EXPRESS Logistics Carrier (NASA) 28

Columbus External Payload Facility (European Space Agency) 31

Russian Segment External Facilities (Russian Federal Space Agency) 32

Funding Opportunities 33

Citations 34

Acronyms 37

Table of Contents
Scientists and engineers have developed advanced materials for manned spacecraft 

and satellites for a range of sophisticated applications in space exploration, 

transportation, global positioning and communication. The materials used on the 

exterior of spacecraft are subjected to many environmental threats that can degrade 

many materials and components. These threats include vacuum, solar ultraviolet 

(UV) radiation, charged particle (ionizing) radiation, plasma, surface charging and 

arcing, temperature extremes, thermal cycling, impacts from micrometeoroids and 

orbital debris (MMOD), and environment‐induced contamination. In terms of 

materials degradation in space, the low-Earth orbit (LEO) environment, defined 

as 200-1,000 km above Earth’s surface, is a particularly harsh environment for 

most non-metallic materials, because single-oxygen atoms (atomic oxygen [AO]) 

are present along with all other environmental components (Yang and de Groh, 

2010). Space environmental threats to spacecraft components vary greatly, based 

on the component materials, thicknesses and stress levels. Also to be considered 

are the mission duration and the specific mission environment, including orbital 

parameters for the mission, the solar cycle and solar events, view angle of spacecraft 

surfaces to the sun and orientation of spacecraft surfaces with respect to the 

spacecraft velocity vector in LEO (Dever et al., 2005). Examples of AO erosion 

and radiation-induced embrittlement of spacecraft materials are provided in 

Figures 1 and 2.

Research Priorities for 

Space Environmental 

Effects on the ISS

Figure 1. Preflight and postflight Long Duration Exposure Facility M0001 Heavy Ions in Space experiment, indicating 

atomic oxygen erosion and ultraviolet degradation.

Preflight Postflight
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Determining how long-term exposure to space conditions impacts various 

materials and, thus, which materials are best suited for spacecraft construction 

can most effectively be accomplished through actual testing in space. Although 

space environment effects testing can be conducted in ground-laboratory facilities, 

ground facilities often do not accurately simulate the combined environmental 

effects and, therefore, do not always accurately simulate performance or degradation 

observed in the space environment. The aspects of the space environment section 

of this document discusses each aspect of the space environment and what ground 

simulation methods translate best to actual flight results. However, the synergism 

of all the elements of the space environment is difficult to duplicate on the ground. 

Therefore, actual spaceflight experiments provide the most accurate spacecraft 

durability data. Materials spaceflight experiments to evaluate the environmental 

durability of various materials and components in space have been conducted since 

the early 1970s, including 57 experiments on the Long Duration Exposure Facility 

(LDEF), which was retrieved in 1990 after spending 69 months in LEO (de Groh 

et al., 2011). The ISS provides an ideal platform for long-term space environment 

effects testing, particularly as experiments can be returned to Earth for postflight 

analyses. The Materials International Space Station Experiment (MISSE) is a series 

of materials flight experiments, the first two of which were delivered to the ISS 

Figure 2. Space-exposure damage to Hubble Space Telescope multilayer insulation.

Large radiation-induced cracks in the 

outer layer of multilayer insulation after 

6.8 years of space exposure (Townsend 

et al., 1999).

Severe degradation to the aluminized-Teflon® outer layer of multilayer 

insulation after 19 years of space exposure (Yang and de Groh, 2010).
during STS-105 in 2001 (de Groh et al., 2008; de Groh et al., 2009; de Groh et al., 

2011). Consisting of a pair of trays hinged together like a suitcase (called a Passive 

Experiment Container [PEC]) and containing an array of individual experiments, 

the PEC is attached to the exterior of the ISS, providing long-duration exposure 

to space conditions. In the MISSE suite (MISSEs 1 through 8), 10 PECs (and one 

smaller tray), together containing thousands of samples, have been flown in various 

external locations on the ISS. With participants from NASA, the Department of 

Defense, industry, and academia, MISSE is the longest running multi-organization 

technology, development and materials testing project on the ISS. It has provided 

many tangible benefits for the agency and its partners, with the flight data 

impacting many space programs. Like LDEF, MISSE flight data have provided a 

great wealth of space performance and environmental durability data and many 

lessons learned beneficial to further investigators (Banks et al., 2008). Published 

MISSE data, photographs and some raw data files are being gathered in a MISSE 

database in the Materials and Processes Technical Information System (MAPTIS).

All prospective experimenters should register for an account on MAPTIS, found at 

http://maptis.nasa.gov/. MAPTIS contains a wealth of information for designers 

and materials engineers, particularly the Materials Selection Database. This is 

a useful reference to consult before building hardware so that safety, structural, 

pressure vessel and line, fracture-critical and contamination requirements are met. 

This database is open to all registered users and holds 50 years of analyzed results of 

tests conducted on metallic and nonmetallic materials:

• Metals data include analyzed results of tests relevant to corrosion, crack growth, 

creep rupture, flammability, fluid compatibility, fracture mechanics, frictional 

heat, high-cycle fatigue, low-cycle fatigue, mechanical impact, particle impact, 

pneumatic impact, promoted ignition, stress corrosion, and tensile strength. 

A list of manufacturers is also provided.

• Nonmetals data consist of test results, including flammability, fluid 

compatibility, odor, outgassing (thermal vacuum stability), toxicity (offgassing) 

and vacuum condensable material compatibility with optics (VCMO). 

Materials data relevant to spacecraft design may also be found in Silverman (1995). 

This guidebook gathered many materials experiments’ results from the LDEF, shortduration shuttle flights and selected ground simulations.
10

Vacuum 

The hard vacuum of space (10-6 to 10-9 torr) will cause outgassing, which is the 

release of volatiles from materials. The outgassed molecules then deposit on lineof-sight surfaces and are more likely to deposit on cold surfaces. This molecular 

contamination can affect optical properties of vehicle and payload surfaces and 

spacecraft performance, particularly for sensitive optics. To mitigate this problem, 

the ISS has specified in NASA SSP 30426, Space Station External Contamination 

Control Requirements, what the limits are for molecular deposition, induced 

molecular column densities and the release of particulates.

An investigator must compile a list of all materials used in a flight experiment 

and submit this list for a timely review. NASA maintains a database of test results 

from ASTM E1559, Standard Test Method for Contamination Outgassing 

Characteristics of Spacecraft Materials, and the older ASTM E595, Standard Test 

Method for Total Mass Loss and Collected Volatile Condensable Materials from 

Outgassing in a Vacuum Environment. Material identification, location, vacuum 

exposed surface area, operating temperature range and condensable outgassing 

rate data are used by the ISS Program’s return flux model to calculate any impact 

to the vehicle (Soares and Mikatarian, 2003). A material known to outgas should 

be thermal vacuum baked for a minimum of 24 hours at a temperature above that 

expected in orbit or, if that is not known, at 100 °C. Assemblies may be thermal 

vacuum baked prior to flight. 

This is one component of the space environment that translates well between 

ground simulation of 10-6

 to 10-9 torr vacuum and flight.

Atomic Oxygen 

AO is produced when short-wavelength UV radiation reacts with molecular 

oxygen in the upper atmosphere. It is the most significant component of the space 

environment at ISS altitude in terms of material degradation. AO oxidizes many 

metals, especially silver, copper, and osmium. AO reacts strongly with any material 

containing carbon, nitrogen, sulfur and hydrogen bonds, meaning that many 

polymers react and erode. Polymers containing fluorine, such as Teflon®, react 

synergistically, meaning that the reactivity to AO increases with longer exposure 

to UV radiation (Pippin et al., 2004). Even materials with AO protective coatings 

can degrade because of AO undercutting erosion at protective coating defect sites 

(Figure 3; Banks et al., 2004).

Aspects of the Space 

Environment
The AO exposure, characterized by the fluence (atoms/cm2) to an experiment will 

depend not only on the orientation and spacecraft altitude but also on the solar 

activity at the time of flight. There can be an order of magnitude difference in 

AO flux between solar maximum (higher flux) and solar minimum (lower flux). 

The current solar cycle at time of publication, Solar Cycle 24, promises to be the 

least active cycle in the last hundred years (Hathaway, 2013). An experiment to 

determine AO reactivity, also known as erosion yield (cm3/atom), of materials needs 

an exposure that is long enough for measureable erosion to occur. Software such as 

the Mass Spectrometer Incoherent Scatter (MSIS) Model can be used to estimate 

the AO fluence for a particular mission, which is then used to determine if the 

experiment will receive a high enough level of AO for the desired measurements 

and also whether the sample thickness is sufficient if the expected fluence is high.

Translating ground simulation results to flight depends heavily on the source of 

AO used in the laboratory. AO in orbit is ~5.2-eV energy, principally as a result of 

the ISS orbital velocity. Plasma ashers create AO but also produce heating, intense 

Lyman-alpha UV radiation and a significant percentage of ions rather than just 

neutral atoms. In addition, the arrival direction of AO may vary more than in 

space. These differences may give erroneous results, particularly for polymers with 

low glass transition temperatures. Stambler (2011) compared ground simulation 

to flight results for 40 different polymers, and the erosion yield in the plasma 

asher was higher than that in space for every material. Thermal energy AO sources 

eliminate the heating problem but do not have the necessary energy to break some 

chemical bonds and may need a much higher fluence to replicate the same erosion 

in orbit. 

AO beam facilities may use laser-detonation or microwave sources to generate AO 

at energies close to that at ISS altitude. For polymers that do not contain fluorine, 

simulations in these facilities are generally close to flight, with the AO reactivity 

Figure 3. Image shows atomic oxygen undercutting degradation of the solar array wing blanket box cover on the International 

Space Station after one year of space exposure.
12

being ±10 percent. Polymers with fluorine are more sensitive to the UV generated 

simultaneously with the AO and consequently have a higher reactivity in the beam 

facility than observed in orbit.

In general, laboratory results are normalized to flight results by flying materials 

with known reactivity to AO in addition to the test samples. Kapton® H and 

Kapton® HN polyimide are the most common materials selected for this purpose; 

polyethylene, polypropylene, and pyrolytic graphite have also been used. Following 

ASTM E2089, Standard Practices for Ground Laboratory Atomic Oxygen 

Interaction Evaluation of Materials for Space Applications, can reduce variability in 

results.

Ultraviolet Radiation 

Earth’s atmosphere filters out most of the sun’s damaging light, but ISS materials 

bear the brunt of solar photon damage. While AO may bleach materials, UV 

generally darkens them (Figure 4), particularly in the presence of contamination. 

UV radiation also damages polymers by either cross-linking (hardening) or chain 

scission (weakening). UV under high vacuum can also create oxygen vacancies in 

oxides, leading to significant color changes.

Normalizing ground simulations to flight results is very dependent on the material 

and the type of lamp used, whether deuterium, mercury, or mercury-xenon and 

the intensity of the lamp (number of suns). Accelerated UV testing may also cause 

material heating, so the accuracy of a simulation may depend on whether the sample 

temperature is controlled or not. Any UV exposure on the ground should be done 

with the samples in vacuum. Also, any near-UV source such as a mercury-xenon 

Figure 4. Preflight and postflight images of the Optical Properties Monitor shown with ultraviolet-darkened insulation 

after nine months of exposure on the Mir Space Station.

Preflight Postflight
lamp that is not integrated into the vacuum system should illuminate the samples 

through a UV-grade window. For exposure of metal oxides, consider the ability 

to make in-situ reflectance measurements after vacuum UV ground exposures, as 

atmospheric oxygen has been observed to reverse the effects of oxygen vacancies.

Particulate or Ionizing Radiation 

The three main sources of charged particle radiation naturally occurring in space 

are galactic cosmic rays, solar proton events, and the trapped radiation belts. For 

most materials on the ISS, the effects of AO and UV can overshadow any effects by 

particulate radiation. Depending on the polymer, particulate radiation can result 

in cross-linking or chain scission, similar to damage by UV, resulting in polymer 

embrittlement. A greater effect is seen in avionics, namely single-event upsets, 

bit errors and latchups. This can be mitigated by selecting avionics that are “rad 

hard” or by placing shielding around the electronics. Radiation effects may also 

be alleviated by using error-correction circuitry and triple-module redundancy, 

where two good process results “outvote” a corrupted one. Most of the particulate 

radiation dose to the ISS occurs when flying through the South Atlantic Anomaly 

(Figure 5) and is in the form of electrons, although proton dose can be significant 

for some materials and components.

Laboratory results can be normalized to flight results by understanding the dosedepth profile of radiation for the maximum effect on that material. For example, 

radiation testing of a thermal control coating that is 3 to 5 mil thick could include 

electrons and protons in the 40- to 700-keV range. Particles in this energy range 

would have the most effect on optical properties and coating chemistry, while 

higher energy particles would pass through the coating and affect the substrate. 

Figure 5. Charged particle map shows the South Atlantic Anomaly. [Credit: S.L. Snowden, http://heasarc.gsfc.nasa.gov/

docs/rosat/gallery/misc_saad.html]
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Conversely, higher-energy particles (in the 1- to 70-MeV range) have been used to 

study single-event upsets (Swank et al., 2008), whereas the lower-energy particles 

would not have much of an effect.

Plasma

Similar but separate from the higher energy particulate radiation is the space 

plasma. The plasma environment around the ISS is composed of approximately 

equal amounts of positively charged oxygen ions (O+) and free electrons and varies 

with solar activity and altitude. Because of the differences in spacecraft velocities, 

ion thermal energy and electron thermal energy, electrons can impact any spacecraft 

surface, while ions can only impact ram (leading edge) surfaces. This can lead to 

a negative charge buildup, which can lead to ion sputtering, arcing and parasitic 

currents in solar arrays, as well as reattraction of contamination (James et al., 1994).

Examples of experiments that would be affected by the plasma environment would 

be high-voltage solar arrays and conductive coatings designed to bleed off static 

charge on a spacecraft. A principal investigator for such an experiment would need 

to work in conjunction with the Spacecraft Charging Assessment Team at NASA’s 

Johnson Space Center (JSC). Specific information about the plasma environment 

around the ISS can be provided through operation of the Floating Potential 

Measurement Unit. ISS plasma conditions may be modeled in the laboratory with 

a hollow cathode plasma source to create a low-density (106/cm3), low-temperature 

(≤1‐eV electron temperature) plasma. Argon is often used, but some sources may 

use xenon, oxygen or helium.

Temperature Extremes and Thermal Cycling (Coefficients of 

Thermal Expansion [CTE] Mismatch)

As the ISS moves in and out of sunlight during its orbit around Earth, the degree 

to which a material experiences thermal cycling temperature extremes depends on 

its thermo-optical properties (solar absorptance and thermal emittance), its view of 

the sun, its view of Earth, its view of other surfaces of the spacecraft, durations of 

time in sunlight and in shadow, its thermal mass and the influence of equipment 

or components that produce heat (Dever et al., 2005). A rule of thumb for these 

cyclic temperature variations is -120 °C to +120 °C, but high solar absorptance with 

low infrared emittance will contribute to greater temperature swings. Large areas 

of material with poor thermal properties may not be allowed on the ISS because of 

exceeding touch temperature limits for the astronauts’ gloves.
Protective coatings may degrade in the ISS environment if there is a mismatch in 

CTE between the coating and the substrate. Sixteen thermal cycles a day (the ISS 

orbits Earth approximately once every 92 minutes) may lead to cracking, peeling, 

spalling or formation of pinholes in the coating, which then allows AO to attack 

the underlying material (Figure 6).

Micrometeoroid/Orbital Debris Impact

All areas of a spacecraft may be impacted by micrometeoroids traveling as fast 

as 60 km/s. Surfaces facing the ram direction are more likely than those in the 

wake direction to be hit with space debris, traveling at an average velocity of 

10 km/s. Space debris varies with the solar cycle: as the Sun’s activity increases, the 

atmosphere heats up, increasing the drag on space debris in orbit. Large space debris 

is tracked so that the ISS can perform avoidance maneuvers, but there is no current 

way to avoid small debris impacts. Most of the impacts on returned experiments 

have been small, creating ≤0.5-mm diameter craters. The first two Materials 

International Space Station Experiments (MISSE-1 and MISSE-2) averaged less 

than two impacts/ft2/year (Pippin 2006), with a much smaller shielding factor 

from the then-incomplete ISS structure. As more space debris is added to the 

Figure 6. Postflight photograph displays thin film samples from the Materials International Space Station Experiment-4 

with catastrophic failure of both coatings related to combined thermal cycling and atomic oxygen erosion.
16

environment, however, the impact risk changes. For example, the deliberate 

fragmentation of the Fengyun 1-C spacecraft in 2007 and the collision between the 

Iridium 33 and Cosmos 2251 satellites in 2009 increased the trackable space debris 

by 60 percent. As a result of these two events, the number of untrackable particles 

may have increased by 250,000 or more (Orbital Debris Quarterly News, 2010).

Micrometeoroid or space debris impacts on an experiment may crater the material, 

spall off a coating (Figure 7) or short out a solar cell. It would be difficult to shield 

an experiment from impacts and still have it exposed to the rest of the space 

environment, but the debris environment (at time of publication) is relatively 

benign as it typically affects a very small area of exposure. By the same token, if an 

experiment will be testing various space debris shielding designs, it needs a large, 

unshielded area in space for a long time. Laboratory testing is usually conducted 

with slower velocities (<8 km/s) than those found in orbit, so it may not model 

vaporization, spalling or penetration accurately. Analysis tools such as the Smooth 

Particle Hydrodynamic Code are available and used to predict impact effects. These 

tools have been verified only to the velocity limit of current ground test facilities.

Figure 7. A micrometeoroid or orbital debris impact results in a 0.6-mm diameter crater and approximately 3-mm 

diameter spall (loss of the coating). 2.54 cm diameter Tiodize coating on titanium flown on the Long Duration Exposure 

Facility A0171 experiment.
The orientation of materials experiments are typically in the ram, wake, zenith and/

or nadir directions. Ram refers to the velocity vector of the vehicle and has the 

greatest fluence of AO. Zenith, which points into space in the opposite direction 

of Earth, has the most solar illumination. Wake and nadir are the opposing faces 

of ram and zenith, respectively. The wake direction is good for studying UV 

effects with typically an order of magnitude less AO as the ram direction, and 

some experimenters may wish to fly duplicate samples (ram- and wake-facing) to 

differentiate between AO and UV effects. A nadir orientation is desired for Earthviewing experiments.

Note that the wake direction does not mean “no AO.” During a nominal orbit, the 

ISS currently flies in what is called a local vertical, local horizontal (LVLH) frame, 

with the X-axis in the velocity vector (+XVV; Figure 8, NASA SSP 50699-03, ISS 

Certification Baseline Volume 3: Flight Attitudes). Pitch, yaw and roll are all held 

at zero, so the pressurized mating adapter 2 (PMA-2) always faces exactly forward 

along the velocity vector. During Soyuz and Progress dockings and undockings, the 

ISS may be pitched -90° so PMA-2 faces exactly nadir and, during this orientation, 

wake surfaces can receive AO flux. 

Orientation and 

Location on the ISS

Figure 8. +XVV Z nadir flight attitude.
18

Contamination in both molecular and particulate form can compromise numerous 

types of experiments (materials, optical devices, etc.), as well as affect line-of-sight 

surfaces on the ISS. As mentioned earlier, the ISS has specified in NASA SSP 30426 

what the limits are for molecular deposition, induced molecular column densities 

and the release of particulates. The MAPTIS database for thermal vacuum stability 

(TVS) may assist in choosing materials with low outgassing rates.

If there is no substitute for a higher outgassing material, a thermal vacuum bakeout 

may be performed during assembly to meet ISS outgassing requirements. The 

bakeout should be at a higher temperature than expected in orbit, in a vacuum of 

10-6

 torr or better and for at least 24 hours or until a temperature-controlled quartz 

crystal microbalance (TQCM) in the vacuum chamber indicates no more evolving 

material. Silicone materials that have not been vacuum-baked can frequently 

contain short chain molecules that are volatile and readily transported onto 

neighboring surfaces. When these contaminated spacecraft surfaces are exposed to 

AO in LEO, the silicones oxidize to form silica (or silicates). Hydrocarbons can 

also be trapped on the surface during this process. The resulting deposit can form 

an AO-protective coating that can darken as a result of further solar radiation 

exposure. This contamination can also skew erosion yield measurements, increase 

solar absorptance and irreparably damage optical materials. One such case was 

Approaches to Mitigate 

Contamination

Figure 9. Postflight photograph shows an room temperature vulcanization sample from the Long Duration Exposure Facility 

A0171 experiment with deposited silicone contaminant on baseplate and sample. Also note oxidized silver fastener.
the room temperature vulcanization (RTV) 511 samples of an LDEF experiment 

(Figure 9). The sheet of RTV was baked out for a short time, which was inadequate, 

and then the samples were cut out of the sheet, exposing fresh material. The holes 

drilled for fastening also exposed new material, with outgassed deposits seen in rings 

on top of the sample.

Along with thermal vacuum bakeout and proper materials selection, another 

approach to mitigate contamination is to be aware of line-of-sight to sensitive 

optics and to design vent paths accordingly. If purges or dumps are necessary to 

an experiment, they must not cause any damage to the ISS surfaces from either 

direct or subsequent orbital impingement. NASA SSP 30426 also specifies that the 

limit on particulate generation from an experiment is to be 1 particle (≥100-μm 

diameter) per orbit per 1x10–5 steradian field of view as seen by a 1-m diameter 

aperture telescope. This particulate generation requirement also applies to moving 

parts and vents.
20

Plan for Flight Recovery Contingencies

The biggest lesson learned for passive experiments is that, historically, few materials 

experiments have been returned to Earth after exactly the time planned. The LDEF 

was planned for a 9- to 12-month mission; the Challenger accident left it in orbit 

for 5.8 years. MISSEs-1 and -2 were expected to fly for one year; the Columbia 

accident extended that to nearly four years. Planning for every contingency may 

not be possible, but samples should be chosen to provide the needed science, even if 

return is significantly delayed.

Control Samples and Preflight Testing

The second biggest lesson learned is to perform as much preflight analyses as 

possible on duplicate sets of flight and control samples. It may not be possible 

to have a control sample for every flight sample, but there should be at least one 

control sample retained for each material, orientation and surface treatment, etc. 

The control samples should be maintained appropriately during the experiment 

flight, e.g., bagged in nitrogen or held in a desiccator out of direct sunlight. Many 

polymers, such as Kapton®, are hygroscopic (absorbing up to two percent of their 

weight in moisture) and can fluctuate in mass with humidity and temperature. 

Therefore, for accurate mass loss measurements to be obtained, it is necessary 

that the samples be fully dehydrated immediately before measuring the mass 

both Preflight and postflight. Ideally, this is done by dehydrating the samples 

in a vacuum desiccator maintained at a pressure of ≤13.3 Pa (≤100 mtorr) 

with a mechanical roughing pump for a minimum of 72 hours. Four days are 

recommended. Samples should be removed one at a time and the desiccator 

pumped back down after a sample is removed to keep other samples under 

vacuum. Records should be kept of the following: the sequence of sample weighing, 

the number of samples in each set (if a multilayered sample), the time under 

vacuum before weighing, the temperature and humidity in the room, the time air 

was let into the desiccator and the time a sample was taken out of the desiccator, 

the time of each weighing (with a minimum of three measurements taken as 

quickly as possible) and the mass. Typically, the measurements are averaged, or 

they can be back-extrapolated to time zero (when air is first introduced into the 

desiccator). A balance with at least 0.01‐mg readability is needed. A faster method is 

to place one sample at a time in a vacuum desiccator and pump down to a pressure 

of exactly 6.67 Pa (50 mtorr). When 6.67 Pa is reached, the timer is started at the 

same moment the chamber is vented. The sample is quickly removed from the 

chamber and placed on the balance. Mass measurements are taken every 30 seconds 

Lessons Learned
for 5 minutes; then regression analysis is used to calculate the mass at time zero. 

This method should be used only with a small vacuum chamber that can be vented 

quickly and opened easily. The same procedure and sequence should be repeated 

with the same samples postflight.

In addition to dehydrated mass measurements, some suggested preflight analyses 

include normal and black-light photography, thickness measurements for erosion 

yield calculations, surface roughness, bi-directional reflectance distribution function 

(BRDF), transmission, ellipsometry, solar absorptance, infrared emittance, electrical 

conductivity and solar cell current-voltage curves. These non-destructive tests can 

then be repeated postflight, minimizing any uncertainties related to equipment 

calibration or even use of different equipment after flight. Experiments that are 

not well planned or that occur at the last minute risk yielding poor data, so designs, 

materials selections and analyses should be completed well in advance of experiment 

integration activities.

Understand Sample Geometry

The geometry of sample holders can influence the flux of AO. Commonly used 

passive Evaluation of Oxygen Interaction with Materials (EOIM)/MISSE sample 

trays have chamfered circular apertures that allow AO to scatter from the chamfered 

surfaces onto the samples, thus locally increasing the flux impacting the samples, as 

shown in Figure 10 (Banks et al., 2008). A consequence of the perimeter-scattered 

AO is that the erosion around the sample perimeter is greater than in the central 

area. If the AO is arriving off normal, then there will be a variation in flux around 

the perimeter of each sample, depending on the scattering geometry. As can be 

seen in Figure 11 (de Groh et al., 2008; Banks et al., 2008), two MISSE-2 samples 

peeled up from their lower left edges, and AO was found to be arriving at 8° from 

normal and coming from the upper right 

direction. Thus, there appears to be a flux 

concentration near the perimeter of the 

samples from the AO that was deflected off 

the chamfered surface. 

For the MISSE-2 trays with circular 

2.54-cm-diameter samples, the chamfer 

was at 45o, and the lip was 0.763 mm 

thick. The maximum possible additional 

fluence for 2.54-cm-diameter samples 

Figure 10. Flux concentration from chamfered Evalution 

of Oxygen Interaction with Materials/Materials International Space Station Experiment sample holders.
22

caused by AO scattering would be about15 percent but is more likely around 3.3 

percent. Therefore, the concern is not of a higher average fluence but rather of 

sample peeling and potential release before full sample erosion, which could lead to 

incorrect erosion yields.

The problem of flux concentration and premature peeling could be eliminated if a 

reverse chamfer were used on the sample holders, which would not allow scattering 

of AO onto the surface of the samples. A potential disadvantage of this would be 

the loss of intimate contact at the edge of the sample for profiling purposes, but this 

would not be a concern for mass loss measurements.

Other Lessons Learned

Written instructions and procedures used for mounting and assembly of spaceflight 

hardware are almost always necessary to properly integrate an experiment as 

specified by quality assurance and flight documentation requirements. This step 

Figure 11. The Materials International Space Station Experiment-2 Polymer Erosion and Contamination 

Experiment tray is seen with samples peeling up on their lower-left sides.
occasionally is glossed over by those tasked with final installation. Be clear about 

experiment or sample orientation (“this side faces space”) and properly label in red 

any “REMOVE BEFORE FLIGHT” covers.

Another lesson learned is that an experiment should be able to withstand the 

corrosive salt environment of Cape Canaveral or the Mid-Atlantic Regional 

Spaceport while waiting for launch. The experiment should also be able to 

withstand the power being off, both on the ground and in space. 

Some materials should not be flown in LEO. NASA-STD-6012, Corrosion 

Protection For Space Flight Hardware, section 4.6.2, states that cadmium plating 

and zinc plating shall not be used because of contamination concerns. Silver and 

osmium react strongly with AO and should not be used without a protective 

coating.

Some commercial off-the-shelf electronics may use “lead-free” or pure tin 

components and soldering. This may lead to tin pest (tin plague) or tin whisker 

growth in orbit, potentially adversely affecting the electronics through the release of 

conductive particulate. 

More lessons learned can be found in Banks et al. (2008) and on the NASA Lessons 

Learned database located at http://llis.nasa.gov/llis/search/home.jsp.
24

The current external sites on the ISS for research include:

• Columbus External Payload Facility (EPF) (European Space Agency [ESA])

• EXpedite the PRocessing of Experiments to the Space Station (EXPRESS) 

Logistics Carrier (ELC [NASA])

• Japanese Experiment Module (JEM) External Facility (JEM EF) (Japan 

Aerospace Exploration Agency [JAXA])

• JEM Small Satellite Orbital Deployer (J-SSOD [JAXA])

• Universal workstations with capabilities for re-equipping (URM-Ds), biaxial 

pointing platforms, magnetomechanical anchors/locks and portable working 

platform (Roscosmos).

More information about these facilities can be found in the section on 

ISS External Accomodations.

An experiment setup will depend on what the principal investigator wants to 

study. Issues to consider include: whether the experiment needs to be returned to 

Earth for analysis; if a power source and data downlink are required, i.e., for active 

experiments, such as testing solar cells; can the experiment be robotically deployed 

through the JEM airlock? Spacewalks will be limited in the future, so deployment 

by an astronaut might not be possible. Which environment is better suited for 

the experiment: ram for the most AO; zenith for the most UV radiation; wake for 

UV and little AO; nadir for Earth-observing? How sensitive is the experiment to 

contamination? Table 1 shows the large variation in silica-based contamination on 

surfaces from experiments placed at different locations on the ISS (Dever et al., 

2006; Steagall et al., 2008). The ram-facing surface of MISSE-2 had two orders 

of magnitude less contaminant thickness than the three JAXA experiment units 

attached to the Russian Service Module. This is probably related to differences 

in the total arrival of silicones based on each experiment’s respective view of, and 

distance from, contaminant sources on the ISS. 

Care must be taken to avoid self-contamination of the experiment, as well as to 

ensure that the experiment is out of the view of sources of silicone to be sure that 

AO does not produce silica deposits that can affect erosion yields or cause changes 

in solar absorptance.

Developing and Flying 

Research to the ISS
Location Contaminant thickness (nm) Duration of exposure (yr) Contaminant thickness/

year (nm)

MISSE-2, Tray 1: 

Ram facing

1.3 – 1.4 3.95 0.326 – 0.351

JAXA:

Unit 1, Ram facing

Unit 2, Ram facing

Unit 3, Ram facing

30.0

75.0

93.5

0.863

2.37

3.84

34.8

31.7

24.3

Table 1. Silica-Based Contamination on International Space Station Experiment Surfaces.

The experiment will need to be strong enough to survive the launch loads but not 

overdesigned with excess weight. Vibration and shock tests must be performed on 

the flight hardware, with the test levels dependent on the launch vehicle. Hardware 

must pass sharp-edge inspection to ensure the hardware poses no snagging danger 

to a gloved and/or tethered astronaut on an Extravehicular Activity (EVA). Active 

experiments will have to meet electromagnetic interference/electromagnetic 

compatibility (EMI/EMC) and grounding requirements so as not to interfere with 

the avionics and communications on the ISS. In addition, experiments will need to 

undergo a thermal-vacuum bake-out prior to flight.

Experiments should have adequate internal data storage capacity, as no data storage 

capacity is available from the ISS. Some compression within the experiment may be 

needed for transmission of high volumes of data. 

Pictures taken at various time intervals throughout the duration of the experiment 

can also provide valuable information as to experiment function and material 

degradation. For example, images taken of MISSE-1 and MISSE-2 experiments 

showed deterioration of certain 

samples early in the mission and years 

before the experiment was retrieved. 

In Figure 12, near the center, the 

Dyneema tether has failed in tension, 

as has the dielectric foam sample 

above and to the right. To the right 

of the tethers experiment, one sample 

of Kapton® has already eroded away, 

while the adjacent sample of Kapton® 

with aluminum acetylacetonate is still 

intact. Along the right edge, part of 

a beta cloth experiment has already 

turned brown. This was with only a 

little more than 6 months’ exposure, 

and MISSE-2 would stay in orbit for 

nearly four years.

Figure 12. Image shows sample deterioration on Materials 

International Space Station Experiment-2 after six months 

in orbit. 
26

This section gives an overview of existing ISS External Facilities (EF) that are 

multidisciplinary in nature (Figure 13), provide access to multiple sites that are exposed 

to the space environment and include structural attachment points and utility interfaces.

Japanese Experiment Module Exposed Facility 

(Japan Aerospace Exploration Agency [JAXA]) 

The JEM Exposed Facility (EF) is an unpressurized, multipurpose pallet structure 

attached to the JEM, or “Kibo,” meaning “hope” in Japanese (Figure 14). This 

external platform is used for research in the areas of communications, space science, 

engineering, technology demonstration, materials processing and Earth observation. 

Experiments can be monitored by six video cameras, two located on the JEM EF, 

two located externally on the JEM Pressurized Module and two on the JEM Remote 

Manipulator System (RMS).

The JEM EF is accessible from the internal pressurized volume of the ISS via the 

JEM Airlock. Articles interfacing with this structure are grappled and moved using 

the JEM Remote Manipulator System (RMS). The JEM EF is roughly 5.7 m x 5.0 

m x 3.8 m, weighs approximately 4,000 kg, and includes utilities at each of the 12 

attachment sites (Figure 13). The JEM EF specification and available resources are 

listed below:

ISS External 

Accommodations

Figure 13. Kibo Exposed Facility. The yellow and black numbers show the payload attach points.
• Mass: 500 kg (10 standard sites, mass includes payload interface unit [PIU]); 

2,500 kg (two heavy sites, mass includes PIU)

• Size: 1.85 m x 1 m x 0.8 m

• Power: 3 kW or 6kW for each payload, and 11kW max in total, 120 VDC

• Thermal: Active thermal control (fluid loop)

• Data: Low Rate – 1 mbps MIL-STD-1553; High Rate – 43 mbps (shared); 

Ethernet – 100Base-TX. Downlink and uplink are available through NASA 

communication system or the Japanese Inter-orbit Communication System(ICS).

Multi-Purpose Experiment Platform (MPEP [JAXA])

Multi-Purpose Experiment Platform (MPEP) (Figure 15) is available for the 

medium-sized space-exposed experiment, which aims for the monitor of space 

environment for Earth observation. MPEP accommodates up to a 100 kg 

experiment, and it provides the appropriate attitude and position for payloads by 

the support of JEM RMS. Available resources for payloads on MPEP are power, 

communication and video interface.

One of the applications includes the JEM Small Satellite Orbital Deployer 

(J-SSOD) in the next section.

Figure 14. Multi-Purpose Experiment Platform installed on the Japanese Experiment Module Slide Table.
28

JEM Small Satellite Orbital Deployer (J-SSOD [JAXA])

The J-SSOD (Figure 15) holds up to three small one-unit (1U; 10 cm x 10 cm x 10 

cm) CubeSats per satellite-install case (six total), though other sizes (up to 55 cm x 

55 cm x 35 cm) can be used from 2015 (Figure 16). Weight limit is 1.33 kg per 1U, 

8 kg total. Crew members unpack and install the satellite-install case onto the MultiPurpose Experiment Platform (MPEP), which is on the JEM Slide Table. The small 

satellites are then transferred out through the JEM Airlock and deployed by the JEM 

RMS. Nine satellites were successfully deployed in two missions since 2012. 

EXPRESS Logistics Carrier (NASA)

The ELC is a pallet designed to support external research hardware and store 

external spares (Orbital Replacement Units [ORUs]) needed over the life of the ISS. 

Currently, four ELCs are mounted to ISS trusses, providing unique vantage points 

for space, technology and Earth observation investigations. Two ELCs are attached 

to the integrated truss segments starboard truss 3 (ITS‐S3) and two ELCs to the 

integrated truss segments port truss 3 (ITS-P3). By attaching at the S3/P3 sites, 

a variety of views such as zenith (deep space) or nadir (Earthward) direction with 

a combination of ram (forward) or wake (aft) pointing allows for many possible 

viewing opportunities. The ELC Adapter capabilities are listed below:

• Mass: 227 kg (8 sites across four ELCs; not including adaptor plate)

• Volume: 1.2 m3

• Size: 0.8 m x 1.2 m x 1.2 m

• Power: 750 W, 113-126 VDC; 500 W, 28 VDC

• Data: Low Rate: 1 Mbps MIL-STD-1553 Medium Rate: 6 Mbps (shared)

Figure 15. The Japanese Experiment Module Small 

Satellite Orbital Deployer (J-SSOD).

Figure 16. CubeSats deployed from the J-SSOD and 

photographed against an ISS solar array background.
MISSE-7, consisting of PECs, called MISSE-7A and MISSE-7B and later MISSE-8, 

were located on ELC 2, Site 3, as shown in Figure 17. MISSE-7A was flown in 

a zenith/ram orientation, and MISSE-7B was flown in a ram/wake orientation. 

MISSE-8, along with a small passive sample tray called the Optical Reflector 

Materials Experiment III (ORMatE III), were deployed in the same location as 

MISSE-7A and -7B, respectively, (Figure 18). MISSEs 1 through 5, 6A and 6B were 

located in various other (non-permanent) external locations on the ISS.

Figure 17 a. Materials International Space Station Experiment-7 on EXPRESS Logistics Carrier 2, Site 3 shortly after 

deploy during STS-129 in November 2009.

Figure 17 b. Photo of the zenith facing surface of Materials International Space Station Experiment (MISSE) 7A (left) and 

the ram facing surface of MISSE-7B (right) during STS-130 in February 2010.
30

Figure 18 b. Materials International Space Station Experiment-8 (left) and ORMatE III (lower center) on EXPRESS Logistics 

Carrier 2, Site 3. Images in both 18a & 18b were taken in July 2011 during STS-135.

Figure 18 a. Closeup of the zenith-facing surface of Materials International Space Station Experiment-8.
Figure 19. European Space Agency Columbus External Payload Facility.

External

Payload

Facility

Columbus External Payload Facility (European Space Agency)

The EPF (Figure 19) has four powered external attachment site locations for scientific 

payloads or facilities. To date, the EPF has been used by ESA and NASA. Each of the 

four attachment sites may hold a mass of up to 290 kg, and each is provided utility 

connections for power and data. The EPF capabilities are listed below:

• Mass: ≤ 290 kg per site (4 sites; uses Columbus External Payload Adapter 

[CEPA])

• Volume: 1.39 m3

• Size: 0.864 m x 1.168 m x 1.375 m

• Power: 1,250 W (total limited to 2,500 W for all four external payloads), each 

external facility is connected to 2 – 120 VDC power feeds. Note: equipment 

can only be radiatively passively cooled.

• Data: Low Rate: up to 100 Mbps
32

Russian Segment External Facilities (Russian Federal 

Space Agency)

Russian Segment External Facilities [Roscosmos] consists of a whole host of 

multi-user external Facility interface locations and support structures. There are 

two external mounting platforms (URM-D) on the Service Modulus (SM), also 

known as Zvezda, located on the starboard and port sides. The URM-D provides 

power and data connection to the payloads. Passive experiments can be flown in 

the Replaceable Cassette-Container (SKK or CKK) and can be attached to Zvezda, 

the Poisk MRM, or the Pirs Docking Compartment-1 (Figure 20). Payloads are 

installed and removed by the crew during an EVA.

Figure 20. Russian Replaceable Cassette-Container on the Russian Service Module.
NASA and the National Science Foundation fund a variety of research efforts 

through grants, fellowships, contracts, interagency transfers and cooperative 

agreements. The NASA Solicitation and Proposal Integrated Review and Evaluation 

System (NSPIRES; http://nspires.nasaprs.com/) lists the solicitations and research 

opportunities. A guidebook for responding to NASA Research Announcements 

(http://www.hq.nasa.gov/office/procurement/nraguidebook/) guides the reader 

through notice of intent to propose, preparation of proposals, conduct of the 

research itself and the sharing of new knowledge through publications, public 

outreach and education. Research proposals are peer reviewed and should meet 

national research objectives.

NASA receives thousands of research proposals every year and may only fund a 

limited number. However, in accordance with Federal statutes and NASA policy, no 

eligible applicant is to be excluded from participation or be denied the benefits of 

or be subjected to discrimination under any program or activity receiving financial 

assistance from NASA on the grounds of race, color, creed, age, sex, national origin 

or disability. Proposals from all qualified sources are welcome. Historically Black 

Colleges and Universities (HBCUs), Other Minority Universities (OMUs), small 

disadvantaged businesses (SDBs), veteran-owned small businesses (VOSBs), servicedisabled veteran-owned small businesses (SDVOSBs), Historically Underutilized 

Business Zone (HUBZone) small businesses and women-owned small businesses 

(WOSBs) are encouraged to submit proposals. 

Funding Opportunities
34

ASTM E595, Standard Test Method for Total Mass Loss and Collected Volatile 

Condensable Materials from Outgassing in a Vacuum Environment.

ASTM E1559, Standard Test Method for Contamination Outgassing 

Characteristics of Spacecraft Materials.

ASTM E2089, Standard Practices for Ground Laboratory Atomic Oxygen 

Interaction Evaluation of Materials for Space Applications.

Banks, B. A., K. K. de Groh, and S. K. Miller. “Low Earth Orbital Atomic 

Oxygen Interactions with Spacecraft Materials.” 2004 Materials Research Society 

Fall Meeting, Boston, MA, November 29-December 3, 2004, Materials Research 

Society Symposium Proceedings 2004, NN8.1. New York: Cambridge University 

Press, 2004. Also published as NASA/TM-2004-213400 (Nov. 2004). 

Banks, B. A., K. K. de Groh, S. K. Miller, and D. L., Waters. “Lessons Learned 

from Atomic Oxygen Interaction with Spacecraft Materials in Low Earth Orbit.” 

Proceedings of the 9th International Conference: Protection of Materials and 

Structures from Space Environment, Toronto, Canada, 20–23 May 2008. Melville, 

NY: American Institute of Physics. Also published as NASA/TM-2008-0215264, 

July 2008. 

de Groh, K. K., B. A. Banks, C. E. McCarthy, R. N. Rucker,, L. M. Roberts, and 

L. A. Berger. “MISSE-2 PEACE Polymers Atomic Oxygen Erosion Experiment on 

the International Space Station.” High Performance Polymers 20 (2008): 388-409. 

de Groh, K. K., D. A. Jaworske, W. H. Kinard, H. G. Pippin, and P. P. Jenkins. 

“Tough Enough for Space: Testing Spacecraft Materials on the ISS.” Technology 

Innovation 15, no. 4 (2011): 50-53.

de Groh, K. K., B. A. Banks, J. A. Dever, D. J. Jaworske, S. K. Miller, E. A Sechkar, 

and S. R. Panko. “NASA Glenn Research Center’s Materials International Space 

Station Experiments (MISSE 1-7).” Proceedings of the International Symposium on 

“SM/MPAC&SEED Experiment,” Tsukuba, Japan, March 10-11, 2008. JAXA-SP08-015E, March 2009, pp. 91 – 119. Also published as NASA TM-2008-215482, 

December 2008.

Dever, J. A., B. A. Banks, K. K. de Groh, and S. K. Miller. “Degradation of 

Spacecraft Materials.” In Handbook of Environmental Degradation of Materials, 

edited by Myer Kutz, 465-501. Norwich, NY: William Andrew Publishing, 2005. 

Citations
Dever, J. A., S. K. Miller, E. A. Sechkar, and T. N. Wittberg. “Preliminary 

Analysis of Polymer Film Thermal Control and Gossamer Materials Experiments 

on MISSE-1 and MISSE-2,” Proceedings of the 2006 National Space & Missile 

Materials Symposium in conjunction with the 2006 MISSE Post-Retrieval 

Conference, June 26-30, 2006, Orlando, FL. 

Hathaway, David H. “Solar Cycle Prediction.” http://solarscience.msfc.nasa.gov/

predict.shtml (accessed April 30, 2013).

NASA SSP 30426, Space Station External Contamination Control Requirements.

NASA SSP 50699-03. ISS Certification Baseline Volume 3: Flight Attitudes

James, Bonnie F., O. W. Norton, and Margaret B. Alexander. “The natural space 

environment: Effects on spacecraft.” NASA STI/Recon Technical Report N 95 

(1994): 25875.

Pippin, H. G., E. Normand, S. L. B. Woll, and R. Kamenetzky. “Analysis of 

Metallized TeflonTM Thin-Film Materials Performance on Satellites.” Journal of 

Spacecraft and Rockets 41, no. 3 (May 2004): 322-325.

Pippin, Gary. Summary Status of MISSE-1 and MISSE-2 Experiments and Details 

of Estimated Environmental Exposures for MISSE-1 and MISSE-2, AFRL-MLWP-TR-2006-4237, Air Force Research Laboratory: OH July 2006.

Schwank, James R., Marty R. Shaneyfelt, and Paul E. Dodd. “Radiation Hardness 

Assurance Testing of Microelectronic Devices and Integrated Circuits: Radiation 

Environments, Physical Mechanisms, and Foundations for Hardness Assurance.” 

Sandia National Laboratories Document SAND-2008-6851P. 21 May 2013. http://

www.sandia.gov/mstc/services/documents/Sandia_RHA_Foundations_FINAL.pdf.

Silverman, E. M. Space Environmental Effects on Spacecraft: LEO Materials 

Selection Guide. NASA CR 4661, Part 1, TRW Space & Electronics Group: 

Redondo Beach, CA. August 1995. 

Soares, C. and R. Mikatarian. “Understanding and Control of External 

Contamination on the International Space Station.” ESA SP-540 Proceedings of the 

9th International Symposium on Materials in a Space Environment, June 16-20. 

2003, Noordwijk, The Netherlands. Noordwijk: ESA. September 2003. 
36

Stambler, A. H., K. E. Inoshita, L. M. Roberts, C. E. Barbagallo, K. K. de Groh, 

and B. A. Banks. “Ground-Laboratory to In-Space Atomic Oxygen Correlation for 

the PEACE Polymers.” Proceedings of the 9th International Conference Protection 

of Materials and Structures from Space Environment, May 19-23, 2008, Toronto, 

Canada. Ed. J .I. Kleiman, AIP Conference Proceedings 1087, pp. 51-66, 2009. 

Also NASA TM-2011-216904..

Steagall, C., K. Smith, A. Huang, C. Soares, and R. Mikatarian, “Induced 

Contamination Predictions for JAXA’s Micro-Particles Capturer and Space 

Environment Exposure Devices,” Proc. Int. Symp. on SM/MPAC&SEED 

Experiment, Japan, 2008 (JAXA-SP-08-015E, 2009), pp. 19-25. 

Townsend, J. A, P. A, Hansen, M. W. McClendon, K. K. de Groh, and B. A, Banks. 

“Ground-Based Testing of Replacement Thermal Control Materials for the Hubble 

Space Telescope” High Performance Polymers 11 (1999): 63-79.

“Update on Three Major Debris Clouds.” Orbital Debris Quarterly News 14. 2 

(April 2010). 14 May 2013 <http://orbitaldebris.jsc.nasa.gov/newsletter/pdfs/

ODQNv14i2.pdf>.

Visentine, J., W. Kinard, R. Pinkerton, D. Brinker, D. Scheiman, B. Banks, J. 

Zwiener, K. Albyn, T. Farrell, S. Hornung, and T. See. “Mir Solar Array Return 

Experiment”, AIAA Paper 99-0100 presented at the 37th AIAA Aerospace Sciences 

Meeting, Reno, NV, Jan. 11-14, 1999.

Yang, J. C. and K. K. de Groh. “Materials Issues in the Space Environment.” MRS 

Bulletin 35 (January 2010): 12-19.
AO Atomic Oxygen

BRDF Bi-Directional Reflectance Distribution Function

CEPA Columbus External Payload Adapter

CKK Replaceable Cassette-Container

CTE Coefficients of Thermal Expansion

EF External Facility

ELC EXPRESS Logistics Carrier

EMI/EMC Electromagnetic Interference/Electromagnetic Compatibility

EOIM Evaluation of Oxygen Interaction with Materials

EPF External Payload Facility

ESA European Space Agency

EVA Extravehicular Activity

EXPRESS EXpedite the PRocessing of Experiments to the Space Station

HBCUs Historically Black Colleges and Universities

HUBZone Historically Underutilized Business Zone

ISS International Space Station

ITS-S3 Integrated Truss Segments Starboard Truss 3

ITS-P3 Integrated Truss Segments Port Truss 3

JAXA Japan Aerospace Exploration Agency

JEM Japanese Experiment Module

J-SSOD JEM Small Satellite Orbital Deployer

LDEF Long Duration Exposure Facility

LEO Low-Earth Orbit

LVLH Local Vertical, Local Horizontal

MAPTIS Materials and Processes Technical Information System

MISSE Materials International Space Station Experiment

MMOD Micrometeoroids and Orbital Debris

MPEP Multi-purpose Experiment Platform

MRM2 Mini Research Module 2

MSIS Mass Spectrometer Incoherent Scatter

NSPIRES NASA Solicitation and Proposal Integrated Review and Evaluation System

OMUs Other Minority Universities

PEACE Polymer Erosion and Contamination Experiment

PEC Passive Experiment Container

PIU Payload Interference Unit

PMA-2 Pressurized Mating Adapter 2

RMS Remote Manipulator System

Roscosmos Russian Federal Space Agency

RTV Room Temperature Vulcanization

SDBs Small Disadvantaged Businesses

SDVOSBs Service-Disabled Verteran-Owned Small Businesses

SKK Replaceable Cassette-Container

TQCM Temperature-Controlled Quartz Crystal Microbalance

TVS Thermal Vacuum Stability

URM-Ds Universal Workstations with Capabilities for Re-Equipping

UV Ultraviolet

VCMO Vacuum Condensable Material Compatibility with Optics

WOSBs Women-Owned Small Businesses

+XVV X-axis in the Velocity Vector

Acronyms
38

The Complete ISS Researcher’s 

Guide Series

1. Acceleration Environment

2. Cellular Biology

3. Combustion Science

4. Earth Observations

5. Fluid Physics

6. Fruit Fly Research

7. Fundamental Physics

8. Human Research

9. Macromolecular Crystal Growth

10. Microbial Research

11. Microgravity Materials Research

12. Plant Science

13. Rodent Research

14. Space Environmental Effects

15. Technology Demonstration

38
For more information...

Space Station Science

http://www.nasa.gov/iss-science

Facilities

http://www.nasa.gov/mission_pages/station/research/

facilities_category.html

ISS Interactive Reference Guide:

http://www.nasa.gov/externalflash/ISSRG/index.htm

Researchers/Opportunities

http://www.nasa.gov/mission_pages/station/research/

ops/research_information.html

ISS Research Customer Helpline

JSC-ISS-research-helpline@mail.nasa.gov

281-244-6187
40

National Aeronautics and Space Administration

Johnson Space Center

http://www.nasa.gov/centers/johnson

www.nasa.gov

NP-2015-03-015-JSC