JOURNAL OF SPACECRAFT AND ROCKETS
Vol. 51, No. 3, May–June 2014
Review of Evaluation Methodologies for Satellite Exterior
Materials in Low Earth Orbit
Devarakonda Angirasa∗
European Aeronautic Defence and Space Company N.V., Stevenage,
England SG1 2AS, United Kingdom
and
Portonovo S. Ayyaswamy†
University of Pennsylvania, Philadelphia, Pennsylvania 19104-6315
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DOI: 10.2514/1.A32742
The objective of this review paper is to present a formal technical evaluation methodology for the materials used on
a satellite in low Earth orbit. It attempts to succinctly describe the current materials evaluation, selection, and design
processes from the perspective of space systems engineering. The low Earth orbit has a severe environment that can
degrade the exterior materials of a satellite, thereby adversely affecting its ability to function. The environmental
factors that cause the degradation are first described and quantified. The exterior elements/equipment/units of a
satellite are identified. Materials exposed to the environment and their thermal, chemical, and electromagnetic
interactions with the various environmental factors, individually and in combination, are analyzed. A methodology
that combines the evaluations of individual environmental factors is proposed. Various materials are evaluated, and
where possible, recommendations are made for the suitability of their use on the exterior of low-Earth-orbit satellites.
Nomenclature
A
A
Ef
=
=
=
Et
Eη
00
FAO
00
_
FAO
G
h
I
Keff
=
=
=
=
=
=
=
=
L
l
=
=
surface area, cm2 , m2
cross-sectional area, m2
atomic oxygen erosion yield factor of a material
(nondimensional)
atomic oxygen erosion thickness of a material, cm
atomic oxygen erosion yield of a material, cm3 ∕O atom
atomic oxygen fluence, number of O atoms∕cm2
atomic oxygen flux, number of O atoms∕cm2 · s
incident radiation heat flux on a surface, W∕m2
distance, cm
infrared radiation flux from Earth, W∕m2
effective thermal conductivity of insulation
blanket, W∕m · K
length, m
thickness of insulation blanket, m
m
Q
Q_ 0 0
q
R
RS
TC
TH
TS
t
t
tE
=
=
=
=
=
=
=
=
=
=
=
=
V
W
α
ε
=
=
=
=
mass, g
heat transfer, W
conduction heat flux, W∕m2
heat flux, W∕m2
electrical resistance, Ω
surface resistivity or sheet resistance, Ω∕□
cold-side temperature of insulation blanket, K
hot-side temperature of insulation blanket, K
equilibrium temperature of a surface, K
mission duration time, s
layer thickness, m, cm
atomic oxygen erosion thickness number of
a material, μm
satellite relative velocity (∼7.6 km∕s)
width, m
surface absorptivity of solar radiation
surface emissivity of infrared radiation
Devarakonda Angirasa received his Ph.D. in mechanical sciences from the Indian Institute of Science, Bangalore in
1989. Afterward, he worked in the United States, the United Kingdom, Holland, and Germany in academia, industry,
and at NASA in spacecraft systems engineering and design; mechanical, thermal, and fluids engineering;
development and manufacturing of two-phase thermocapillary devices; high-temperature materials and
applications; environmental testing; and spacecraft materials technologies. His scientific interests include planetary
atmospheres and space environment.
Portonovo S. Ayyaswamy received his Ph.D. in engineering from the University of California, Los Angeles in 1971.
Since 1974, he has been on the faculty at the School of Engineering and Applied Science, University of Pennsylvania,
Philadelphia. He is at present the Asa Whitney Professor of Dynamical Engineering at the University of Pennsylvania.
His research is in the area of mechanical engineering, with focuses in modeling, simulations, and experimentation of
multiphase flow/heat and mass transfer. Over the years, he has contributed to many diverse areas of heat transfer,
mass transfer, and fluid mechanics.
Received 21 June 2013; revision received 22 October 2013; accepted for publication 2 December 2013; published online 21 April 2014. Copyright © 2013 by the
American Institute of Aeronautics and Astronautics, Inc. All rights reserved. Copies of this paper may be made for personal or internal use, on condition that the
copier pay the $10.00 per-copy fee to the Copyright Clearance Center, Inc., 222 Rosewood Drive, Danvers, MA 01923; include the code 1533-6794/14 and $10.00 in
correspondence with the CCC.
*Principal Engineer, Astrium Satellites, Gunnels Wood Road.
†
Asa Whitney Professor of Dynamical Engineering, Department of Mechanical Engineering and Applied Mechanics, 231 Towne Building (Corresponding
Author).
750
751
ANGIRASA AND AYYASWAMY
ε
θ
ρ
ρ
ρAO
ρS
σ
□
↑
↓
=
=
=
=
=
=
=
=
=
=
effective emittance or E-star of insulation blanket
angle
electrical resistivity, Ω · m, Ω · cm
density of a material, g∕cm3
atomic oxygen density, number of O atoms∕cm3
solar reflectivity of a surface
5.67 × 10−8 W∕m2 · K4 , Stephan–Boltzmann constant
square area
increase
decrease
I.
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T
Introduction
HE orbits of Earth for satellites may be classified on the basis of
their altitude. The orbit nearest to Earth is called low Earth orbit
(LEO) in the range 160–1500 km altitude. An example is the Hubble
Space Telescope, which orbits at an LEO of about 600 km.
Geosynchronous orbit (GEO) is at an altitude of 35,786 km above
equator. Most communications and weather satellites orbit in GEO
because in this orbit the satellite appears stationary above a fixed
location on Earth, approximately equaling the Earth’s rotational
speed and duration of rotation ∼24 hours. Orbits between LEO and
GEO may be called medium Earth orbits. Satellites orbiting Earth
above GEO are called supersynchronous, and those spacecraft that
leave the proximity of Earth are interplanetary. Spacecraft that leave
the solar system altogether are called interstellar. Wertz [1] provides a
comprehensive guide on spacecraft orbit and attitude systems, their
design, and management, and the Wertz et al. compendium [2] is a
comprehensive guide on all aspects of the analysis and design of the
space mission.
LEO satellites usually use polar or highly inclined orbits. A polar
orbit is one that is perpendicular to the equator (90 deg), and a highly
inclined orbit is with reference to the equator. An Earth polar orbiting
satellite can pass over its entire surface as the Earth rotates beneath the
satellite orbital path. The LEO orbits are particularly useful for
scientific missions of Earth observation. For example, measurements
of various atmospheric parameters (natural and anthropogenic
aerosols, their radiative properties, interactions with clouds, to name
a few) over the entire planet can be done by satellites with appropriate
science payload in these orbits. A cluster of such satellites providing
measured data over several years can provide an accurate quantitative
picture of the changing climate of the Earth in extensive detail.
Design of LEO (and polar) satellites have particular challenges.
These can be grouped under 1) orbit-related, and 2) LEO
environmental. The orbit-related challenges arise from the requirement that more fuel needs to be carried on the satellite and more
power generated. Hence, the power and propulsion modules are
different for LEO satellites that carry more mass, and this in general
affects the satellite design. The other factor is that LEO has an
environment rich in atomic oxygen (AO), plasma, and ultraviolet
(UV) radiation. All of these, individually and in combinations,
adversely affect the materials used to build the satellite, especially the
exterior. Anderson and Smith [3] give a detailed description of the
LEO environment covering plasma, charged particles, electromagnetic radiation, meteoroids and orbital debris, magnetic field, and
thermal environment.
Atomic oxygen environment is quantified indirectly by measuring
the effects it produces on various material coupons. The report by
Visentine and Whitaker [4] and its revision by Dooling and Finckenor
[5] describe these effects. A more detailed and extensive report is that
by Silverman [6], which surveys the effects of all LEO environmental
factors on spacecraft materials.
In evaluating the environmental effects on satellite materials, the
duration of the mission is an important consideration. Material
suitable for a short-duration mission may fail when used for a long
duration. The data that are available on the LEO environmental
effects on materials are derived from short-term and long-term flight
experiments. Short-term space flights (∼40 h) include Space Shuttle
flights, and long-duration experimental data are from retrieved
satellites, particularly Long Duration Experimental Facility (LDEF)
and Solar Maximum Mission. There are several ground-based
simulators for experimental studies of a material’s exposure to a given
space environmental parameter; for example, AO or combined AO
and UV radiation. However, there is no test facility on Earth that can
simulate the complete orbital environment.
Cost is understandably the deciding factor in the choice of materials.
The huge expense involved to design, develop, manufacture, test, and
qualify the materials for space use is one of the parameters taken into
consideration when performing tradeoffs on material choices. Some
materials thus developed at large expense for space use might soon go
out of production because of a lack of demand. Availability of a given
material is another factor that must be considered.
The aims of this paper are to combine the best practices from space
agencies and satellite industry and to succinctly provide an evaluation
methodology for appropriate choice of exterior materials in the
design of LEO satellites. The paper does not, however, survey all the
available materials and their interactions with the LEO environment.
It will not also elucidate the physical science of the LEO space
environment in detail but only discuss enough to enable a sound
engineering judgment on the choice of materials. For a brief and lucid
introduction to the physical science of space environment, see, for
example, Tascione [7]. The current article addresses the needs of the
space system engineers who need to make decisions on materials
without having all the interdisciplinary scientific knowledge required
for such a task.
The review is organized as follows. The environmental factors to
be considered for the material evaluation are first described. This is
followed by a brief description of the exterior of a satellite from the
point of subsystems/equipment/elements exposed to the space
environment. This serves to identify the type of materials and their
function on the satellite as well as the properties required to perform
those functions. Next, the effects of the environmental factors on
these properties, individually or in combination, is discussed through
a quantified evaluation process. Finally, some general recommendations are offered. It needs to be emphasized, however, that it is not a
goal of this review to make recommendations on further research on
materials and their interactions with the LEO environment.
II.
Low Earth Orbit Space Environment
Smith and West [8] provide data on the physical environment
of Earth’s terrestrial space. This region of space is defined as
extending approximately from the lower boundary of the ionosphere
(∼80 km altitude) to the boundary of the magnetosphere where
the unperturbed interplanetary medium is reached. A spacecraft
operating within this space is affected by the environmental factors:
Earth’s gravity, neutral atmosphere, solar radiation and albedo,
electric and magnetic fields, thermal plasma, fast charged particles,
meteoroids, and space debris. Anderson and Smith [3] contains
NASA’s guidelines for the development of spacecraft that operate in
near-Earth space environment, with emphasis on the low Earth orbit
of about 200 to 1000 km altitude. The European Space Agency (ESA)
guidelines are contained in [9], while [10] is a comprehensive source
of information on the near-Earth space environment.
In interplanetary space, there is no AO, but a large amount of AO is
present in the altitude range 200–700 km above Earth. UV radiation is
of two types based on the wavelength: solar and vacuum. Vacuum
ultraviolet (VUV) radiation is more damaging, but it is not present at
this altitude because it is absorbed by the atmospheric gases. Solar
ultraviolet (SUV) radiation is present and will be considered in
conjunction with AO for their combined effect on the spacecraft
exterior material erosion and degradation of the properties.
The low Earth orbit atmosphere comprises ∼90% atomic oxygen.
AO is an extremely reactive chemical species, which can severely
erode exterior materials. In the case of hydrocarbons, the AO reaction
resultant products such as CO, CO2 , H2 O, and others are volatile and
evaporate from the surface, thereby exposing the interior materials
for further reaction with AO. Material evaluation has to take into
consideration the particular space environment of the satellite so that
appropriate materials may be chosen and any additional mitigating
measures taken.
752
ANGIRASA AND AYYASWAMY
In high-inclination LEO, because of the presence of plasmas in
some of the regions, electrical charge may build up on the exterior
surfaces of the satellite. In that case, electrical grounding has to be
considered. Small amounts of charged particles are present in
the polar LEOs, but their effect is not significant. Particulate
contamination is a concern that has to be addressed in the context of
the manufacturing processes.
Micrometeoroid and space-debris effects are not evaluated in this
review.
Downloaded by GA Inst of Technology on May 9, 2023 | http://arc.aiaa.org | DOI: 10.2514/1.A32742
III.
Space-Exposed Materials of a Satellite
in Low Earth Orbit
A typical satellite has a base platform that structurally supports the
hardware of various subsystems, some parts mounted internally and
some externally and science payload instruments (mostly external).
Figure 1 is an artist’s impression of a LEO satellite, which shows the
base platform, some instruments, and the solar panel array that forms
an external part of the power subsystem.
As seen in Fig. 1 and in any photograph of a satellite, its surface
generally looks golden in color. The golden color emanates from
the outer layer of the multilayer insulation (MLI) blankets. The
spacecraft exterior is mostly blanketed with MLI with cutouts for
radiator surfaces. The spacecraft external surface is mostly composed
of either insulation blanket or radiator surface materials. Both of these
are classified under the thermal control subsystem. The design of
insulation blankets and radiator surface finishes is a vital aspect of not
just the satellite thermal control but of the entire satellite system. The
materials for them must be chosen such that they withstand the lowEarth-orbit space environment and perform their thermal function
while meeting the requirements of spacecraft integration.
In this review, special focus will be on the materials for the spaceexposed outer layer of the insulation blankets, solar cell cover slips,
and the radiator surface finish. In addition, many spacecraft exterior
materials used for other subsystems and instruments will also be
covered.
The external thermal control materials are mainly either insulation
blankets or radiator surfaces. Insulation blankets are used on the
spacecraft main structure and are composed of multiple layers of thin
flexible sheets. The radiator surface can be a second surface mirror
(SSM), which is a more rigid blanket or tile attached to the spacecraft
structure. It can also be an optical solar reflector (OSR) or a paint. The
materials used in insulation blankets are different from those in the
blankets or tiles for the radiators.
An SSM is a thin sheet of transparent material (e.g., glass or
quartz), silvered or aluminized on one side and bonded to a rigid
structure (Fig. 2). It is both an emitter and a reflector. Materials used
for SSMs are transparent to solar radiation, and hence they are
Fig. 1
Artistic rendition of a satellite in LEO (courtesy ESA).
metalized on one side so that the radiation reaching the metallic
coating (silver or aluminum) is reflected back to space. The coated
rear surface thus is called a second surface. The main characteristics
of SSM materials are excellent emission of infrared radiation, so that
the waste heat generated within the spacecraft can be conducted to
them and then radiated to space, and high resistance to plasma and
SUV. SSMs are widely used as radiators of the spacecraft and cover
slips for solar cells. OSR is a material that reflects the sun’s visible
spectrum of radiation, and it is not synonymous with SSM, although
they may be functionally similar. White paints, for example, are
OSRs but not SSMs. Quartz and glass SSMs are sometimes
confusingly referred to as OSRs, while Teflon based reflectors as
SSMs. In this review, SSM is as described through Fig. 2, irrespective
of the material used. We avoid using the term OSR and instead use the
material name for the sake of clarity.
White paints are composed of a white pigment in a clear binder.
They have low solar absorptivity and high infrared emissivity. They
also have high reflectivity, which is diffuse, while SSM’s reflection is
specular. Ceramic coatings have similar thermo-optical properties as
white paints and have the advantage that thinner and uniform coatings
can be applied on a substrate. Black paints have nearly black body
properties of high absorptivity and high emissivity near unity. They
are usually preferred inside the spacecraft, and therefore are not
considered in this review. Black Kapton may be used as high-gain
antenna shield.
Metal side-exposed blankets are necessary in a small number of
applications (e.g., propellant lines). Metallic shields are used to block
the radiative view from hot thrusters to other areas of spacecraft.
Some components that are conductively linked to the spacecraft may
also sometimes have a metallic finish. Metallic surfaces are avoided
for most of the exterior of the spacecraft because they have very high
value of the ratio α∕ε and may attain high temperatures.
IV.
Evaluation Methodology for Material–
Atomic-Oxygen Reactions
Atomic oxygen is formed in the upper atmosphere by UV
photolysis and dissociation of molecular oxygen. The amount of AO
present in the upper atmosphere is given by the term AO density ρAO,
which is defined as the number of atoms of oxygen present in a unit
volume of one cubic centimeter. The AO density is dependent on the
solar activity (low, medium, high) and varies inversely with altitude.
The VUV component of the solar irradiance causes the photolysis.
AO density is calculated based on a mass spectrometer incoherent
scatter model [11]. NASA developed a web software based on this
model that is available for such calculations. The model takes in the
input of year, day of the year, universal time, altitude, geodetic
latitude and longitude, and local apparent solar time. It provides
output of the number density as defined previously for He, O, N2 , O2 ,
Ar, H, and N; total mass density; and neutral temperature and
exospheric temperature. Values of AO density for the satellite orbits
may vary ∼O107 –1011 atoms∕cm3.
AO density thus obtained for a given mission is used to define an
AO flux as the product of the AO density and the mean velocity V of
the satellite relative to Earth’s rotation given by
Fig. 2
Second surface mirror.
753
ANGIRASA AND AYYASWAMY
atoms
atoms
cm
00
F_ AO
·
V
ρ
AO
s
cm2 · s
cm3
(1)
Downloaded by GA Inst of Technology on May 9, 2023 | http://arc.aiaa.org | DOI: 10.2514/1.A32742
The average satellite relative velocity is 7.6 km∕s. The flux is the
number of oxygen atoms that would strike the satellite per second
per square centimeter area of its external surface in the ram
direction (θ 0).
From this mean value of AO flux for the mission, the total AO
fluence on a given external surface of the satellite is obtained by
atoms
atoms
00
00
_ AO
FAO
(2)
F
· tMission fsg · cos θ
cm2
cm2 · s
In this calculation, t is the total mission time expressed in seconds,
and θ is the angle between the ram direction and the normal to the
given surface. Thus, AO fluence is calculated for the total mission for
each of the external surfaces of the satellite. The fluence on the side
with θ 90 deg is about 4.5% of the highest fluence in the ram
direction (θ 0). Note that, in the calculations so far, the atomic
oxygen content of the atmosphere in the region of the orbit during the
specified mission period, total mission time, the average velocity of
the satellite in orbit, and the orientation of the various external
surfaces of the satellite with reference to the direction of motion of the
satellite are considered. No other factors, for example, the actual
surface area of the satellite, materials of the external surfaces, and the
various thermal loads on them, have been taken into account.
A measure of AO erosion susceptibility of a material is AO erosion
yield Eη , which is defined for a given material as the volume eroded
by exposure to and reaction with one atom of oxygen, thus giving it
the unit centimeters cubed per atom (oxygen). Values of erosion yield
for various materials are experimentally obtained, with most data
coming from the experiments on NASA’s Space Shuttles and Long
Duration Experimental Facility (LDEF), the latter a material test
satellite program in LEO during early 1980s. AO erosion yield for a
given material is obtained from the following relation:
mfgg
3
cm
ρfg∕cm3 g
Eη
00
(3)
atom
FAO fatoms∕cm2 g · Afcm2 g
where A is the exposed area of the surface, m is the mass of the
material eroded or lost, and ρ is the density of the material.
The AO erosion thickness of a material is calculated for a given
mission as the product of AO erosion yield Eη and the AO fluence for
that surface given by
Et fcmg Eη
cm3
atoms
00
· FAO
atom
cm2
(4)
The erosion thickness coming as it does from measured data in AOrich LEO give a reasonably accurate prediction of the suitability of a
given material for the satellite exterior. Degradation of thermo-optical
properties, solar radiation absorptivity and infrared emissivity (Δα
and Δε), are also measured along with the AO erosion yield Eη . These
data and the evaluations based on them together lead to appropriate
material choice and design. The best value of Et from the standpoint
of AO resistance is zero, meaning the measured value of Eη for that
material is zero in LEO environment exposure experiments. A
caution, however, with the data is that there is a larger uncertainty in
the evaluation of AO density. It is not simultaneously measured with
AO erosion yield in the experiments. Instead, it is evaluated from an
analysis of such factors as the altitude, solar activity, and so on, as
described previously, not all of which can be accurately quantified by
analysis alone. For a detailed discussion on the difficulties of accurate
measurement of atomic oxygen density and comparison with the
modeled data, see Hedin et al. [12].
AO erosion yields can be conveniently expressed as a multiplication factor Ef of 10−24 cm3 ∕O atom as
Eη Ef × 10−24 cm3 ∕atom
(5)
where the nondimensional Ef may be stated as the AO erosion yield
factor for the given material. It must be noted that here we have
introduced the term “AO erosion yield factor” as an “engineering rule
of thumb”. The formal term in scientific literature is “AO erosion
yield”, as given by Eq. (3).
Banks [13] summarized AO erosion yield data for various
materials from the LDEF experiments. More recently, the Polymer
Erosion and Contamination Experiment was flown as part of
the Materials International Space Experiment 2; see [14]. For this
experiment, 40 different materials were tested. The selection
represented a variety of polymers used in space and different
chemical structures. Carefully measured data were obtained for
dehydrated mass, density, fluence, and erosion yields. The data were
used to develop a predictive tool for LEO erosion yields of new
polymers based on their chemical structure and laboratory test data
[15,16]. It helps to avoid in-space LEO testing of every new polymer
developed. The predictive tool is applicable to all polymers that
contain any of the following elements: carbon, hydrogen, nitrogen,
oxygen, fluorine, chlorine, and sulphur. Table 1 provides exhaustive
data of erosion yields of various materials provided by Banks [13],
Banks et al. [15,16], and [17].
Materials in the first column (Ef ≤ 0.1) have good resistance to
AO. A few in the second column may also be acceptable. Silverman
[6] compiled detailed data of erosion yields on polymers, composites,
metals, lubricants, coatings, and paints along with other materials.
For polymers, the recent data of Banks et al. [15,16] are
recommended.
The erosion thickness Et may also be expressed in terms of the
erosion yield factor by combining Eqs. (4) and (5) as
Et fcmg Ef × 10−24
cm3
atoms
00
· FAO
atom
cm2
(6)
To calculate the erosion thickness of a given material if it were present
on a given side of a satellite for the specified mission, the fluence for
that surface is multiplied by the measured value of the erosion yield.
The erosion thickness is also expressed as a number in micrometers
given by
tE fμmg Et fcmg ×
104 fμmg
1fcmg
(7)
The erosion thickness does not indicate the chemical reaction that
might have taken place at the surface of a given material exposed to
AO. It indicates only the thickness of the eroded material: mass blown
away. In the case of metals, reaction with AO produces oxides, which,
when subjected to the day and night thermal cycling, may result in
the erosion of the surface. A few metals erode more than others
depending on their susceptibility to AO. For example, silver and
osmium readily react with AO. For some other metals, oxides form a
thin coating on the surface, thereby inhibiting the exposure of the
interior layers. The surface oxide formation, however, may degrade
thermo-optical properties. Titanium in general is AO-resistant.
Titanium foil can be used as a single layer near high-temperature
regions like thrusters. In general, a metallic foil is not suitable for the
rest of the outer surfaces because of less-than-adequate properties as
well as the high surface temperatures it may attain.
Organic materials and coatings significantly react with AO while
pure carbon less so. Polyimides (for example, Kapton) are highly
reactive to AO. The erosion thus caused exposes the interior layers for
further reaction and erosion while degrading the thermo-optical
properties. AO and solar UV radiation together interact with the
external surface in a complex way, bleaching and radiation darkening
the material, thereby causing optical property degradation. In an AOrich environment, it is essential to predict the erosion and property
degradation of the external layer of the blankets and design the
material either to eliminate these effects or to mitigate them.
For a given four-year mission input, the AO density is calculated
from the NASA web model (MSISE 1990) to give a value of
ρAO 1.6 × 108 atoms∕cm3 . We wish to determine a suitable
754
ANGIRASA AND AYYASWAMY
Table 1
1.0–1.99
Different types of carbon
(0.5–1.7)
Crystalline polyvinyl-fluoride with
Acrylonitrile butadiene
white pigment (white Tedlar) (0.101) styrene (1.09)
Nb film, Ni film, Pt film,
PTFE or Teflon (0.142)
Aquadag E (graphite in
W, WC (0.0)
aqueous binder) (1.23)
YB-71 (ZOT) White Paint Perfluoroalkoxy copolymer resin
Polyvinylidene fluoride
(0.0)
(PFA) (0.173)
(Kynar) (1.29)
Ir film (0.0007)
Amorphous fluoropolymer (0.198)
Poly(p-phenylene-2 6bezobisoxazole) (Zylon)
(1.36)
Polyacrylonitrile (1.41)
SiO2 (650 Å) on Kapton H Urethane (black, conductive) (0.3)
(<0.0008)
401-C10 (flat, black) (0.3)
Polyphenyleneisophthalate
SiO2 (650 Å) with ≤ 4%
PTFE (<0.0008)
(Nomex) (1.41)
Indium Tin Oxide (ITO)
Siloxane polyimide (25% Sx) (0.3)
Polybenzimidazole (1.5)
(0.002)
Mo (1000 Å) (0.006)
Chemglaze Z306 (flat, black) (0.35) Polyurethane (1.56)
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<0.1
Al, Cu, Au, Pt, SiO2 , Pb,
Mo, Mg, Ni, Ta (0.0)
Nichrome (100 Å) (0.0)
AO erosion yield factor Ef
0.1–0.99
Al-Kapton (0.1)
2.0–4.0
Graphite epoxy (2.1–2.6)
>4.0
Polyetherimide (>3.31)
Polybenzimidazole
(>2.21)
Polysulfone (2.4)
Polyamide 6 or nylon 6 (3.51)
Polystyrene (3.74)
Polypropylene (2.68)
Epoxide or epoxy (4.21)
Mylar polyester (2.7–3.9)
Polycarbonate (4.29)
Polyimide (Kapton HN)
(2.81)
Polycarbonate (≫2.9)
Cellulose acetate (5.05)
Polysulphone (2.94)
Polyetheretherketone
(2.99)
Cr (125 Å) on Kapton H
(0.006)
Cu (1000 Å) (0.0064)
Fluorinated ethylene propylene
(Teflon FEP) (0.37)
Silverised Teflon (0.37) [17]
Epoxy (1.7)
Polyimide (Kapton H) (3.0)
Polystyrene (1.7)
TiO2 (1000 Å) (0.0067)
Black paint Z306 (0.3–0.4)
Ethylene-chlorotrifluoroethyline (Halar) (1.79)
Cu (1000 Å) on sapphire
(0.007)
MgF2 on glass (0.007)
White paint A276 (0.3–0.4)
Polyamide 66 or nylon 66
(1.80)
Polyimide (CP1) (1.91)
Polyethylene terephthalate
(Mylar) (>3.0)
High temperature
polyimide resin (PMR-15)
(>3.02)
Polyvinyl fluoride Tedlar
(clear) (3.2)
Polyethylene (3.3–3.7)
(>3.74)
Z302 (glossy black) (3.9)
ITO-Kapton-Al (0.01)
SiOx -Kapton-Al (0.01)
Al2 O3 (700 Å) on Kapton
H (<0.02)
Diamond (0.021)
Al2 O3 (<0.025)
Fluoro-polymers (0.03–
0.05)
Silicones (>0.05)
Electrodag 402 (silver in a
silicone binder) (0.057)
Pyrolytic graphite (0.415)
Polyvinylidene Fluoride (0.6)
Siloxane Polyimide (7% Sx) (0.6)
Tedlar (white) (0.6)
Polyethylene oxide (1.93)
Polymethyl methacrylate
(>5.6)
Allyldiglycolcarbonate
(CR-39) (>6.8)
Polyoxymethylene; acetal;
polyformaldehyde (Delrin)
(9.14)
Silver (>10)
Polyimide (Upilex-S) (9.22)
Osmium
Apiezon Grease (2 μm) (>0.625)
Poly-(p-phenyleneterephthalamide)
(Kevlar) (0.628)
Chlorotrifluoro-ethylene (Kel-f)
(0.831)
Polybutylene terephthalate (0.911)
Tetrafluorethylene-ethylene
copolymer (Tefzel) (0.961)
material as the outer cover of the MLI blanket. The procedure is as
follows.
To accommodate the model uncertainty in the calculation of
ρAO , a margin of 25% is added to this value so that ρAO 2.0 × 108 atoms∕cm3 . Taking the average satellite relative velocity
to be 7.6 km∕s, we obtain the AO flux as
atoms
00
5 cm
2.0 × 108
·
7.6
×
10
F_ AO
s
cm3
atoms
1.52 × 1014
cm2 · s
blanket on this surface of the satellite for the mission defined
previously. The evaluation of tE for ITO-Kapton-Al is detailed next.
From Table 2, we have Ef 0.01. Indium tin oxide (ITO) is
atomic-oxygen-durable, but pin holes in the coating cause mass loss
that appears as an erosion in the overall. Using Eq. (6) and with the
0 0 as calculated previously, the erosion thickness is
value of FAO
The following evaluation is done for the surface with the worst
case AO exposure (θ 0). The AO fluence for this surface is
calculated as
0 0 1.52 × 1014 atoms
· 4 × 365 × 24 × 60 × 60fsg · cos 0
FAO
cm2 · s
atoms
1.92 × 1022
cm2
For ITO-Kapton-Al on the given surface, for this mission,
Because this example evaluation considers only AO resistance and no
other factors (like surface electrical resistivity and thermo-optical
properties), we first take a list of commonly used materials for
blankets for satellites [18,19] as given in Table 2 and evaluate the
erosion thickness of each if it were present as the outer cover of the
Et
0.01 × 10−24
cm3
atoms
22
· 1.92 × 10
atom
cm2
0.0192 × 10−2 fcmg
Table 2
AO erosion thickness calculation for MLI outer cover
materials
Material
ITO-Kapton-Al
Beta cloth (woven fiberglass
impregnated with PTFE) [19]
Tedlar-Al
Teflon-silver-inconel
(SSM-radiator surface)
Ef
0.01
0.3 (PTFE eroded;
fiberglass intact)
0.6
0.37
tE ,
μm
∼2
25
Standard layer
thickness, μm
13–254
200
115
71
120
13–254
755
ANGIRASA AND AYYASWAMY
tE fμmg 0.0192 × 10−2 fcmg ×
104 fμmg
1.92
1fcmg
AO reaction evaluations, as illustrated previously, are used along with
other factors (thermo-optical properties, UV radiation effects, and
surface electrical conductivity) in choosing the appropriate materials
and their design for manufacturing and integration with the satellite.
In the following sections, more details of this procedure will be
provided. Useful information on the design aspects of MLI blankets
may be found in [20]. Information on standard practices for the
evaluation of material–atomic-oxygen interactions in ground
laboratory tests is provided in [21].
Downloaded by GA Inst of Technology on May 9, 2023 | http://arc.aiaa.org | DOI: 10.2514/1.A32742
V.
Plasma Environment and Electrostatic Discharge
To assess the effect of the plasma environment on spacecraft
materials, the vertical profile of the Earth’s atmosphere is first
classified based on temperature variation from the surface upward.
This is because the thermal structure of the low-Earth-orbit regions
has important influence on the physical processes. This classification
is based on atmospheric and meteorological nomenclature [22].
From the Earth’s surface upward, the temperature decreases in the
troposphere. Above this is the stratosphere, in which the temperature
increases with altitude. A temperature maximum separates the
stratosphere from the mesosphere. Within the mesosphere,
temperature decreases with increasing altitude. The top of the
mesosphere is the coldest part of Earth’s atmosphere (∼ − 90°C).
Above the mesosphere, at about ∼80 km height, the temperature
significantly increases with height, and this region is called the
thermosphere. The altitude range for the thermosphere is often stated
as 80 to 500 km from the point of view of satellite orbits, and above
the thermosphere is the exosphere (up to ∼1000 km). For a detailed
discussion of the composition, analysis, and models of the regions
of interest for satellites, see Montenbruck and Gill [23]. The
International Space Station (ISS) orbits in the middle of the
thermosphere, between 320 and 380 km above the Earth’s surface.
All meteorological processes of interest occur below the
thermosphere producing turbulent and neutral gas mixture. Above
the meteorological boundary of about 80 km height, solar radiation
produces partially ionized plasma composed of O, N2 , O2 , He, H,
−
O , H , He , NO , O
2 , N2 , and e . In the thermosphere and
exosphere, the most abundant neutral gas is atomic oxygen (AO), as
discussed previously, and the most abundant ion is O . Above the
exosphere, H and He are dominant. Plasma is a semi-electrically
neutral gas of charged and neutral particles, exhibiting collective
behavior: a “plasma cloud”. Earth’s magnetic field and solar wind
control the movement of the cloud, which generates its own electric
and magnetic fields. These, in turn, affect the movement of the
plasma cloud and also the motion of charged particles far away.
The Earth atmospheric regions are also classified based on
electrical conductivity and electron density. The ionosphere is
defined as the region from ∼85 to 600 km altitude, characterized by
ionization by solar radiation. Magnetosphere is the region formed by
the interaction of Earth’s magnetic field and the solar wind. The
properties of Earth’s ionosphere and magnetosphere and the solar
wind are important for understanding the satellite-materials–orbitenvironment interactions. A lucid introduction to these regions and
properties is given by Hargreaves [24].
The composition and evolution of space plasma environment over
time are very complicated; see, for example, DeForest and McIlwain
[25]. The usual practice to describe a plasma state is to give the values
of a temperature in electron volts and a density expressed in terms of
the number of particles per unit volume of centimeters cubed. These
and other properties and plasma distributions are obtained from
various models. In particular, the global international reference
ionosphere model (IRI-2000, for example) provides estimates
of plasma concentrations, composition, and temperatures, under
varying solar activity; see Bilitza [26]. IRI is the recommended
standard by both NASA [3] and ESA [9].
Plasma interactions with the spacecraft are complicated, depending on the component of spacecraft, passive and active (e.g., solar
array and connected structure), and the region of the satellite orbit. In
general, the surfaces of the spacecraft accumulate electric charge to
be in equilibrium with the plasma. This electrostatic buildup process
is called spacecraft charging. In general, two types of charging are
visualized. Absolute charging occurs when the entire spacecraft
potential relative to the ambient plasma uniformly changes by
interacting with the charging environment. The second type, called
differential charging, occurs when parts of the spacecraft are charged
to different negative potentials with reference to one another. In this
case, strong local electric fields may develop. The static build-up
must be inhibited by appropriate measures of grounding the external
surfaces of the satellite such as MLI and radiators and employing
surface materials of adequate electrical conductivity or low resistivity
or resistance. If the charge build-up is large enough, it will discharge
through the surface layer/finish to the satellite structure. This is called
electrostatic discharge (ESD) and can damage the electrical and
electronic components or otherwise disrupt their functioning.
Absolute charging by itself is not a serious ESD concern. However,
it increases the surface contamination, which degrades thermooptical properties. It also compromises science missions. Spacecraft
surfaces are not uniform in their properties. Furthermore, the surfaces
may have exposure to different environments, and the ambient fluxes
may be anisotropic. These and other charging effects can produce
potential differences between spacecraft surfaces and between
spacecraft surfaces and spacecraft ground. When the potential
difference exceeds a threshold, ESD will occur. With multiple
charges, the satellite may wobble. Ultimately, the satellite operation
may be disrupted.
Because of the frequency and severity of plasma–spacecraft
interactions, much effort has been expended in developing computer
models to quantify spacecraft charging that help in the mitigation
strategies, design, and materials choice. In the United States, a
comprehensive code called NASA/Air Force Spacecraft Charging
Analysis Program is available for U.S. citizens. In Europe, ESA’s
space environment information system provides a suite of models
for predicting space environment and spacecraft–environment
interactions.
To characterize the required surface electrical conductivity (or less
resistivity), the concept of surface resistance or sheet resistance is
used. For a three-dimensional conductor (for example, a rod, see
Fig. 3), the electrical resistance is defined as
RfΩg ρfΩ · mg
Lfmg
Afm2 g
(8)
where R is the electrical resistance, ρ is the electrical resistivity, L is
the length of the conductor, and A is its cross-sectional area.
For a two-dimensional surface on the spacecraft such as a
coating or the outer layer of a blanket, the electrical resistance in the
plane of the layer (not perpendicular) is defined by the following
relation:
RfΩg ρfΩ · mg
Lfmg
Wfmgtfmg
(9)
W
t
L
A
L
Fig. 3 Concept of surface electrical resistance.
756
ANGIRASA AND AYYASWAMY
The appropriate dimensions are shown in Fig. 3. The thickness t ≪ L
or W, and hence it may be considered as two-dimensional in L and W.
Following this, Eq. (9) may be rewritten as
RfΩg ρfΩ · mg Lfmg
Lfmg
·
RS fΩg ·
tfmg
Wfmg
Wfmg
(10)
Downloaded by GA Inst of Technology on May 9, 2023 | http://arc.aiaa.org | DOI: 10.2514/1.A32742
in which RS is called the surface resistivity or sheet resistance. Note
that R and RS have the same units in ohms. However, to characterize
RS as a surface quantity, L can be set equal to W, and we may define
the units of the surface resistivity as
Ω
≡ Ω∕□
(11)
RS ≡
square
This makes RS independent of the units of the length dimension
(meters, inches, or centimeters). In a physical sense, sheet resistance
denotes the resistance of the layer over a square area of its surface
independent of the units chosen for the area. Note that surface
resistivity is not bulk electrical resistivity divided by the layer
thickness. Measurement techniques for surface resistivity are given
by Schroder [27].
Purvis et al. [28] have provided spacecraft design guidelines to
harden spacecraft systems to charging and consequent effects.
Although these were primarily meant for satellites in GEO, there is
also much relevant information on materials evaluation for LEO.
Because the focus of the current review is evaluation of exterior
materials in LEO, attention is focused on the required and
recommended surface characteristics of the satellite exterior
materials: primarily solar cell cover slides, the outer cover of MLI
blankets, and radiator surface treatments. For this purpose, the
differential charging as discussed previously is an important
consideration. For spacecraft control, all of its exterior surfaces need
to be at least partially conductive and grounded to the spacecraft
structure. This is quantified by a maximum allowed surface electrical
resistivity as defined earlier (RS ). However, as discussed in Sec. III,
satellite surfaces contain much insulating materials, which are good
thermal insulators. Thermal insulators are also electrical insulators.
Therefore, to improve their ESD control characteristics, the outer
covers are often coated to provide the desired surface electrical
conductivity.
As an illustrative example, consider a thin indium tin oxide (ITO)
coating that is applied over Kapton (see Table 2). Note that ITO is also
resistant to atomic oxygen exposure. Purvis et al. [28] recommend the
following guidelines for evaluating the surface conductivity
requirements. Conductive materials (e.g., metals) must be grounded
to structure with the smallest resistance possible such that
R · A < 109 Ω · cm2
(12)
where A is the exposed surface area of the conductor in centimeters
squared. Partially conductive surfaces (e.g., paints) applied over a
conductive substrate must have the following resistivity-thickness
product:
ρ · t ≤ 2 × 109 Ω · cm2
(13)
where ρ is the material (i.e., paint) resistivity in ohm-centimeters, and
t is its thickness in centimeters. Partially conductive surfaces applied
over a dielectric and grounded at the edges must have material
resistivity such that
ρ · h2
≤ 4 × 109 Ω · cm2
t
(14)
where h is the largest distance from a point on the surface to a ground
point given in centimeters.
The previous guidelines are dependent on particular geometry and
application. They are simplified and made independent of geometry
and in the process more stringent as given in Table 3.
Spacecraft exterior materials have to be chosen such that the
previous conditions are fulfilled. Grounding methods must be able to
withstand the current bleed-off from ESD, vacuum exposure, thermal
expansion and contraction, and other events.
NASA has issued specific LEO spacecraft charging guidelines in
recent years for high-voltage (HV) space power systems (>55 V)
operating in a plasma environment; see Ferguson and Hillard [29].
With reference to such a HV system, the phenomenon of “snap-over”
has to be considered. Snap-over may be understood as follows.
Consider a flat conducting plate covered with an insulator containing
a pinhole. If the plate is biased by a power supply and exposed to
space plasma environment, current will flow to the plate. For low
voltages, this current is linear with bias voltage. The plate acts as
a source of increasing electric field. This field produces ion
bombardment of the insulator and secondary electron emission. The
result is a rapidly growing sheath collecting charge and funneling
it to the pinhole. As some critical point is reached, the current
flow increases exponentially. This shift to an exponential current
collection is called snap-over. Snap-over poses a threat of damage to
the surfaces as well as the possibility of a significant loss of power.
For example, on a solar array, interconnects, or wire traces, cell edges
act like pinholes. They perform the role of conductors to which
current is funneled. The solar cell substrate or the cover slides act like
the insulator as described previously, furnishing the secondary
electrons and becoming a current collecting plate.
Ferguson and Hillard [29] recommend that, for such HV power
systems, a major design consideration is to avoid snap-over. In such
cases, the strategies may include insulating all surfaces, if possible,
and choosing insulators with low secondary electron emission
characteristics.
The differential charging of the spacecraft exterior surfaces can be
offset by careful consideration and choice of available materials. The
current practice is to make all surfaces conductive to the extent given
by the requirements in the right-hand column of Table 3 and
connecting them to a common ground. Surface coatings are widely
employed to obtain the desired level of electrical conductivity;
conductive conversion coatings on metallic surfaces, conductive
paints, and transparent partially metallic vacuum-deposited films
(e.g., indium tin oxide, ITO) serve the purpose. Purvis et al. [28]
recommend the coatings and other surfaces materials given in
Table 4, which have had successful space use, while Table 5 provides
a list of materials to avoid, if possible. Detailed evaluations and
recommendations of materials and surface treatments for MLI outer
covers and radiator surfaces are presented in subsequent sections.
Materials used as conducting surfaces on the spacecraft are given
in Table 6 under various categories.
Because the choice of materials depend on the configuration of
various applications, grounding techniques correspondingly differ.
For many useful suggestions on grounding techniques to ensure
reliable performance during spacecraft flight and orbit, see [28].
If the spacecraft surface cannot be made fully conductive, an
analysis needs to be performed based on the materials used to
demonstrate that the design is ESD proof. The charging and ESD
characteristics of dielectric materials vary. Thus, a particular
dielectric material can, in principle, significantly affect the surface
voltage profiles. For example, cerium-doped microsheet charges to
much lower potentials under electron irradiation than fused silica,
and it may be preferred as a solar array cover-slide material [30]. All
Table 3
Guidelines for avoiding ESD
Material type and characteristic
Requirement
Isolated conductors must be grounded to structure such
R < 106 Ω
that
Materials applied over a conductive substrate must have
ρ < 1011 Ω · cm
bulk resistivity
Materials applied over a dielectric area must be grounded RS < 109 Ω∕□
at the edges and must have a surface resistivity (GEO)
For LEO polar applications, the last condition above is
RS < 108 Ω∕□
changed to
757
ANGIRASA AND AYYASWAMY
Table 4
Material
Carbon black paint
ITO coating
∼250 nm
Zinc ortho-titanate
(white paint)
Alodyne
Table 5
Comments and recommendations
In addition to ESD requirements, other needs to be
met: thermal properties, adhesion
For use on solar cells, optical solar reflectors, and on
Kapton foils; must be properly grounded; can be used
where some transparency is needed
Highly conductive; careful application is required as
adhesion is difficult
Conductive conversion coatings of magnesium,
aluminum, etc are acceptable
Fiberglass material
(e.g., β cloth)
White paints
Mylar (uncoated)
Teflon
Silica cloth
Quartz and glass
surfaces
Comment
Produces a surface with high resistivity; to be
avoided. If stored energy is proven to be small,
acceptable.
Resistivity is too high.
In general, not acceptable. However, some have
good surface conductivity. Surface resistance to be
measured before use is considered.
Resistivity is too high. Not recommended.
Resistivity too high. Demonstrated long-term
energy storage capability. Can cause catastrophic
damage.
Was used in antenna housing. It is a dielectric, but
because of fibers, or used with embedded conductive
materials, ESD sparks may be small.
Solar cell cover slides and second surface mirrors
have no ESD avoiding substitutes. ESD tests must be
performed to determine the effects on neighboring
electronics.
dielectrics on the spacecraft surface must be carefully examined for
their differential charging. Each dielectric region must be assessed for
its own breakdown voltage, its energy storage capability, and the
disruption and damage it can cause on neighboring electronics and
erosion and contamination of surface materials.
VI.
The thermal performance of an MLI blanket is defined by an
effective emittance ε or an effective thermal conductivity Keff . The
radiative heat transfer Q (in watts), through the blanket with an area A
(in meters squared), is given by
Q ε AσT 4H − T 4C Materials for Insulation Outer Cover
The primary purpose of the MLI blankets is to limit the radiative
heat transfer from and to the satellite. They also protect it from
micrometeoroids, AO, electron charge accumulation, and rocketengine plume impingement. The outer layer of MLI must be AOresistant for protecting the inner layers from exposure to AO. The
interior layer materials of the blankets do not necessarily need to be
AO resistant if the outer layer is adequately resistant to AO reactions.
Table 6
Materials for conducting surfaces
Categories
Vacuum metalized dielectrics in
the form of sheets, or strips, tiles
Thin, conductive front-surface
coatings
Conductive paints or fog (thin)
Exposed conductive face-sheet
materials
Metallic foils or metalized plastic
film tapes
Etched metal grids, or bonded,
heat embedded metal meshes on
nonconductive substrates
Various conductive adhesives
Materials
The metal on the substrate list includes
aluminum, gold, silver, and Inconel on
Kapton, Teflon, Mylar, and fused silica
ITO, SiOx on Kapton, or Teflon,
dielectric stack, fused silica
Carbon-filled Teflon, or carbon-filled
polyester on Kapton
Graphite, or epoxy, metal
Aluminum, titanium
(15)
where ε is the effective emittance, or “E-star” of the blanket; and
T H and T C are the hot and cold side temperatures, respectively,
in Kelvins. The Stephan–Boltzmann constant is σ 5.67 ×
10−8 W∕m2 · K4 . The heat transfer Q may also be related to an
effective thermal conductivity K eff of the blanket of thickness l,
given by
ESD concerns on exterior materials and coatings
Material
Anodyzing
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Exterior surface materials and coatings
QfWg Keff
W
TH − TC K
· Afm2 g ·
m·K
m
l
(16)
The two quantities ε and Keff are thus related by
ε K eff T H − T C lσT 4H − T 4C (17)
Effective blanket thermal performance is related to minimizing the
value of ε or Keff . Because these two quantities are directly
interrelated as given previously, minimizing one of them leads to the
same effect. The values of ε or K eff are experimentally obtained for
blankets of varying area and number of layers with a given material
combination and over a wide range of temperatures. Correlations are
available for the thermal design of blankets; see, for example, [19].
For the outer cover of an insulation blanket (i.e., the layer exposed
to the space environment), the required surface thermal radiation
properties are characterized by the solar absorptivity α and the
infrared emissivity ε. The outer cover should absorb little solar
radiation incident on it and should be a good emitter of infrared
radiation. It is also common to state this requirement in terms of a
ratio of α∕ε, so that a combined effect of minimum solar absorptivity
and maximum emissivity may be realized.
The outer layer of the insulation blanket must have adequate
surface electrical conductivity. In addition, it must be AO-resistant.
Any erosion of the layer should not lead to the degradation of its
thermal radiation properties and its surface conductivity.
Uncoated polymers (e.g., Kapton) readily erode with AO exposure
and are therefore not suitable to be the outer layer of the external MLI.
Fluoro-polymers (e.g., Teflon) have low AO reaction rates. However,
these tend to lose all mechanical strength with exposure to plasma and
thermal cycling. In addition, Teflon is a difficult material for electrical
grounding purposes. Mylar is highly reactive with AO, and so is
polyethylene. Also, Mylar is highly susceptible to UV exposure.
The desired properties of the outer layer, viz., AO resistance,
thermo-optical properties, and surface electrical conductivity are
obtained by suitable coatings over a substrate. Before selecting a
coating, it must be first established that the coating material and the
substrate material are chemically compatible in the operational
temperature range for the mission. Some substrate materials with
high AO reaction may first be considered with the view that a suitable
coating may be applied that is AO-resistant while providing other
desirable properties.
In Table 7, various MLI outer-layer materials are given together
with their properties: AO erosion yield factor Ef , surface electrical
resistivity RS, beginning of life (BOL) solar absorptivity α and
infrared emissivity ε, change in their values due to material exposure
to AO, ultraviolet radiation, plasma, and thermal cycling. The sources
for the data are AO erosion yield [13,15,16] and other properties [6].
Woven fiberglass is called β cloth. It is impregnated with polytetra-fluoro-ethylene (PTFE), popularly known as Teflon, which
improves the strength and mechanical handling of the blanket.
Although fiberglass is unaffected by atomic oxygen, PTFE is known
to erode. The solar reflectivity ρS of β cloth is ∼0.7. Combined with
its high emissivity (0.9), it has the desired thermo-optical properties
[31]. Aluminum is vacuum-deposited on the interior face. The
758
ANGIRASA AND AYYASWAMY
Table 7
Downloaded by GA Inst of Technology on May 9, 2023 | http://arc.aiaa.org | DOI: 10.2514/1.A32742
Outer layer materials
β cloth (fiberglass)
Siloxane polyimide (25% Sx)
SiOx -Kapton-VDA
ITO-Kapton-VDA
SiO2 (650 Å) -Kapton
Al2 O3 -Kapton
Germanium-Kapton
Insulation blanket outer layer properties
Surface
resistivity
RS , Ω∕□
Highly resistive
AO erosion
yield factor Ef
0
0.3
0.01
0.01
<0.0008
<0.025
∼0
<104
∼107
process, however, is difficult because of PTFE. β cloth has very high
surface resistivity, and ESD is a concern. Conductive coatings (ITO,
SiO2 etc) are difficult to apply.
Kapton without outer coating is not suitable as the outer layer of the
insulation blanket because of its high AO erosion yield factor of 3.0
[32]. Siloxane polyimide (25% Sx) has low AO erosion yields. Its
properties, however, are not readily available.
Polyimide Kapton is often used as the baseline material for the
flexible solar array panels (e.g., Hubble Space Telescope). Coatings
for Kapton were developed to protect it from atomic oxygen reaction
[33,34] while providing the requisite surface electrical conductivity
and thermo-optical properties. Polycrystalline ceramic films, such as
SiOx where 1.9 < x < 2.0, SiO2 , and Al2 O3 were space-tested and
found to be effective in protecting Kapton from atomic oxygen
erosion in LEO [35–37]. ISS Alpha used SiOx -Kapton for its solar
array panels.
Sputter-deposited Germanium on Kapton provides resistance to
AO and has adequate surface electrical conductivity. Nominal
coating thickness is 1500 Å. Ge-Kapton is used as an interstitial layer
between photovoltaic cells and the face sheet on solar array panels.
Useful information on surface finishes for thermal control materials
is available in [38].
The choice of selecting the suitable AO resistant insulation blanket
cover with all the desirable properties is a challenging task in longduration polar LEO satellite system design.
VII.
Radiator Surfaces
The basic purpose of a spacecraft radiator is to radiate waste heat
generated from various heat dissipaters away into space. The
radiators may be of various types: spacecraft structure, or projected
panels, panels deployed in space. Whatever be the type, the radiator
surface is treated to obtain the desired thermo-optical properties of
minimum solar absorptivity and maximum infrared emissivity. In
addition, in LEO, the surfaces should be resistant to atomic oxygen
and, for polar orbits, have adequate electrical conductivity. The
radiative heat transfer is dependent on the surface temperature, and
the thermal control system design has to take into account a good
conductive path from the location of the heat dissipater to the
radiative surface. The smaller the temperature drop between the
dissipater and the radiator, the more efficient is the thermal control
system. For design aspects of radiators, see [39]. Radiator surface
finishes are mostly SSMs and white paints. Ceramic coatings,
which have similar properties as white paints, may also be used as
Table 8
Solar
absorptivity
α (BOL)
0.33
Infrared
emissivity
ε (BOL)
0.9
0.195
0.44
0.33
0.34
0.483
0.85
0.62
0.72
0.71
0.617
Δα
0.14 ↑
Δε
0.001 ↓
0.005 ↓
3% ↑
0.01 ↑
0
0.002 ↑
0.003 ↓
Comments
ESD concern
Properties not available
ITO fragile
appropriately needed. Some of the earlier discussions for MLI outer
covers are also valid for radiator surface treatments [38].
In Table 8, SSM properties are presented. There are extensive flight
test data for Teflon-Silver SSM without any outer coating on Teflon
[6]. The differences among the various tests are the duration of the
LEO environment exposure, AO fluences, UV exposure and the
number of thermal cycles. An equivalent sun hour (ESH) is one hour
of solar irradiation at its maximum intensity. It is a measure of the
magnitude of the material’s exposure to UV radiation.
The AO erosion yield factor for FEP Teflon is higher than it is for
Teflon (PTFE) in long-duration experiments [18]. FEP Teflon also
is highly resistive, and without a conductive coating, ESD is a
concern. It has very good thermo-optical properties with acceptable
degradation over long-duration exposure in LEO environment. The
solar reflectivity of silvered FEP Teflon SSM, ρS (BOL), is 0.92 with
a degradation of 0.01 (↓). The reflection tends to be more diffuse at
end of life. The combined effect of AO and SUVexposure makes both
PTFE and FEP Teflon brittle. This feature, combined with somewhat
higher value of AO erosion yield factor, should be considered in
choosing the appropriate thickness of the FEP Teflon layer along with
a coating to provide the mechanical strength, AO and UV resistance,
and adequate surface electrical conductivity. Coatings on FEP Teflon
include ITO, SiOx , and In2 O3 . Of these, only ITO coated Teflon
SSMs seem to be available, although all the stated coatings appear to
have the desired properties. SiOx -coated Kapton is available, which is
more suitable for use in insulation blankets than in radiators.
Organic paints are not suitable in LEO for the radiator surface
finish because of their ready reaction with AO. Inorganic white paints
have flight heritage. Table 9 lists the properties of three suitable white
paints for LEO environment. Important considerations in using the
white paints are ease and uniformity of application, thermochemical
compatibility with the substrate material, and surface electrical
conductivity. The solar absorptivity of white paints tends to increase
with the LEO environment exposure. The solar reflection of the
paints is diffuse.
Ceramic materials by themselves or coatings on a substrate may be
used as radiators. TiO2 and ZnO2 are resistant to AO. Surface
electrical conductivity data, however, are not readily available.
Table 10 lists their thermo-optical properties [6].
A. Rigid Second Surface Mirror
As described in Sec. III, rigid second surface mirrors based on
glass or quartz are widely used as solar cell cover slips. They are rigid
in the sense that they can be attached only to flat surfaces. Table 11
lists the properties of fused-silica-based SSM [31]. They are tolerant
Radiator surface finish properties: SSMs
Solar
Infrared
Surface
absorptivity emissivity
resistivity
AO erosion
α (BOL)
ε (BOL)
Δα
Δε
Comments
RS , Ω∕□
SSM materials
yield factor Ef
FEP Teflon (∼125 μm) — silver — Inconel
0.37
Highly resistive
0.07
0.81
0.03 ↑ 0.03 ↓ ESD concern; teflon difficult material
FEP Teflon — aluminum
0.14
0.75
No flight data
0.14
0.8
0.04 ↑ 0.009 ↓ Available
ITO-FEP Teflon — silver — Inconel
0.01
<104
<2000
Known to have adequate properties
In2 O3 - FEP Teflon — Silver
0.19
0.75
No flight data
ITO-FEP Teflon — aluminum
<104
Coating known to be possible; no data
SiOx –FEP Teflon- silver
759
ANGIRASA AND AYYASWAMY
Table 9
Inorganic white
paint
PCB-Z
Z-93
YB-71
a
AO erosion yield
factor Ef
∼0
∼0
∼0
Surface resistivity RS,
Ω∕□a
<103
<103
Solar absorptivity α
(BOL)
0.27
0.14
0.13
Infrared emissivity ε
(BOL)
0.87
0.90
0.85
Δα
Δε
0.08 ↑ 0.006 ↑
0.01 ↑ 0.01 ↑
0.02 ↑ 0.01 ↓
Reference
[31]
[40]
[40]
Suppliers’ data.
Table 10
Ceramic
material
TiO2
ZnO2
Radiator surface properties: ceramics
Solar absorptivity α (BOL) Infrared emissivity ε (BOL)
0.2
0.9
0.29
0.78
Table 11
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Radiator surface finish properties: inorganic white paints
In addition, it also receives IR radiation from Earth. From the surface
thermo-optical properties given in Table 8, the maximum equilibrium
surface temperature T S can be determined as follows. At the surface,
the energy balance for a unit radiator surface area (Fig. 5) is given as
Rigid SSM surface properties
Rigid SSM material
Fused silica-silver-Inconel
Solar reflectivity Infrared emissivity
ε (BOL)
ρS (BOL)
0.94
0.8
of atomic oxygen and UV radiation. Contamination related
degradation of optical properties, however, is a concern, and
appropriate measures need to be taken.
qabsorbed qconduction qemitted
(18)
αG εI Q_ 0 0 εσT 4S
(19)
in which G is the combined solar heat flux and the albedo from
Earth, and I is the infrared radiant flux from Earth. These values
are available from the mission profile. The conduction heat flux from
the interior heat dissipaters in the spacecraft is an input value for
the thermal design. The equilibrium surface temperature then is
obtained as
TS 1
αG εI Q_ 0 0 4
εσ
(20)
B. Illustrative Evaluation of Radiator Materials
In Sec. IV, we have carried out an illustrative evaluation of MLI
outer cover materials for atomic oxygen resistance for a given
mission. Here, we wish to evaluate spacecraft radiator materials for
the mission. In the process, some of the system and mission level
matters will be examined.
The widely used silvered Teflon is considered with an appropriate
coating. ITO was evaluated before for AO reaction with reference to
MLI outer cover coating. The worst-case erosion thickness was
calculated to be <2μm. Note, however, that even this mild erosion is
caused by contaminants, but not due to AO. A coating of ITO with
adequate margin of thickness will provide protection of the radiator
surface from AO for this mission (Fig. 4).
Assuming that this is a polar LEO mission, the condition for
surface electrical resistance given by the final requirement in the
right-hand column of Table 3 applies. From Table 8, we find that this
condition is met. All the components of a spacecraft, subsystems,
instruments, equipment, and so on, have temperature limits for
on-ground, in-flight, and operational phases of the mission. The
temperature limits of various spacecraft elements are determined
based on their function, materials properties, and mission profile. For
the current example, say, the radiator temperature limits are stated as
between −40 and 65°C. The supplier of the materials would have to
demonstrate that they would confirm to these temperature and all
other requirements of the mission.
The worst-case high temperature of the radiator surface is when it
receives the highest heat flux from the sun and the albedo from Earth.
Fig. 4
The radiator surface area and configuration are determined at system
level by the locations of the heat dissipaters, conduction paths to the
radiators and the heat loads.
Equation (20) shows the important and complex connection
between the thermo-optical properties of the radiators that are
influenced by the LEO environmental factors of AO, UV radiation,
and plasma through their effects on the materials during the mission
and spacecraft temperatures and thermal control. The “health” of the
satellite depends on these factors, hence the necessity to perform
careful and methodical evaluation of materials.
Fig. 5
Heat balance at unit surface area of radiator.
Radiator surface materials: example.
760
ANGIRASA AND AYYASWAMY
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VIII. Conclusions
This review is an attempt to synthesize the knowledge accumulated
over the past two decades on the significant effect of the lowEarth-orbit environment on the spacecraft exterior materials in a
manner spacecraft system development engineers can readily use.
Evaluations of materials due to the exposure to atomic oxygen,
ultraviolet radiation, and plasma have been formalized based on the
extensive data of long-duration exposure facility, space shuttles, and
other flight data.
Low-Earth-orbit space environment is described in terms of the
components and parameters that affect the materials. Analyses,
models, and correlations to quantify various environmental parameters and their interactions with the materials are given. The factors
that significantly interact with satellite materials and degrade their
performance are identified as ultraviolet radiation, atomic oxygen,
and plasma–spacecraft interactions. Satellite equipment/subsystems/
components that are exposed to the space environment are described,
and the materials used in their manufacture are identified.
Based on the extensive space measured data, a detailed sequential
quantification methodology is proposed to choose materials and
design satellite hardware that withstands the chemical reactions with
atomic oxygen in low Earth orbit in combination with ultraviolet
radiation. This takes into consideration the current difficulties in
measuring the density of atomic oxygen in the low Earth orbit
simultaneously with the erosion it produces on materials. Extensive
data on atomic oxygen–ultraviolet-radiation interactions with
spacecraft materials in a readily usable format are presented.
The physics of spacecraft–plasma interactions is described, and
the resultant effect on the satellite, electrostatic discharge, is shown to
be damaging to the functioning of materials, devices, and the satellite
itself. The material characteristics and mitigation strategies are
quantified through measurable parameters. Spacecraft exterior
materials are evaluated through this methodology for their suitability
in multilayer insulation blankets, radiator surfaces, and solar cell
cover slides. Again, readily usable data for many materials are
presented for their evaluation. A realistic low-Earth satellite design
example is worked out to demonstrate the proposed methodology in
the evaluation and choice of materials for the insulation outer cover
and radiator surfaces.
The methodology and the evaluations have shown where
developmental efforts might be required, mainly flight qualification
of coatings and white paints. Contamination related degradation of
thermo-optical properties is a concern to be addressed in the
manufacturing, assembly, and spacecraft integration phases.
Acknowledgment
Bruce Banks (NASA Glenn Research Center) provided most of the
data on the atomic oxygen exposure of materials as well as much
insight on the subject. His contribution is gratefully acknowledged.
References
[1] Wertz, J. R., Mission Geometry: Orbit and Constellation Design and
Management, Microcosm Press, El Segundo, CA, 2001.
[2] Wertz, J. R., Everett, D. F., and Puschell, J. J., (eds.), Space Mission
Engineering: The New SMAD, Microcosm Press, Hawthorne, CA,
2011.
[3] Anderson, B. J., and Smith, R. E., “Natural Orbital Environment
Guidelines for Use in Aerospace Vehicle Development,” NASA TM4527, 1994.
[4] Visentine, J. T., and Whitaker, A. F., “Material Selection Guidelines to
Limit Atomic Oxygen Effects on Spacecraft Surfaces,” NASA TM100351, 1989.
[5] Dooling, D., and Finckenor, M. M., “Material Selection Guidelines to
Limit Atomic Oxygen Effects on Spacecraft Surfaces,” NASA TP209260, 1999.
[6] Silverman, E. M., “Space Environmental Effects on Spacecraft: LEO
Materials Selection Guide: Parts 1 and 2,” NASA CR-4661, 1995.
[7] Tascione, T. F., Introduction to the Space Environment, 2nd ed., Krieger
Publ., Malabar, FL, 1994.
[8] Smith, R. E., and West, G. S., “Space and Planetary Environment
Criteria Guidelines for Use in Space Vehicle Development,” NASATM82478, Vol. 1, Sec. 2, 1983.
[9] “Space Environment,” European Space Agency, European Space
Research and Technology Centre, Rept. ECSS-E-ST-10-04C, Noordwijk,
The Netherlands, 2008.
[10] Jursa, A. S. (ed.), Handbook of Geophysics and the Space Environment,
4th ed., U.S. Air Force Geophysics Lab., Rept. ADA 167000, Hanscom
AFB, MA, 1985.
[11] Hedin, A. E., “MSIS-86 Thermosphere Model,” Journal of Geophysical
Research, Vol. 92, No. A5, 1987, pp. 4649–4662.
doi:10.1029/JA092iA05p04649
[12] Hedin, J., Gumbel, J., Stegman, J., and Witt, G., “Use of O2 Airglow for
Calibrating Direct Atomic Oxygen Measurements from Sounding
Rockets,” Atmospheric Measurement Techniques, Vol. 2, 2009,
pp. 801–812.
[13] Banks, B. A., “Atomic Oxygen,” Proceedings of LDEF Materials Data
Analysis Workshop, edited by Stein, B. A., and Young, P. R., NASA CP10046, 1990, pp. 191–216.
[14] de Groh, K. K., Banks, B. A., McCarthy, C. E., Rucker, R. N., Roberts,
L. M., and Berger, L. A., “MISSE PEACE Polymers Atomic Oxygen
Erosion Results,” NASA TM-214482, 2006.
[15] Banks, B. A., Backus, J. A., and de Groh, K. K., “Atomic Oxygen
Erosion Yield Predictive Tool for Spacecraft Polymers in Low Earth
Orbit,” NASA TM-215490, 2008.
[16] Banks, B. A., Backus, J. A., Manno, M. V., Waters, D. L., Cameron, K.
C., and de Groh, K. K., “Atomic Oxygen Erosion Yield Prediction for
Spacecraft Polymers in Low Earth Orbit,” NASA TM-215812, 2009.
[17] Levadou, F., Froggatt, M., Rott, M., and Schneider, E., “Preliminary
Investigations into UHCRE Thermal Control Materials: Part 2,”
Proceedings of LDEF — 69 Months in Space: First Post-Retrieval
Symposium, edited by Levine, A. S., NASA CP-3134, 1991, pp. 875–
898.
[18] Finckenor, M. M., and Dooling, D., “Multilayer Insulation Material
Guidelines,” NASA TP-209263, 1999.
[19] Kauder, L., “Preliminary Results for LDEF/HEPP Thermal Control
Samples: Part 2,” Proceedings of LDEF — 69 Months in Space: First
Post-Retrieval Symposium, edited by Levine, A. S., NASA CP-3134,
1991, pp. 797–800.
[20] Donabedian, M., Gilmore, D. G., Stultz, J. W., Tsuyuki, G. T., and Lin,
E. I., “Insulation,” Spacecraft Thermal Control Handbook Vol. 1:
Fundamental Technologies, 2nd ed., edited by Gilmore, D. G.,
Aerospace Press, El Segundo, CA, 2002, Chap. 5.
[21] “Standard Practices for Ground Laboratory Atomic Oxygen Interaction
Evaluation of Materials for Space Applications,” American Society for
Testing and Materials, Rept. ASTM-E2089-00, West Conshohocken,
PA, 2006.
[22] Goody, R., Principles of Atmospheric Physics and Chemistry, Oxford
Univ. Press, New York, 1995.
[23] Montenbruck, O., and Gill, E., Satellite Orbits, Springer-Verlag, Berlin,
2000.
[24] Hargreaves, J. K., The Solar-Terrestrial Environment, Cambridge Univ.
Press, Cambridge, England, U.K., 1992.
[25] DeForest, S. E., and McIlwain, C. E., “Plasma Clouds in the
Magnetosphere,” Journal of Geophysical Research, Vol. 76, No. 16,
1971, pp. 3587–3611.
doi:10.1029/JA076i016p03587
[26] Bilitza, D., “International Reference Ionosphere 2000,” Radio Science,
Vol. 36, No. 2, 2001, pp. 261–275.
doi:10.1029/2000RS002432
[27] Schroder, D. K., Semiconductor Material and Device Characterization,
3rd ed., Wiley, Hoboken, NJ, 2006.
[28] Purvis, C. K., Garrett, H. B., Whittlesey, A. C., and Stevens, N. J.,
“Design Guidelines for Assessing and Controlling Spacecraft Charging
Effects,” NASA TP-2361, 1984.
[29] Ferguson, D. C., and Hillard, G. B., “Low Earth Orbit Spacecraft
Charging Design Guidelines,” NASA TP-212287, 2003.
[30] Bever, R. S., and Staskus, J., “Tank Testing of a 2500 cm2 Solar Panel:
Spacecraft Charging Technology,” NASA CP-2182, 1981.
[31] Guillaumon, J. C., and Paillous, A., “Spacecraft Materials: Comparison
Between Flight Results Obtained on LDEF and MIR,” Proceedings of
LDEF Materials Results for Spacecraft Applications, NASA CP-3257,
1994, pp. 485–498.
[32] Visentine, J. T., Legar, L. J., Kuminecz, J. F., and Spiker, I. K., “STS-8
Atomic Oxygen Effects Experiment,” 23rd Aerospace Sciences
Meeting, AIAA Paper 1985-0415, Jan. 1985.
[33] Banks, B. A., Mirtich, M. J., Rutledge, S. K., and Swec, D. M.,
“Sputtered Coatings for Protection of Spacecraft Polymers,” NASATM83706, 1984.
ANGIRASA AND AYYASWAMY
Downloaded by GA Inst of Technology on May 9, 2023 | http://arc.aiaa.org | DOI: 10.2514/1.A32742
[34] Banks, B. A., Mirtich, M. J., Rutledge, S. K., and Nahra, H. K.,
“Protection of Solar Array Blankets from Attack by Low Earth Orbit
Atomic Oxygen,” Proceedings of the 18th IEEE Photovoltaic
Specialists Conference, IEEE Press, Piscataway, NJ, 1985, pp. 381–
386.
[35] Knopf, P. W., Martin, R. J., Damman, R. E., and McCargo, M.,
“Correlations of Laboratory and Flight Data for the Effects of Atomic
Oxygen on Polymeric Materials,” AIAA 20th Thermophysics
Conference, AIAA Paper 1985-1066, June 1985.
[36] Banks, B. A., Mirtich, M. J., Rutledge, S. K., Swec, D. M., and Nahra, H.
K., “Ion Beam Sputter-Deposited Thin Film Coatings for the Protection
of Spacecraft Polymers in Low Earth Orbit,” NASA TM-87051, 1985.
[37] Rutledge, S. K., and Olle, R. M., “Durability Evaluation of Photovoltaic
Blanket Materials Exposed on LDEF Tray (S10003): Part 3,”
Proceedings of LDEF — 69 Months in Space: First Post-Retrieval
Symposium, edited by Levine, A. S., NASA CP-3134, 1991, pp. 1379–
1394.
761
[38] Gilmore, D. G., Stuckey, W. K., and Fong, M., “Thermal Surface
Finishes,” Spacecraft Thermal Control Handbook Vol. 1: Fundamental
Technologies, 2nd ed., edited by Gilmore, D. G., Aerospace Press, El
Segundo, CA, 2002, Chap. 4.
[39] Gilmore, D. G., “Radiators,” Spacecraft Thermal Control Handbook
Vol. 1: Fundamental Technologies, 2nd ed., edited by Gilmore, D. G.,
Aerospace Press, El Segundo, CA, 2002, Chap. 6.
[40] Wilkes, D. R., Brown, M. J., Hummer, L. L., and Zwiener, J. M., “Initial
Materials Evaluation of the Thermal Control Surfaces Experiment
(S0069): Part 2,” Proceedings of LDEF — 69 Months in Space: First
Post-Retrieval Symposium, edited by Levine, A. S., NASA CP-3134,
1991, pp. 899–917.
T. Minton
Associate Editor
C H A P T E R
28
Degradation of Spacecraft Materials
Kim K. de Groh, Bruce A. Banks, Sharon K.R. Miller and Joyce A. Dever
NASA Glenn Research Center, Cleveland, OH, United States
O U T L I N E
28.1 Introduction
601
28.2 Atomic Oxygen Effects
28.2.1 Environment Description
28.2.2 Interaction With Materials
28.2.3 Mitigation Techniques
602
602
606
618
28.3 Contamination Effects
28.3.1 Sources and Transport of Spacecraft
Contamination
28.3.2 Contamination Effects on Spacecraft
Surfaces
28.3.3 Mitigation of Contamination Effects
621
28.4 Space Radiation Effects
28.4.1 Solar UV Radiation
28.4.2 Ionizing Radiation
623
623
627
621
621
622
28.4.3 Mitigation of Radiation Degradation
631
28.5 Thermal and Thermal Cycling Effects
28.5.1 Environment Description
28.5.2 Effects on Materials
631
631
632
28.6 Micrometeoroid and Orbital Debris Effects
28.6.1 Environment Description
28.6.2 Interactions With Materials
28.6.3 Mitigation Techniques
638
638
639
639
28.7 Concluding Remarks
641
References
642
Further Reading
645
28.1 INTRODUCTION
Materials used on exterior spacecraft surfaces are subjected to many environmental threats that can cause degradation. These threats include photon radiation, charged-particle radiation, temperature effects and thermal
cycling, impacts from micrometeoroids and debris, contamination, and low Earth orbit atomic oxygen. Space
environmental threats to materials vary greatly, based on both the material and its environment. Environmental
variables include orbital parameters for the mission, mission duration, 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 low Earth orbit. It is clear that each mission has its own unique set of environmental exposure
conditions that must be well understood for purposes of selecting durable spacecraft materials and interpreting
observed degradation.
Fundamentally important properties for exterior spacecraft surfaces include structural integrity and thermooptical properties. Problems occur when spacecraft materials become too thin or brittle to support a required
load, or when protective thermal insulation film layers crack and peel away from the spacecraft. Operating
temperatures of spacecraft systems rely on exterior surfaces possessing the required solar absorptance and
Handbook of Environmental Degradation of Materials.
DOI: https://doi.org/10.1016/B978-0-323-52472-8.00029-0
601
© 2018 Elsevier Inc. All rights reserved.
602
28. DEGRADATION OF SPACECRAFT MATERIALS
thermal emittance values. Solar absorptance, αS, is the fraction of incident solar energy that is absorbed by a surface. Thermal emittance, ε, is the ratio of radiated energy emitted from a surface to that which would be emitted
from a perfect (black body) emitting surface. Degradation of these thermo-optical properties can cause an undesirable change in temperature of the spacecraft or its components. Loss of transmittance through solar cell coverglass materials, such as by contamination, can result in decreased output of solar arrays and thus a reduction in
overall spacecraft power. For electrical wiring and cables exterior to the spacecraft, polymer insulation can
become degraded upon space radiation exposure.
Understanding the degradation of spacecraft materials can be determined through actual space exposure and
ground laboratory studies. Each has advantages and disadvantages. Opportunities for examining space-flown
materials, through retrieved flight hardware or dedicated experiments, are rare. Dedicated space-flight experiments are expensive and require long lead times from planning to flight. Differences between the experiment
environment, the intended mission environment, and synergistic environmental effects require cautious interpretation of results. When it is not possible to retrieve spacecraft hardware, data available from satellite operations,
such as power output and spacecraft surface temperatures, can also be used to assess material performance to
some extent. Where severe degradation is evident, however, it is often not possible to conclusively identify the
cause or mechanism of degradation, since comprehensive analysis of materials is not possible. Ground laboratory
studies can be used to examine individual environmental effects or a combination of environmental effects.
Laboratory tests can be conducted in a timely manner using accelerated levels for some environmental effects,
but, due to the difficulties in exactly simulating the space effects, complex calibrations and cautious interpretation
of the results are required. A combination of space exposures, ground laboratory studies, and computational
modeling is most useful for assuring durability of spacecraft materials.
This chapter includes descriptions of specific space environmental threats to exterior spacecraft materials. The
scope will be confined to effects on exterior spacecraft surfaces and will not address environmental effects on
interior spacecraft systems, such as electronics. Space exposure studies and laboratory simulations of individual
and combined space environmental threats will be summarized. Significant emphasis is placed on the effects of
Earth orbit environments, because the majority of space missions have been flown in Earth orbits that have provided a significant amount of data on materials effects. Issues associated with interpreting materials degradation
results will be discussed, and deficiencies of ground testing will be identified. Recommendations are provided on
reducing or preventing space environmental degradation through appropriate materials selection.
28.2 ATOMIC OXYGEN EFFECTS
28.2.1 Environment Description
Atomic oxygen is formed in the low Earth orbital (LEO) environment by photo-dissociation of diatomic
oxygen. Short wavelength (,243 nm) solar radiation has sufficient energy to break the 5.12 eV O2 diatomic bond
(Dickerson et al., 1979) in an environment where the mean-free path is sufficiently long (B108 meters) that the
probability of reassociation or the formation of ozone (O3) is small. As a consequence, between the altitudes of
180 2 650 km, atomic oxygen is the most abundant species (Fig. 28.1) (N.O.A.A., N.A.S.A. and U.S.A.F, 1976).
Although excited states of atomic oxygen can be formed, their lifetimes are sufficiently short (Banks et al., 1991)
that the P ground state dominates the LEO atomic oxygen formation and is dependent on the diatomic oxygen
density and solar UV flux. Solar heating of the Earth’s atmosphere causes an increase in the number density of
atoms at a given altitude as the Earth rotates from sunrise toward solar noon. Because the atmosphere
corotates with the Earth, the solar-heated bulge in the atmosphere is pushed forward such that the peak of the
atomic oxygen density occurs at approximately 3 P.M. rather than solar noon. As a consequence, antisolar facing
surfaces, such as the back side of solar arrays, receive 25% more atomic oxygen fluence than the solar-facing surfaces as the spacecraft orbits the Earth (Banks et al., 1991).
Solar-caused variations in the UV radiation impinging upon the LEO atmosphere can greatly change the
atomic oxygen production rate (and, therefore, the arriving flux on spacecraft surfaces). Periods of high and low
solar activity can change the arriving flux by a factor of up to 500, depending on altitude (Fig. 28.2) (N.O.A.A., N.
A.S.A. and U.S.A.F, 1976). Thus, the atomic oxygen flux cannot be accurately predicted due to uncertainty in the
solar activity. The average atomic oxygen fluence per year varies as a result of the solar activity consistent with
the 11-year sun spot cycle, as shown in Fig. 28.3, for which data were calculated using the MSIS-86 atmospheric
IV. CASE STUDIES
603
28.2 ATOMIC OXYGEN EFFECTS
Altitude, km
700
600
O
N2
O2
500
Ar
He
H
400
300
200
100
1e+4
FIGURE 28.1
1e+5
1e+9
1e+7
1e+8
Number density, cm−3
1e+6
1e+10
1e+11
1e+12
Density of atmospheric species as a function of altitude.
1000
Solar minimum
Standard atmosphere
Solar maximum
Altitude, km
800
600
400
200
1e+4
FIGURE 28.2
1e+5
1e+6
1e+7
1e+8
1e+9
Number density, cm−3
1e+10
1e+11
Atomic oxygen flux versus altitude for solar minimum, nominal (standard atmosphere), and solar maximum conditions.
model (Hedin, 1987). Atomic oxygen can also be produced in other planetary environments where oxygen is
present, such as the low Mars orbit.
As a spacecraft orbits the Earth at velocities of the order of 7.7 km/sec, it runs into the atomic oxygen (hence
the term “ram” atomic oxygen). If the spacecraft is in an orbit that has zero inclination then the average angle of
attack of the atomic oxygen is perpendicular to surfaces whose surface normal points in the direction of travel.
However, most spacecraft have orbits that are inclined with respect to the Earth’s equatorial plane. This causes
the average angle of attack of the arriving atomic oxygen to sinusoidally vary around the orbit as a result of the
vectoral addition of the orbital spacecraft velocity vector and the atmosphere’s corotation velocity vector (de
Groh and Banks, 1994). In addition, atomic oxygen atoms have thermal velocities associated with their
MaxwellBoltzman velocity distribution at the high temperatures of LEO, which are typically B1,000 K (N.O.A.
A., N.A.S.A. and U.S.A.F, 1976). The high-velocity tail of the MaxwellBoltzman distribution actually allows
some atomic oxygen atoms to catch up with the trailing, or wake, surfaces of a LEO spacecraft to produce a small
flux, which is orders of magnitude lower than the ram flux. Thus, the thermal velocities of the atomic
oxygen associated with their MaxwellBoltzman velocity distribution contributes an additional component to
IV. CASE STUDIES
604
28. DEGRADATION OF SPACECRAFT MATERIALS
Total atomic oxygen ram fluence
for one year (atoms/cm2)
the overall impact velocity of the atomic oxygen. If we add the three vectoral components and average over a
typical 400-km orbit at 28.5 degrees inclination, we come to the angular distribution of arriving atoms shown in
Fig. 28.4, where the arrival distribution in the horizontal plane is shown as a function of incidence angle for
surfaces normal to the ram direction (Banks et al., 1999).
Atomic oxygen can arrive at angles beyond 90 degrees from the orbital direction. For example, Fig. 28.5 shows
that a surface whose normal is 90 degrees with respect to the ram direction receives approximately 4% of the flux
that occurs for a surface whose normal is parallel to the ram direction.
The impact energy of arriving atomic oxygen atoms also is dependent on the following three contributions to
the resulting velocity vectors: the orbital spacecraft velocity, the Earth’s atmosphere corotation velocity, and
atomic oxygen thermal velocity. Fig. 28.6 is a plot of the energy distribution of atomic oxygen atoms as a function
of altitude for a circular orbit with 28.5 degrees inclination and 1,000 K thermosphere (Banks et al., 1990) As can
be seen, the average impact energy is 4.5 eV 6 1 eV for a 400-km orbit and the impact energy decreases with
3.5E+21
3.0E+21
2.5E+21
2.0E+21
1.5E+21
1.0E+21
5.0E+20
0.0E+00
2014
2016
2018
2020
2022
2024
2026
2028
Year at 400 km circular orbit, 28.5 degree inclination
FIGURE 28.3 Atomic oxygen fluence per year during a solar cycle, based on MSIS-86 atmospheric model (Hedin, 1987).
FIGURE 28.4 Atomic oxygen arrival flux relative to the ram direction for a 400-km orbit at 28.5 degrees inclination and 1000K
thermosphere.
IV. CASE STUDIES
605
28.2 ATOMIC OXYGEN EFFECTS
FIGURE 28.5 Polar plot of relative atomic oxygen flux as a function of the angle between the ram direction and the normal of the arrival
surface for a LEO spacecraft in a 400-km orbit at 28.5 degrees inclination and 1000K thermosphere.
500
450
Altitude (km)
400
−σ
350
+σ
300
250
200
3.0
3.5
4.0
4.5
Energy, eV
5.0
5.5
6.0
FIGURE 28.6 Energy distribution of atomic oxygen atoms as a function of altitude for a circular orbit at 28.5 degrees inclination and
1000K thermosphere.
altitude. For highly elliptical orbits, the perigee ram impact energy can be significantly higher than for circular
LEO orbits. Such elliptical orbits can also produce high fluxes near perigee due to the low altitudes involved. If a
spacecraft is spinning with its axis of rotation per-pendicular to the Earth then the average flux to any surface is
simply 1/π of that of the ram direction.
IV. CASE STUDIES
606
28. DEGRADATION OF SPACECRAFT MATERIALS
28.2.2 Interaction With Materials
Although LEO atomic oxygen possesses sufficient energy to break most organic polymer bonds and sufficient
flux to cause oxidative erosion of polymers, there was little known or interest in atomic oxygen interaction with
materials until the start of Space Shuttle missions. This is primarily because most prior missions occupied highaltitude orbits where atomic oxygen densities are rather inconsequential.
One example of LEO environmental interaction with materials is the glow phenomena that occurs when
atomic oxygen and other LEO atmospheric species impact spacecraft surfaces causing the creation of short-lived
excited state species that emit visible radiation near the surfaces of spacecraft as shown in Fig. 28.7. Fig. 28.7A
was taken during daylight and Fig. 28.7B was taken at night (Mende et al., 1984; Rice and Russell, 1992;
Caledonia, 1989).
The reaction of atomic oxygen with spacecraft materials has been a significant problem to LEO spacecraft
designers. Atomic oxygen can react with polymers, carbon, and many metals to form oxygen bonds with atoms
on the surface that are exposed. For most polymers, hydrogen abstraction, oxygen addition, or oxygen insertion
can occur (Fig. 28.8). With continued atomic oxygen exposure, all oxygen interaction pathways for hydrocarbons
eventually lead to volatile oxidation products accompanied by the gradual erosion of hydrocarbon materials.
The surfaces of polymers exposed to atomic oxygen also develop an increase in oxygen content as shown in
Fig. 28.9 (Mata et al., 2003).
The sensitivity of hydrocarbon materials to reaction with atomic oxygen is quantified by the atomic oxygen
erosion yield of the material. The atomic oxygen erosion yield is the volume of a material that is removed
(through oxidation) per incident atom impacting the material. The most well-characterized atomic oxygen erosion
yield is that of polyimide Kapton H, which has an erosion yield of 3.0 3 10224 cm3/atom for LEO 4.5 eV atomic
oxygen (Banks and Rutledge, 1988; Leger, 1983).
FIGURE 28.7 Low Earth orbital glow phenomena: (A) Photo of the Space Shuttle tail during daylight; (B) photo of the Space Shuttle tail at night.
IV. CASE STUDIES
607
28.2 ATOMIC OXYGEN EFFECTS
RCH2CH3
N
2 RCH2CH2− + HOH
IO
RCH2CH2− + OH−
T
AC
R
ST
AB
VOLATILES
REPLACEMENT
RECOMBINATION/
RCH2CH3 + O
R’O− + R’’CH2CH3−
(CO, H2O, HCO,
IN
FURTHER
REACTION/
SE
CH3, ETC.)
RT
FRAGMENTATION
IO
N
+
[RCH2CH2OH]
FRAGMENTATION (VOLATILES)
ALKANES
RECOMBINATION/
FURTHER REACTION/
FRAGMENTATION
ALKENES
H
R
C
N
TIO
H
R
C
H
+
C
AC
TR
INSERTION
S
AB
R
O A
DD
H
H
C
C
−
O
−
FIGURE 28.8
H
C
FRAGMENTATION (VOLATILES)
OH
*
H
+ HOH
C−
*
ON
H
C
2
H
C
R
H
R
CHRCH2
H
ITI
H
+ OH−
C−
VOLATILES
ION
INAT
ELIM
RCCH2 −
+ H − (VOLATILES)
O
H
R
C
EPOXIDE FORMATION
H
TR
I
INT PLET
RCCH3
E
RE RCO SING
AR
L
N
O
E
RA VE
T
NG RS
EM ION
/
EN
T
*
C
O
H
FRAGMENTATION
(VOLATILES)
*
FRAGMENTATION
(VOLATILES)
Atomic oxygen reaction pathways with polymers.
Surface oxygen content (%)
12
10
8
6
4
2
0
Unexposed
1e+17
1e+18
1e+19
1e+20
1e+21
Atomic oxygen effective fluence (atoms/cm2)
FIGURE 28.9 Surface oxygen content (measured using XPS) for chlorotrifluoroethylene as a function of atomic oxygen effective fluence,
where exposures were conducted in ground facility RF air plasma.
IV. CASE STUDIES
608
28. DEGRADATION OF SPACECRAFT MATERIALS
The most common technique for determining the erosion yield of flight samples is through mass-loss measurements. These measurements are made by obtaining mass measurements of the sample before and after flight.
The erosion yield of the sample, ES, is calculated through the following equation:
ES 5
ΔMS
;
ðAS ρS FÞ
ð28:1Þ
where
ES 5 erosion yield of flight sample (cm3/atom)
ΔMS 5 mass loss of flight sample (g)
AS 5 surface area of flight sample exposed to atomic oxygen attack (cm2)
ρS 5 density of sample (g/cm3)
F
5 fluence of atomic oxygen (atoms/cm2)
The atomic oxygen fluence, F, can be determined through the mass loss of a Kapton witness sample because
Kapton has a well-characterized erosion yield in the LEO environment. Therefore, the atomic oxygen fluence can
be calculated using:
F5
ΔMK
;
ðAK ρK EK Þ
ð28:2Þ
where
ΔMK 5 mass loss of Kapton witness sample (g)
AK 5 surface area of Kapton witness sample exposed to atomic oxygen (cm2)
ρK 5 density of Kapton witness sample (1.42 g/cm3)
EK 5 erosion yield of Kapton witness sample (3.0 3 10224 cm3/atom)
Thus:
E S 5 EK
ΔMS AK ρK
:
ΔMK AS ρS
ð28:3Þ
One of the critical issues with obtaining accurate erosion yield data is related to the humidity of the environment where the mass measurements are taken. Many polymer materials, such as Kapton, are very hygroscopic
(absorbing up to 2% of their weight in moisture) and can fluctuate in mass significantly with humidity and temperature. Therefore, for accurate mass loss measurements to be obtained, it is necessary that the samples be fully
dehydrated (e.g., in a vacuum desiccator) prior to measuring the mass, both preflight and postflight.
Materials spaceflight experiments for atomic oxygen erosion yield determination have been flown on the
Shuttle, the Long Duration Exposure Facility (LDEF), the Russian space station MIR, and other spacecraft.(de
Groh, 2010) Erosion yield values obtained from Shuttle flight experiments were found to have large variations.
This is partly because Shuttle flight experiments were exposed to low atomic oxygen fluences on-orbit. However,
variations in much of the early LEO space data also occurred because some erosion yield data were not determined based on dehydrated mass measurements, introducing large error for hygroscopic materials, especially for
low-fluence exposures or low-erosion yield samples. More recently, long-duration experiments have been flown
on the exterior of the International Space Station (ISS) as a part of the Materials International Space Station
Experiment (MISSE) missions. In the original MISSE missions, individual flight experiments were flown in
suitcase-like containers called passive experiment containers (PECs). The PECs were closed during launch in the
Shuttle cargo bay to protect the samples. Once on orbit, the PECs were placed on the exterior of the ISS during
an extravehicular activity (EVA), or spacewalk, in either a ram/wake or a zenith/nadir orientation and opened,
exposing the experiments to the space environment. The PECs were left in place for 14 years and then closed
up and retrieved during another EVA and returned to Earth for postflight analyses. Table 28.1 provides a summary of the MISSE 1-8 missions.
Atomic oxygen erosion yield values from a wide variety of polymers flown in the ram direction as part of the
MISSE 2, 4, 6, 7, and 8 missions are provided in Table 28.2. The erosion yield data was obtain using preflight and
postflight dehydrated mass measurements. Variation in erosion yields for polymers obtained during different
MISSE missions is likely due to the variation in the atomic oxygen fluence and/or the variation in solar exposure
for the various missions. For example, the erosion yield of fluoropolymers such as Teflon fluorinated ethylene
IV. CASE STUDIES
TABLE 28.1 MISSE 1-8 Mission Exposure Summary
MISSE PEC
Launch
mission
Date placed outside
ISS
Location on ISS
Tray orientation
Retrieval
mission
Retrieved
from ISS
Exposure
duration (Years)
1&2
STS-105
8/16/2001
PEC 1: High Pressure Gas Tank (HPGT)
PEC 2: Quest Airlock
Ram & Wake
STS-114
7/30/2005
3.95
3&4
STS-121
8/3/2006a
PEC 3: HPGT
PEC 4: Quest Airlock
Ram & Wake
STS-118
8/18/2007
1.04
5
STS-114
8/3/2005
Aft P6 Trunion Pin Handrail
Zenith & Nadir
STS-115
9/15/2006
1.12
6 A & 6B
STS-123
3/22/2008
Columbus Laboratory
Ram & Wake
STS-128
9/1/2009
1.45
7A & 7B
STS-129
11/23/2009
EXPRESS Logistics Carrier 2 (ELC 2) on
the S3 Truss
7A: Zenith & Nadir 7B: Ram &
Wake
STS-134
5/20/2011
1.49
8 & ORMatEIII R/W
STS-134
**STS-135
8: 5/20/2011
ORMatE-III R/W:
7/12/2011**
EXPRESS Logistics Carrier 2 (ELC 2) on
the S3 Truss
8: Zenith & Nadir ORMatE-III R/
W: Ram & Wake
SpaceX-3
Dragon
7/9/2013
MISSE 8: 2.14
ORMatE-III: 2.00
a
Deployed during Expedition 13.
ORMatE-III R/W, Optical Reflector Materials Experiment III Ram/Wake.
TABLE 28.2 Atomic Oxygen Erosion Yields from MISSE 2-8 Ram Polymer Experiments
MISSE 2 ram
MISSE 4 ram
MISSE 6B ram
(nonstressed)
MISSE 7B ram
MISSE 8 ram
AO F* 5 8.43E121 at/cm2
Solar 5 6,300 ESH Mission: 4
years Ref (de Groh et al.,
2008)
AO F 5 2.15E121
at/cm2
Av Solar 5 1,400
ESH Mission: 1
year Ref (Dever
et al., 2010)
AO F 5 1.97E121 at/cm2
Solar 5 2,600 ESH
Mission: 1.5 years Ref
(de Groh et al., 2013)
AO F 5 4.22E121 at/cm2
Solar 5 2,400 ESH
Mission: 1.5 years Ref
(de Groh et al., 2016)
AO F 5 4.62E121
at/cm2 Solar 3,200 ESH Mission:
2 years Ref (de
Groh et al., 2017)
Thickness
(mils)
MISSE ID (#
layers) Ref (de
Groh et al., 2014)
AO Ey
(cm3/at)
MISSE
ID (#
layers)
MISSE ID
(# layers)
AO Ey
(cm3/at)
MISSE ID
^^(# layers)
AO Ey
(cm3/at)
MISSE
ID (#
layers)
Acrylonitrile butadiene
ABS
styrene (Cycolac, Absylux)
5
2-E5-6 (3)
1.09E-24
Allyl diglycol carbonate
(CR-39, Homalite H-911)
ADC
31, 47
2-E5-14 (1)
. 6.80E24a
W3-6/ N14S (1)
8.28E-24
Amorphous
fluoropolymer (Teflon AF
1601)
AF
2
2-E5-45 (1)
1.98E-25
Cellulose acetate (Clarifoil, CA
Tenite Acetate; Dexel)
2
2-E5-7 (13)
5.05E-24
Cellulose nitrate
(Celluloid)
CN
5
N5-14 (7)
. 6.58E-24
Chlorotrifluoroethylene
(Kel-F, Neoflon M-300)
CTFE
5
Carbon Diamond
(single crystal natural
Class IIA (100 plane))
Diamond
10
B7-6 (1)
1.58 6 0.40
E-26 (θ)
Carbon - Highly Oriented
Pyrolytic Graphite
(HOPG), Basal Plane
HOPG-BP 58
B7-6
1.05 6 0.08
E-24 (θ)
Carbon - Highly Oriented HOPG-EP 58
Pyrolytic Graphite
(HOPG), Edge (a, b) Plane
B7-6
5.38 6 0.90
E-25 (θ)
1.07E-25
Material (Trade Name)
Abbrev.
2-E5-39 (1)
2-E5-25 (1)
AO Ey
(cm3/at)
8.31E-25
Carbon - Pyrolytic
graphite (C plane)
PG
80
4.15E-25
Copolymer of
polytetrafluoro-ethylene,
perfluoropropyl vinyl
ether, perfluoromethyl
vinyl ether (Hyflon MFA
620)
MFA
4
M7BT-E15$
(1)
Crystalline
polyvinylfluoride (PVF)
PVF-W
2
B7-1 (1): AO 1.48E-25
F 5 4.22E 1 21
AO Ey
(cm3/at)
w/ white pigment (White
Tedlar)
Crystalline PVF w/ white
pigment (White Tedlar
with 0.5 mil Kapton H
cover)
PVF-W
2
B7-2 (1): AO 1.54E-25
F 5 3.79E 1 21
Crystalline PVF w/ white PVF-W
pigment (White Tedlar
with two 0.5 mil Kapton H
covers (1.0 mil))
2
B7-3 (1): AO 1.67E-25
F 5 3.37E 1 21
Crystalline PVF with
white pigment (White
Tedlar TW10B53)
PVF-W
1
Cyanate ester (CE)
composite (120FG
/EX1515 fiber glass/CE)
CE
39
M7BT-E19
(1) $
2.41E-25
Cyanate ester composite
(M55J /954-3 carbon
fiber/CE)
CE
35
M7BT-E20
(1) $
5.38E-25
Cyanate ester composite
(M55J /996 carbon fiber/
CE with 10% silicone)
CE Si
35
M7BT-E21
(1) $
3.44E-25
Epoxide or epoxy (Hysol
EA 956)
EP
88-92
2-E5-19 (1)
4.21E-24
Ethylenechlorotrifluoroethylene
(Halar 300)
ECTFE
3
2-E5-40 (3)
1.79E-24
EthyleneETFE
tetrafluoroethylene (Tefzel
ZM)
3
2-E5-41 (2)
9.61E-25
Fluorinated ethylene
propylene (Teflon FEP)
FEP
2
2-E5-42 (1)
2.00E-25
FEP (Teflon FEP)
FEP
5
FEP back-surface
aluminized (FEP/Al))
(Aluminized-Teflon FEP)
FEP/Al
5
FEP back-surface carbon
painted
FEP/C
2
FEP back-surface
aluminized, preflight
heated 380 hrs @200 C
Heated
FEP/Al
5
2-E5-11 (13)
2-E6-13 (1)b
1.01E-25
2.11E-25
W2-10/ N2
(4)
1.69E-25
W6-6 (1) %
2.28E-25
W6-11 (1) %
N5-4 (1)
1.81E-25
M7BT-E16
(1) $
1.57E-25
M8-R5
(7)
1.45E-25
M9-R9
(1)
2.37E-25
M8-R2
(1)
2.39E-25
2.26E-25
(Continued)
TABLE 28.2 (Continued)
Thickness
(mils)
MISSE 2 ram
MISSE 4 ram
MISSE 6B ram
(nonstressed)
MISSE 7B ram
MISSE 8 ram
AO F* 5 8.43E121 at/cm2
Solar 5 6,300 ESH Mission: 4
years Ref (de Groh et al.,
2008)
AO F 5 2.15E121
at/cm2
Av Solar 5 1,400
ESH Mission: 1
year Ref (Dever
et al., 2010)
AO F 5 1.97E121 at/cm2
Solar 5 2,600 ESH
Mission: 1.5 years Ref
(de Groh et al., 2013)
AO F 5 4.22E121 at/cm2
Solar 5 2,400 ESH
Mission: 1.5 years Ref
(de Groh et al., 2016)
AO F 5 4.62E121
at/cm2 Solar 3,200 ESH Mission:
2 years Ref (de
Groh et al., 2017)
MISSE ID (#
layers) Ref (de
Groh et al., 2014)
MISSE
ID (#
layers)
MISSE ID
(# layers)
AO Ey
(cm3/at)
MISSE ID
^^(# layers)
MISSE
ID (#
layers)
AO Ey
(cm3/at)
AO Ey
(cm3/at)
Material (Trade Name)
Abbrev.
FEP/Ag/Inconel, flown
on LDEF
LDEF Ag-FEP F-04 (10,458
ESH)
Hubble Space Telescope
(HST) FEP/Al 1st
servicing mission (SM1
MSS-D: 3.6 yrs, 11,339
ESH)
LDEF Ag- 5
FEP
W6-13 (1) %
2.37E-25
HST
FEP/Al
5
W6-8 (1) %
2.40E-25
HST FEP/Al 3rd servicing HST
mission (SM3A R1: 9.7 yrs, FEP/Al
13,598 ESH)
5
W6-7 (1) %
2.11E-25
HST FEP/Al 3rd servicing HST
mission (SM3A R5: 9.7 yrs, FEP/Al
.30,321 ESH)
5
W6-12 (1) %
2.34E-25
FEP coated (TiO2/Al2O3/
FEP)
TiO2/
Al2O3/
FEP
2
FEP back-surface vapor
deposited silver/niobium
(FEP/VDS/Nb)
FEP/
VDS/Nb
5
2-E22-29 1.28E-25
(1)
Germanium coated (25 Å)
Kapton XC with Nomex
scrim
Ge/
1
Kapton
XC/Scrim
2-E22-16 1.65E-26
(1)
High-temperature
polyimide resin (PMR-15)
PI (PMR15)
13
High-temperature
polyimide resin (PMR-15)
PI (PMR15)
34
Perfluoroalkoxy
copolymer resin (Teflon
PFA 200 CLP)
PFA
2
M8-R6
(1)
B7-11 (1)
2-E5-34 (1)
. 3.02E24a
W2-7/ N12
(1)
2-E5-44 (4)
1.73E-25
2.68E-24^
AO Ey
(cm3/at)
7.99E-28
AO Ey
(cm3/at)
2.50E-25
Poly-(p-phenylene
terephthalamide) (Kevlar
29 fabric)
PPD-T
2.2
2-E5-8 (3)
6.28E-25
Poly(p-phenylene-2,6benzobis-oxazole) (Zylon,
balanced biaxial)
PBO
1
2-E5-18 (11)
1.36E-24
Polyacrylonitrile (Barex
210)
PAN
2
2-E5-13 (9)
1.41E-24
Polyamide 6 (Nylon 6)
PA 6
2
2-E5-27 (8)
3.51E-24
Polyamide 66 (Nylon 66)
PA 66
2
2-E5-28 (7)
1.80E-24
Polyamide-imide (Torlon
4203)
PAI
29
Polybenzimidazole
(Celazole PBI 22)
PBI
2
Polybenzimidazole
PBI
4
Polybutylene terephthalate PBT
(VALOX 357)
1
2-E5-21 (5)
9.11E-25
Polycarbonate (PEEREX
61)
PC
10
2-E5-36 (2)
4.29E-24
Polyetheretherketone
(Victrex PEEK 450)
PEEK
3
2-E5-37 (6)
2.99E-24
Polyetherimide (Ultem
1000)
PEI
10
2-E5-26 (2)
. 3.31E24a
Polyethersulfone
PES
3
Polyethylene
PE
2
2-E5-9 (6)
. 3.74E24a
Polyethylene oxide (Alkox PEO
E-30 powder)
29
2-E5-17 (1)
1.93E-24
Polyethylene terephthalate PET
(Mylar A-200)
2
2-E5-38 (8)
3.01E-24
Polyethylene terephthalate PET
(Mylar A-500)
5
Polyimide, clear (LaRC
CP1-300)
CP1
3
2-E5-29 (4)
1.91E-24
Polyimide, clear (CP1, SRS CP1
Tech.)
1
2-E6-15 (9)b
1.91E-24 2-E22-26 2.26E-24
(20)
Polyimide CORIN (AO
resistant)
2
CORIN
2-E5-35 (4)
N5-12 (1)
1.74E-24
N5-3 (6)
2.79E-24
N5-1 (18)
4.11E-24
M7BT-E17
(1) $
2.75E-24
N5-15 (4)
3.05E-26
. 2.21E24a
W3-7/N12-S 4.55E-24
(5)
W2-12/N13
(2)
W2-2/N10
(10)
3.37E-24
4.05E-24
W2-13/N9
(4)
3.22E-24
W2-4/N4
(20)
2.16E-24
(Continued)
TABLE 28.2 (Continued)
Thickness
(mils)
MISSE 2 ram
MISSE 4 ram
MISSE 6B ram
(nonstressed)
MISSE 7B ram
MISSE 8 ram
AO F* 5 8.43E121 at/cm2
Solar 5 6,300 ESH Mission: 4
years Ref (de Groh et al.,
2008)
AO F 5 2.15E121
at/cm2
Av Solar 5 1,400
ESH Mission: 1
year Ref (Dever
et al., 2010)
AO F 5 1.97E121 at/cm2
Solar 5 2,600 ESH
Mission: 1.5 years Ref
(de Groh et al., 2013)
AO F 5 4.22E121 at/cm2
Solar 5 2,400 ESH
Mission: 1.5 years Ref
(de Groh et al., 2016)
AO F 5 4.62E121
at/cm2 Solar 3,200 ESH Mission:
2 years Ref (de
Groh et al., 2017)
MISSE ID (#
layers) Ref (de
Groh et al., 2014)
MISSE
ID (#
layers)
AO Ey
(cm3/at)
MISSE ID
(# layers)
AO Ey
(cm3/at)
MISSE ID
^^(# layers)
AO Ey
(cm3/at)
MISSE
ID (#
layers)
W2-15/N3
(20)
1.65E-24
B7-8 (4)
3.00E-24c
N5-5 (4)
3.05E-24
N5-6 (4):
AO F 5 4.22
E 1 21
3.01E-24
N5-7 (4):
2.85E-24
AO Ey
(cm3/at)
Material (Trade Name)
Abbrev.
Polyimide (BPDA)
(Upilex-S)
PI
1
(Upilex-S)
2-E5-32 (11)
9.22E-25 2-E22-19 1.71E-24
(21)
Polyimide (BPDA)
(Upilex-S)
PI
1
(Upilex-S)
2-E6-14 (9)b
9.76E-25
Polyimide PMDA (Kapton PI
CB)
(Kapton
CB)
5
W2-9/N5 (4) 2.70E-24
Polyimide (PMDA)
(Kapton E)
PI
(Kapton
E)
2
W2-8/N7
(10)
2.83E-24
Polyimide (PMDA)
(Kapton H)
PI
(Kapton
H)
5
W2-3/N8
(4), W6-5/
GW-1 (3)
3.00E-24c
Polyimide (PMDA), 0.75v
square (Kapton H)
PI
(Kapton
H)
5
Polyimide (PMDA), 10
layers 0.5 mil Kapton H
with spacers over 2 x
5 mil layers Kapton H
PI
(Kapton
H)
0.5
W6-9/ GW-3 3.25E-24
(10)
Polyimide (PMDA), 10
PI
layers 0.5 mil Kapton H
(Kapton
(no spacers) over 1 x 5 mil H)
Kapton H
0.5
W6-10/ GW- 3.08E-24
2 (10)
Polyimide (PMDA)
(Kapton HN)
PI
(Kapton
HN)
5
2-E5-31 (3)
2.81E-24 2-E22-17 2.83E-24
(7)
Polyimide (PMDA)
(Kapton HN)
PI
(Kapton
HN)
5
2-E6-16 (3)b
2.75E-24
5
2-E5-30 (3),
2-E5-33 (3)
3.00E24c
2-E22-18 3.00E-24c
(7)
M8-R1
(4)
AO Ey
(cm3/at)
3.00E-24c
Polyimide (PMDA),
Kapton HN with 0.5 mil
Kapton H cover
PI
(Kapton
HN)
AO
F 5 3.79E 1 21
Polyimide (PMDA),
Kapton HN with two
0.5 mil (1.0 mil) Kapton H
covers
PI
(Kapton
HN)
5
Polyimide (Kapton XC
(1E-7 Ohms/Sq))
PI
(Kapton
XC)
5
Polyimide (Vespel)
PI
(Vespel)
20
Polymethyl methacrylate
(Plexiglas; Acrylite
(Impact Mod.))
PMMA
2
Polymethylpentene (TPX
DX845 Natural)
PMP
2
Polyoxymethylene; acetal;
poly-formaldehyde
(Delrin Acetal)
POM
10
2-E5-12 (5)
9.14E-24
Polyphenylene
isophthalate (Nomex
Crepe Paper T-410)
PPPA
2
2-E5-24 (6)
1.41E-24
Polypropylene (Contour
28)
PP
20
2-E5-20 (1)
2.68E-24
Polystyrene (Trycite 1000)
PS
2
2-E5-15 (11)
3.74E-24
Polysulfone (Thermalux
P1700-NT11, Udel P-1700)
PSU
2
2-E5-22 (6)
2.94E-24
Polytetrafluoroethylene
(Chemfilm DF 100)
PTFE
2
2-E5-43 (1)
1.42E-25
Polytetrafluoroethylene
(Chemfilm DF 100)
PTFE
5
Polyurethane (Dureflex
PS8010)
PU
2
Polyvinyl Alcohol
(MonoSol M1000)
PVOH
1.5
N5-13 (24)
3.14E-24
Polyvinyl chloride (ClearLay Rigid PVC)
PVC
5
B7-10 (1)
.1.74E-24a
N5-8 (4):
2.20E-24
AO
F 5 3.37E 1 21
W2-14/N6
(4)
2-E5-16 (10)
. 5.60E24a
W3-11/
N15-S (10)
2.20E-24
B7-9 (1)
2.94E-24
N5-2 (9)
4.45E-24
M7BT-E14
(1) $
3.12E-24
5.28E-24
M8-R7
(1)
1.94E-25
W2-5/N1 (4) 1.33E-25
2-E5-23 (9)
1.56E-24
(Continued)
TABLE 28.2 (Continued)
MISSE 2 ram
MISSE 4 ram
MISSE 6B ram
(nonstressed)
MISSE 7B ram
MISSE 8 ram
AO F* 5 8.43E121 at/cm2
Solar 5 6,300 ESH Mission: 4
years Ref (de Groh et al.,
2008)
AO F 5 2.15E121
at/cm2
Av Solar 5 1,400
ESH Mission: 1
year Ref (Dever
et al., 2010)
AO F 5 1.97E121 at/cm2
Solar 5 2,600 ESH
Mission: 1.5 years Ref
(de Groh et al., 2013)
AO F 5 4.22E121 at/cm2
Solar 5 2,400 ESH
Mission: 1.5 years Ref
(de Groh et al., 2016)
AO F 5 4.62E121
at/cm2 Solar 3,200 ESH Mission:
2 years Ref (de
Groh et al., 2017)
MISSE ID (#
layers) Ref (de
Groh et al., 2014)
AO Ey
(cm3/at)
MISSE
ID (#
layers)
MISSE ID
(# layers)
MISSE ID
^^(# layers)
MISSE
ID (#
layers)
Material (Trade Name)
Abbrev.
Thickness
(mils)
Polyvinylfluoride (Tedlar
TTR10SG3)
PVF
1
2-E5-10 (13)
3.19E-24
Polyvinylidene fluoride
(Kynar 740)
PVDF
3
2-E5-46 (2)
1.29E-24
Silicone (DC 93-500 on
fused silica)
DC 93500
10
Triton Oxygen Resistance
polymer (TOR LM)
TOR
1.5
Urethane/Vectran mesh
Ur/
Vectra
10
a
AO Ey
(cm3/at)
AO Ey
(cm3/at)
M8-R10
(1)
2-E22-30 Mass
(1)
gain
Ey is greater than this value because the sample was fully or partially eroded through the sample.
AO fluence for this tray was determined to be 8.51E 1 21 atoms/cm2 (Ref (Dever et al., 2008)).
c
Kapton H was used to determine the mission AO fluence using the LEO Ey from prior flight experiments (3.00E-24 cm3/atom).
*AO F, Atomic oxygen fluence.
$ Taped in a thin Al holder.
^ Residue on the surface, may have impact Ey.
^^ MISSE 7 N5 sample were 0.75vsquare and the B7 samples were 1v square.
% Not previously published data (not included in Ref. de Groh et al., 2013).
θ Not previously published data (not included in Ref. de Groh, 2016).
b
AO Ey
(cm3/at)
N5-16 (1)
4.61E-25
AO Ey
(cm3/at)
3.81E-27
617
28.2 ATOMIC OXYGEN EFFECTS
TABLE 28.3
Erosion Yield (Ey) Comparison for Teflon FEP Flown in Different Orientations on MISSE 8 (de Groh et al., 2017)
MISSE 8
orientation
Solar
exposure
(ESH)
Atomic oxygen (AO)
fluence (atoms/cm2)
Solar/AO ratio (xE19
ESH*cm2/atom)
FEP Ey
(cm3/atom)
Al-FEP Ey
(cm3/atom)
HST Al-FEP Ey
(cm3/atom)
Ram
3200
4.62E 1 21
6.9
2.37E-25
2.39E-25
2.50E-25
Wake
2700
8.80E 1 19
307
1.08E-24
1.10E-24
1.04E-24
Zenith
(Taped)
6100
1.96E 1 20
312
1.37E-24
1.90E-24
Zenith (Tray) 6100
4.04E 1 19
1510
5.71E-24
6.32E-24
6.32E-24
7.E-24
y = 4E-08x + 2E-26
R2 = 0.9951
Atomic oxygen erosion yield (cm3/atom)
FEP
6.E-24
Al-FEP
HSTAl-FEP
5.E-24
y = 4E-08x + 2E-25
R2= 0.9861
Zenith
(Tray)
4.E-24
y = 4E-08x + 1E-25
R2 = 0.9974
3.E-24
Zenith
(Taped)
2.E-24
1.E-24
Ram
0.E+00
0.0E+00
Wake
2.0E-17
4.0E-17
6.0E-17
8.0E-17
1.0E-16
1.2E-16
1.4E-16
1.6E-16
2/atom)
Solar exposure/atomic oxygen fluence (ESH*cm
FIGURE 28.10 Erosion yield versus solar exposure to atomic oxygen fluence ratio for Teflon FEP samples flown in the ram, wake, and
zenith orientations on MISSE 8.
propylene (FEP) have been found to be directly related to the solar exposure with the erosion yield increasing
substantially with increasing solar exposure. Erosion yield values for Teflon FEP flown in ram, wake, and zenith
directions during the MISSE 8 mission are provided in Table 28.3 and graphed in Fig. 28.10.(de Groh et al., 2017)
The samples include clear Teflon FEP, back-surface aluminized-Teflon FEP (Al-FEP), and back-surface aluminized-Teflon FEP retrieved from the Hubble Space Telescope (HST) after 9.7 years of space exposure (HST-AlFEP). As can be seen, there is an excellent correlation of erosion yield with the solar exposure to AO fluence ratio
(for all three types of Teflon FEP). This clearly shows the effect of solar radiation, and/or heating due to solar
exposure, on FEP erosion.
Atomic oxygen can also oxidize the surfaces of metals to produce nonvolatile metal oxides. However, for most
metals, the oxides tend to shield the underlying metal from oxidation. Silver is one exception, because silver
oxide tends to spall from the underlying metal, thus allowing continued oxidation. Such effects caused silver
solar cell interconnects to fail in LEO (Leger, 1983). Atomic oxygen interaction with silicones causes oxidation
and removal of methyl groups and gradual conversion of the surface of silicones to silica (Banks et al., 1999,
1996; de Groh and McCollum, 1995) This frequently results in shrinkage and crack formation in the exposed silicones (Figs. 28.11 and 28.12) as they are transformed from low modulus polymers into the higher modulus silica.
The surfaces of materials with volatile oxidation products (such as hydrocarbon polymers) that are oriented in
a fixed position with respect to the ram direction gradually develop left-standing cones, which point in the direction of arriving atomic oxygen. Thus, the microscopic roughness of the surfaces increases over time. Surface
roughening caused by steady erosion, or deposition, which is independent of location and the locations are
IV. CASE STUDIES
618
28. DEGRADATION OF SPACECRAFT MATERIALS
FIGURE 28.11 Effect of LEO atomic oxygen exposure (fluence of 2.3 3 1020 atoms/cm2, Shuttle flight STS-46) on DC 93-500 silicone: (A)
Photo of unexposed DC 93-500 silicone; (B) photo of the same surface following exposure.
FIGURE 28.12
Scanning electron microscope photo of DC 93-500 silicone showing cracking and subsequent branch cracking after atomic
oxygen exposure to an effective fluence of 2.6 3 1021 atoms/cm2 in a plasma asher facility.
random are both phenomena that obey Poisson statistics. This phenomena also occurs in grit blasting of glass
windows, deposition of snow flakes, and the flux of gamma rays. A consequence of this kind of process is that
the roughness grows as the square root of the duration of deposition or erosion (Banks et al., 2001). Fig. 28.13
shows typical atomic oxygen textured surfaces of Kapton H polyimide, fluorinated ethylene propylene, and
chlorotrifluoroethylene after fixed-orientation exposure to atomic oxygen in LEO (de Groh et al., 1995; Banks
et al., 1991). In addition to polymer thickness loss, this type of texturing causes an increase in diffuse reflectance
and a decrease in specular transmittance of polymers (Banks et al., 1985; Waters et al., 2009). Atomic oxygen
exposure of hydrocarbon or halocarbon polymers that are pigmented or filled with metal-oxide particles causes
erosion of the polymeric content, resulting in gradual exposure of an increasing surface population of metaloxide particles that are poorly attached to each other Silverman (1995b). The metal-oxide particles, which become
loosely attached, remain in contact and gradually shield the underlying polymer content from atomic oxygen erosion. Thus, the erosion yield can gradually decrease with atomic oxygen fluence. The example in Fig. 28.14 shows
decreasing erosion yield values of white Tedlar with increasing atomic oxygen fluence for samples flown on
numerous MISSE missions (de Groh et al., 2017).
28.2.3 Mitigation Techniques
Atomic oxygen erosion of thin polymers in LEO has been a challenging spacecraft performance and durability
problem for many years. Three approaches have been used to reduce or eliminate atomic oxygen erosion of polymers: (1) the application of thin-film protective coatings made of atomic oxygen durable materials; (2) the
IV. CASE STUDIES
619
28.2 ATOMIC OXYGEN EFFECTS
FIGURE 28.13 Scanning electron microscope photos of LEO atomic oxygen textured polymers: (A) Kapton H polyimide (atomic oxygen
fluence 5 2.3 3 1020 atoms/cm2); (B) fluorinated ethylene propylene, silverized-Teflon FEP (atomic oxygen fluence 5 8.43 3 1021 atoms/cm2);
and (C) chlorotrifluoroethylene (atomic oxygen fluence 5 8.99 3 1021 atoms/cm2).
Atomic oxygen erosion yield (cm3/atom)
2.0E–25
1.67E–25
1.54E–25
MISSE 7
1.5E–25
y = -7E-26ln(x) + 4E–24
R2 = 0.9926
[1.48E-25]
1.45E–25
MISSE 8
1.01E–25
MISSE 2
1.0E–25
5.0E–26
0.0E+00
5.0E+21
1.0E+22
AO fluence (atoms/cm2)
FIGURE 28.14 Erosion yield versus atomic oxygen fluence for white Tedlar samples flown on MISSE 2, 7, and 8 showing decreasing erosion yield with increasing atomic oxygen fluence (de Groh et al., 2017).
modification of the surface of the polymers to make them more durable to atomic oxygen; and (3) the use of alternative polymers that contain metal atoms that develop a protective coating with atomic oxygen exposure.
The first and most widely used mitigation approach is the application of thin film metal, metal-oxide, or
fluoropolymer-filled metal-oxide protective coatings to polymers (Banks et al., 1984, 2002; Dever et al., 1996;
Rutledge and Mihelcic, 1990; Goode et al., 1994). Thin-film coatings of SiO2, Al2O3, Indium Tin Oxide, Ge, Si, Al,
IV. CASE STUDIES
620
28. DEGRADATION OF SPACECRAFT MATERIALS
FIGURE 28.15 Scanning electron microscope photos of LEO atomic oxygen exposed Kapton from LDEF at sites of pinwindow and crack
defects in a vacuum-deposited aluminum protective coating: (A) Prior to removal of aluminized coating; (B) after chemical removal of protective coating.
and Au with thickness ranging from a few hundred to more than 100 nm are typically applied by sputter deposition or vapor deposition. For example, the SiO2 coatings on Kapton H polyimide for the solar array blankets on
the International Space Station are 130 nm thick and applied by magnetron sputter deposition (Rutledge and
Olle, 1993). Although metal-oxide coatings as thin as B5.0 nm can provide atomic oxygen protection on ultrasmooth surfaces, thicknesses of B100 nm are usually used to ensure complete coverage over irregularities of
debris, pits, and rills on polymer surfaces. Coatings that are thicker than 100 nm can more easily crack or spall
due to either their intrinsic stress or inability to conform with flexure compression or expansion at their polymer
substrates. The addition of fluoropolymer content to metal-oxide coatings allows greater strain-to-failure in the
coatings. Such coatings can be deposited by cosputter deposition of metal oxide and polytetrafluoroethylene
Teflon (Banks et al., 1985, 1984).
The atomic oxygen durability of polymers protected by thin-film coatings made of materials (which are themselves atomic oxygen durable) is largely dependent on the number and size of the pinwindow and scratch defects
in the protective coatings (Fig. 28.15). The application of 130-nm SiO2 protective coatings on Kapton polymide
can frequently reduce the rate of weight loss due to atomic oxygen erosion of Kapton to less than 1% of that of
unprotected Kapton (Rutledge and Olle, 1993).
Atomic oxygen undercutting oxidation at sites of pinwindow and scratch defects can ultimately lead to
mechanical failure of the polymer when a sufficient number of undercut cavities connect (Banks et al., 2002). The
growth of undercut cavities has been studied for polymer films coated on one side or both sides using Monte
Carlo computational modeling (Banks et al., 1990, 1998, 1992, 2003; Banks and Stueber, 1997).
One approach to reducing the number of pinwindow and scratch defects in atomic oxygen protective coatings
is to apply a surface tension leveling coating to the material prior to applying the protective coating. Studies have
found that leveling coatings can successfully increase the atomic oxygen durability of protective coatings on composite materials based on decreasing pinwindow and scratch defect densities (Jaworske et al., 1994; de Groh
et al., 1990). For example, in one study a low-viscosity epoxy was applied to the surface of several composite cou
pons. A protective layer of 1,000 A of SiO2 was deposited on top of the leveling coating, and the coupons were
exposed to an atomic oxygen environment in a plasma asher. Pinhole populations per unit area were estimated
by counting the number of undercut sites observed by scanning electron microscopy. Defect density values of
180,000 defects/cm2 were reduced to about 1,000 defects/cm2 as a result of the applied leveling coating
(Jaworske et al., 1994). Leveling coatings have also been found to improve the optical performance of composite
concentrator surfaces by improving the specular reflectance (Jaworske et al., 1994; de Groh et al., 1990).
The mitigation approach involving modification to the surface of polymers to make them more durable to
atomic oxygen has primarily involved either implantation of metal atoms into the surface of the polymer
(Iskanderova et al., 1999) or chemical modification of the surface of the polymer to incorporate silicon atoms into
the surface and near the surface. In both surface modification approaches, the degree to which the atomic oxygen
erosion yield is reduced is dependent upon the aerial density of metal atoms that can be placed into the polymer
surface.
The formation of alternative polymers that contain inorganic atoms has been approached through a variety of
chemical formulations including the use of silicone copolymers (Rutledge et al., 1990) polysilsesquioxane,
(Gonzalez et al., 2003), cage coordination compound incorporation of metal atoms, (Illingsworth et al., 2001), and
phosphorous-containing polymers (Shepp et al., 1999). As with the surface alteration approach, the durability of
the alternative polymer is dependent on the aerial density of inorganic atoms that can be achieved. In addition,
IV. CASE STUDIES
28.3 CONTAMINATION EFFECTS
621
the alternative polymers must be durable to other space environmental threats, such as UV and ionizing radiation, in addition to atomic oxygen, to be considerer suitable. Other properties, such as low solar absorbtance,
may be required, depending on the application. A benefit of some of the alternative polymers is lower solar
absorbtance than Kapton (Shepp et al., 1999).
28.3 CONTAMINATION EFFECTS
28.3.1 Sources and Transport of Spacecraft Contamination
Spacecraft contamination can be defined as molecular or particulate matter on or near a spacecraft surface that
is foreign to that surface. Sources of spacecraft contamination can include thruster propellants and burn residue,
outgassing of spacecraft materials, vented gases from spacecraft systems, fluids released from the spacecraft
by dumping or leakage, micrometeoroids and orbital debris, and particles generated or redistributed during
spacecraft mechanical operations or astronaut extravehicular activity (EVA) operations (Chen, 2001).
Comprehensive data on outgassing of spacecraft materials is found in Walter and Scialdone (1997). Space environment interactions with materials can also produce contaminants, such as volatile products of atomic oxygen
reactions and UV-induced or radiation-induced chain-scission products in polymer materials, and residual nonoxidative films left free-standing due to atomic oxygen erosion of underlying material. Space environment effects,
such as atomic oxygen, UV, and radiation interactions, can further modify contaminant species.
Spacecraft contaminants can either deposit onto spacecraft surfaces or remain in the vicinity of the spacecraft.
Molecular contaminants can transport from surface-to-surface through various mechanisms, including line-ofsight transport, nonline-of-sight transport through reflection or scattering, and attraction of positively ionized
contaminants to a negatively charged, sunlit spacecraft surface (Tribble, 1995). These transport mechanisms can
put critical spacecraft surfaces at risk for contamination effects.
28.3.2 Contamination Effects on Spacecraft Surfaces
Buildup of molecular or particulate spacecraft contamination can cause degradation in transmittance, reflectance, solar absorptance, and thermal emittance of surfaces. The impacts of this degradation include reduced performance of solar arrays, radiators, instrument optics, sensors, and other systems (Chen, 2001).
Some particularly detrimental cases of space environment interactions with spacecraft contamination include
atomic oxygen oxidation of outgassed silicones to produce a noneroding silica-based layer, and UV or ionizing
radiation interactions with contaminants to produce a contaminant film. Examples of these cases will be discussed in the following.
28.3.2.1 Atomic Oxygen Interaction With Silicones
Almost all spacecraft have silicones on board in the form of adhesives, potting compounds, and lubricants
used in materials processing. Although most LEO spacecraft designers make efforts to use only silicones that are
vacuum stripped to eliminate or reduce the amount of volatile short-chain content, silicone fragments are often
evolved in the vacuum environment in LEO, with the process being further enhanced with atomic oxygen and/
or radiation-induced bond breaking. The resulting silicone fragments can deposit on surfaces that are exposed to
atomic oxygen. If the surfaces are not receiving atomic oxygen, then simple re-evaporation of the silicone can
occur, providing that the surfaces are the same temperature or hotter than the source of the silicone. Or, UV may
interact with the silicone fragments causing a polymerized contaminant layer to build up, as will be discussed in
the following section. With atomic oxygen arrival, oxidation reactions cause the silicones to lose hydrocarbon
content and convert to a silica-based surface layer that is resistant to atomic oxygen erosion. Such processes
occurred on the MIR Space Station, resulting in the accumulation (over a 10-year period) of a microscopically
rough coating on the solar array (Fig. 28.16), which was up to 4.6 μm thick (Banks et al., 1999). Such coatings tend
to be rather transparent. However, if the silicone deposition is also accompanied by hydrocarbon deposition, a
much more optically absorbing coating can result (Noter et al., 1996). Fig. 28.17 shows the antisolar side of one of
the MIR solar-array solar cells. The contamination is highly absorbing. It appears as a tan-colored silica deposit
formed by atomic oxygen reaction of silicones that were arriving at the same time hydrocarbons arrived from a
polymer mesh behind the array. This contaminant layer was B1.24 μm thick.
IV. CASE STUDIES
622
28. DEGRADATION OF SPACECRAFT MATERIALS
FIGURE 28.16 MIR solar array sun-facing surface after 10 years in LEO: (A) Photo showing silica contamination as a diffuse white deposit
on the front surface of the solar cells; (B) scanning electron microscope photo of oxidized silicone contamination layer.
28.3.2.2 Photochemical Deposition of Contaminants
In the presence of UV light, contaminants can form a polymerized film on a spacecraft surface (Tribble, 1995).
It has also been suggested that charged particles may play a role in this polymerization process, either separately
or synergistically (Chen, 2001). Even though outgassing rates decrease with time, subsequent contaminant deposition rates do not decrease at the same rate. This is because photochemical contamination processes have been
found to continue even when outgassing has subsided (Chen. 2001). This can be a significant concern on long
missions.
The LDEF satellite, which was exposed for 69 months in the LEO environment, provided an interesting study
of contamination processes as the leading edge received a high fluence of directed ram atomic oxygen along with
solar exposure, and the trailing edge received very little atomic oxygen with similar solar exposure. Molecular
contamination processes on the LDEF surfaces varied from the leading edge to the trailing edge. Surfaces on the
leading edge experienced removal of hydrocarbon-based contaminants and oxidation of silicone fragments leading to silica-rich, hydrocarbon-poor contaminant films. Surfaces on the trailing edge experienced the buildup of
hydrocarbon contaminants and silicone fragments, which became UV-darkened hydrocarbon-rich, silicone
contaminant films.
28.3.3 Mitigation of Contamination Effects
Mitigation of spacecraft contamination effects is achieved through careful selection of materials (particularly
low outgassing materials), a spacecraft design that minimizes contamination risk for critical components, and
taking precautions for spacecraft cleanliness during ground assembly and on-orbit operations. Spacecraft
IV. CASE STUDIES
28.4 SPACE RADIATION EFFECTS
623
FIGURE 28.17 Back surface of MIR solar cell after 10 years in LEO: (A) photo of contamination on solar cell back-surface prior to “tape
peel cleaning”; (B) photo of the same solar cell back-surface after “tape peel cleaning” of the upper-right corner to show a comparison of the
contaminated and cleaned areas.
materials outgassing data can be found in Walter and Scialdone (1997), and detailed guidelines for spacecraft
contamination control can be found in Tribble et al. (1996).
28.4 SPACE RADIATION EFFECTS
28.4.1 Solar UV Radiation
28.4.1.1 Environment Description
The Earth’s atmosphere absorbs all UV radiation from the Sun that is less than 0.3 μm in wavelength; (Tribble,
1995) however, spacecraft outside of the Earth’s atmosphere and with a view of the Sun are subject to the full
solar spectrum. Additionally, spacecraft surfaces without a direct view of the Sun, but with a view of Earth, may
still experience solar UV effects due to Earth albedo, which is the Sun’s energy (B31%) reflected back to space by
the Earth’s atmosphere (Tribble, 1995).
The solar spectrum outside the Earth’s atmosphere at one astronomical unit from the Sun is referred to as the
air mass zero (AM0) solar spectrum. Fig. 28.18A shows the AMO solar spectrum in the wavelength range up to
250 nm and 28.18B shows the UV region of the AMO solar spectrum (American Society for Testing and
Materials, 1992). In general, the UV range is defined as the portion of the electromagnetic spectrum including
wavelengths between 4 and 400 nm (Koller, 1965). However, relevant to materials degradation in space, it is
convenient to examine only those wavelengths that are significant in the AM0 solar spectrum. Integrated solar
irradiance in various wavelength ranges are shown in Table 28.4 (American Society for Testing and Materials,
1992) along with the portion of the solar constant represented by each wavelength range. It is evident that wavelengths shorter than approximately 120 nm represent a negligible portion of the solar spectrum. Relevant to the
study of space environment effects on materials, the UV range of wavelengths can be conveniently divided into
two bands: Near UV (NUV) as the 200 2 400 nm range and vacuum UV (VUV) as the 100 2 200 nm range.
IV. CASE STUDIES
624
28. DEGRADATION OF SPACECRAFT MATERIALS
(A) 0.00025
Irradiance (W/cm2 nm)
0.00020
0.00015
0.00010
0.00005
0.00000
0
250
500
750
1000 1250 1500 1750
Wavelength (nm)
2000 2250 2500
(B) 1.E−03
Irradiance (W/cm2 nm)
1.E−04
1.E−05
1.E−06
1.E−07
1.E−08
1.E−09
1.E−10
200
300
Wavelength (nm)
100
400
FIGURE 28.18 The air mass zero (AM0) solar spectrum: (A) the wavelength range up to 2,500 nm; (B) the UV portion of the AM0 solar
spectrum from 100 2 400 nm.
TABLE 28.4
Air Mass Zero (AM0) Integrated Solar Irradiance (American Society for Testing and Materials, 1992)
Wavelength range irradiance
Integrated solar in wavelength range (W/m2)
Percent of solar constant within wavelength range
λ # 1000 μm
1,366.1
100
λ # 400 nm
106.6
7.8
λ # 200 nm
0.10
λ # 120.5 nm
0.007
24
2.28 3 1025
3.12 3 10
28.4.1.2 Effects on Materials
Polymers are particularly susceptible to UV radiation degradation, because many types of bonds in organic
polymers are capable of absorbing UV light, which can lead to photochemical reactions (Tribble, 1995; Schnabel,
1981). Additionally, impurities that are usually contained in synthetic polymers are often more likely to absorb
IV. CASE STUDIES
625
28.4 SPACE RADIATION EFFECTS
1.0
300 nm
250 nm
0.9
0.8
224 nm
200 nm
Transmittance
0.7
190 nm
0.6
184 nm
0.5
0.4
0.3
0.2
0.1
0.0
0
50
100
150
Thickness (µm)
FIGURE 28.19
Transmittance as a function of FEP thickness.
UV light than the polymer itself, and can be significant contributors to the photochemical reactions within the
polymer material (Schnabel, 1981). Photochemical reactions within organic molecules may result in effects such
as discoloration of the material (increase in solar absorptance) or loss of mechanical properties due to chemical
changes in the material. Important considerations for polymer UV durability in space are the wavelengths
required to cause degradation, the depth of degradation, and synergistic effects with other environmental factors.
It is generally thought that most polymers absorb approximately 95% of incident radiation below 250 nm within
0.3 μm from the surface (Adams, 1993). However, based on the data shown in Fig. 28.19, Teflon FEP is an exception,
as it transmits a significant amount of UV radiation through tens of micrometers in depth (Dever et al., 2001).
For polymer films whose thickness is significantly greater than the UV attenuation depth (the depth within
which the majority of UV light as absorbed), the undegraded portion of the polymer thickness provides support
to a degraded surface. However, for applications using polymer films whose thickness is on the order of the UV
attenuation depth, the potential for UV degradation resulting in cracking of the full film thickness is significant.
Glass and ceramic materials have been observed to undergo UV radiation-induced darkening (Koller, 1965;
Gatto et al., 1998), also referred to as solarization. UV-light interactions in glass and ceramic materials can cause
formation of electrons or holes that are trapped in various defects. Some of these trapped species absorb light in
specific wavelength ranges and are referred to as color centers (Gatto et al., 1998). UV darkening can be affected
by purity of the material, particle shape and size, surface chemistry, and thermal history (Harada and
Deshpande, 1994). UV darkening can be detrimental to spacecraft materials such as white-paint coatings and
solar-cell cover glass.
In order to understand the effects of space UV radiation on materials, it is useful to examine the results of the
laboratory testing reported in the literature. However, results should be cautiously interpreted, because there are
a multitude of variables associated with UV testing that make it difficult to compare results directly from one test
to another. These variables include: wavelength range, spectral shape and intensity of the UV source, and
the nature of the exposure environment (air, purge gas, or vacuum). It is important to note that UV-induced degradation reactions in polymers are also influenced by oxygen, where the quantum yield (i.e., number of scission
products per incident photon) is often greater for reactions in the presence of oxygen (Schnabel, 1981; McKellar
and Allen, 1979). Even vacuum systems contain some amount of oxygen from residual air in the system. For
some materials, especially semiconductor pigmented paints, postirradiation exposure to air (as is often required
for postexposure analysis) can cause reversal of the UV-induced damage, sometimes referred to as “air bleaching,” creating an additional complication for interpreting test results (Harada and Deshpande, 1994).
28.4.1.2.1 Effects on Fluoropolymers
UV interactions with fluoropolymers have been found to cause loss of mechanical properties and loss of optical properties. Erosion, mass loss, and surface roughness increases have been observed for Teflon FEP upon
exposure to a broad-spectrum ( . 115 nm) VUV deuterium lamp providing wavelengths greater than 115 nm in
IV. CASE STUDIES
626
28. DEGRADATION OF SPACECRAFT MATERIALS
vacuum (Grossman et al., 1997; Skurat et al., 1997). Erosion of FEP was also observed upon exposure to monochromatic light of 147 nm (Skurat et al., 1997). Broad-spectrum VUV ( . 115 nm) also caused mechanical properties degradation for FEP (Skurat et al., 1997). One study (Dever and McCracken, 2003) examined the effects of
broad-spectrum VUV light on FEP using windows to produce lower cut-off wavelengths of 115 nm, 140 nm, and
155 nm. All wavelength ranges were found to produce degradation in the strength and elongation-to-failure of
FEP, indicating that wavelengths greater than 155 nm are capable of degrading FEP. In another study, surface
hardness of FEP was found to increase as a function of increasing exposure to broad-spectrum VUV ( . 115 nm),
where surface hardness was used as a measure of surface embrittlement. In this test, surface hardness was
analyzed using atomic force microscopy techniques (Dever and de Groh, 2002). Teflon polytetrafluoroethylene
(PTFE) was found to be more susceptible to UV damage than FEP (Skurat et al., 1997).
Various synergistic effects were observed for fluoropolymers exposed to UV radiation during heating or
thermal cycling. Tensile strength and elongation-to-failure were found to decrease for PTFE polymers upon UV
exposure in the 185 2 369-nm wavelength range in vacuum or nitrogen at temperatures of 21 2 315 C, suggesting
that chain scission occurs in PTFE polymers (Jukeikis, 1979). UV exposure of FEP resulted in chain scission at
ambient or slightly higher temperatures, and cross-linking above 80 C (Jukeikis, 1979). VUV exposure (147-nm
light) at temperatures of 120 C and 150 C resulted in a significant decrease in tensile strength and elongation
compared to the negligible change observed for VUV alone (Skurat et al., 1997). Compared to VUV irradiation
with steady-state heating at 100 C, creep deformation increased for VUV irradiation with thermal cycling, which
was attributed to FEP traversing between phases during thermal cycling (Toupikov et al., 1997).
Tedlar (polyvinyl fluoride) is found to undergo loss of strength and elongation due to UV exposure as
reported in the literature for terrestrial uses of Tedlar (DuPont, Technical Information on Tedlar, 2003).
Regarding testing for aerospace applications, Tedlar materials were exposed to at least 2,500 equivalent sun hours
in a facility that provided VUV ( . 115 nm) and NUV (B200 2 400 nm) to simulate LEO and GEO solar conditions (Stuckey and Meshishnek, 2000). Results from this test showed that solar absorptance for an uncoated cloud
white Tedlar increased. Use of an Optical Coating Laboratory, Inc. (OCLI) multilayer coating prevented a significant change in solar absorptance. The OCLI coating was thus found to perform well as a UV protective coating
for white Tedla. However, another investigation showed that an OCLI coating did not adhere well to FEP (Dever
and de Groh, 2002),which will be described in Section 28.5.2.3.
While VUV radiation from a deuterium lamp (λ . 115 nm) was found to cause mass loss of FEP, no mass loss
was observed for unpigmented Tedlar (Grossman et al., 1997). Lack of mass loss for Tedlar was attributed to the
fact that, unlike FEP, Tedlar contains CH bonds. It was suggested that Tedlar undergoes cross-linking that modifies the surface layer.
As illustrated through the examples above, UV radiation and synergistic effects with heating or thermal
cycling are known to cause degradation of mechanical properties and optical properties of fluoropolymers. The
degree to which degradation occurs depends upon irradiation wavelengths, exposure temperatures, and the
chemical nature of the fluoropolymer.
It is important to note that the spectral mismatch between the VUV deuterium lamp and the AM0 Sun may
produce degradation mechanisms that are different between space and laboratory exposures. For example, the
lamp’s peak at around 160 nm coincides with a peak of Teflon UV absorption (Seki et al., 1990). There is no such
peak in the solar spectrum. The VUV lamp peak may result in enhanced surface reactions for Teflon, which
would not occur to as great an extent in space.
28.4.1.2.2 Effects on Polyimide Kapton
Polyimide Kapton shows minor changes in properties upon VUV exposure. In one study, (Dever et al., 2001) a
sample of 25.4-μm Kapton HN exposed to 1,100 equivalent VUV sun hours (115 2 200 nm VUV) did not experience any statistically significant changes in elongation or ultimate tensile strength. Minor changes in the reflectance spectrum in the UV-visible wavelengths due to this exposure may be indicative of some changes in surface
chemistry, however. A minor increase in solar absorptance due to this VUV exposure, from approximately 0.23
to approximately 0.24, was also observed.
28.4.1.2.3 Effects on Epoxy Composites and Adhesives
In one study, it was found that UV exposure (220 2 300 nm wavelength range) of epoxy resin caused degradation within the top 0.5 2 1.0 μm layer with as little as 30 minutes of exposure (Luoma and Rowland, 1989).
Weight loss of epoxy composite materials was observed and was found to increase with increased UV
irradiation time for a test conducted using an accelerated weathering tester consisting of medium-wave UV
IV. CASE STUDIES
28.4 SPACE RADIATION EFFECTS
627
(200 2 300 nm) lamps in a system controlled at 50 C (Liau and Tseng, 1998). Another test found that UV irradiation from medium-wave (200 2 300 nm) UV lamps initiated microcracks in graphite epoxy and glass epoxy
composites, which then propagated upon thermal shock testing between 121 C and liquid nitrogen temperature
(2196 C) (Liau and Tseng, 1998). On the LDEF, materials were exposed to the space environment for 69 months,
and Hysol 934 epoxy adhesive was found to discolor and Hysol EA 9628 was found to have a decreased shear
strength following flight. These effects were attributed to UV degradation (Silverman, 1995a).
28.4.1.2.4 Effects on White-Paint Coatings
Space-stable coatings with low solar absorptance and high thermal emittance are important to spacecraft thermal control. Z93 and YB-71 paints, with zinc-oxide and zinc-orthotitanate pigments, respectively, in potassiumsilicate binders, have proven to maintain their thermo-optical properties in the space environment (Harada and
Deshpande, 1994). A paint formulation with zinc-oxide pigment in a methyl silicone binder, designated as S13G/
LO, provides a more flexible option for white-paint coating (Harada and Deshpande, 1994). Development of these
three types of paints was based on the UV stability of the ZnO and Zn2TiO4 pigments. Other pigments examined
were found to undergo significant UV degradation, including zirconia, alumina, and silica pigments (Harada and
Deshpande, 1994). Although Z93, YB-71, and S13G/LO paints have shown appropriate space stability, the S13G/
LO material degrades in UV somewhat faster than YB-71 or Z93 and thus its use is limited to shorter missions.
For example, for 5,000 equivalent sun hours of UV exposure, solar absorptance increases of 0.02, 0.03, and 0.06
were observed for YB-71, Z93, and S13G/LO, respectively (Harada and Deshpande, 1994).
28.4.2 Ionizing Radiation
The ionizing radiation environment of space includes energetic charged species such as electrons and protons,
and energetic photons such as X-rays and gamma rays. These energetic particles and photons are considered ionizing radiation because they ionize atoms as they move through a material. The extent of high-energy particle
interactions with materials depends upon the type of radiation, its energy, and the material. Details on the interactions of radiation with matter are comprehensively addressed elsewhere (Holmes-Siedle and Adams, 2002;
Hastings and Garrett, 1996). Materials degradation due to ionizing radiation can occur based on accumulation of
absorbed energy in the material, which is referred to as total ionizing dose. Additionally, the rate at which the
energy is deposited, referred to as dose rate, may also influence materials degradation.
Radiation dose, referred to in SI units of grays (Gy), is the amount of radiation that deposits 1 J of energy per
1 kg mass of material. Another commonly used unit is the rad, which is 0.01 Gy. Sources of ionizing radiation
within the space environment are described in the following.
28.4.2.1 Solar Flare X-ray Radiation Environment
Solar flares are releases of intense energy from the Sun occurring over a short duration (minutes to hours) and
observed as sudden brightening of the chromosphere of the Sun and producing energy throughout the electromagnetic spectrum from radio waves to gamma rays (Tribble, 1995; Website of the National Oceanic and
Atmospheric Administration’s Space Environment Center). During the approximately 11-year solar cycle, periods
of high solar activity are correlated with enhanced X-ray emission (Tribble, 1995).
For spacecraft in Earth orbits, surfaces with a view of the Sun during periods of solar flares will be exposed to
a significant flux of X-rays. The Geosynchronous Operational Environmental Satellites (GOES) have been monitoring the space solar X-ray environment since 1986, from their location in geosynchronous orbit at an altitude of
approximately 35,800 km. GOES solar X-ray flux data, available through the National Oceanic and Atmospheric
Administration (NOAA) National Geophysical Data Center (NGDC) (Website of the National Oceanic and
Atmospheric Administration’s National Geophysical Data Center), were used to estimate the solar flare X-ray
environment in wavelength regions of 0.1 2 0.8 nm, 0.05 2 0.4 nm, and 0.0124 2 0.05 nm for the HST mission in
LEO (Dever et al., 2000). These X-ray fluences and estimated dose-depth profiles for FEP are shown in
Fig. 28.20A.
28.4.2.2 Charged-Particle Radiation Environment
The three main sources of charged-particle radiation naturally occurring in space are galactic cosmic rays, solar
proton events, and the trapped radiation belts. Galactic cosmic rays (GCRs) consist of low flux ionized nuclei,
mostly protons, generally providing a very low radiation dose rate, which is highest at solar minimum (Tribble,
IV. CASE STUDIES
628
28. DEGRADATION OF SPACECRAFT MATERIALS
10,000
(A)
Dose (krads)
1000
0.1−0.8 nm, 3.6 yr
0.1−0.8 nm, 6.8 yr
0.1−0.8 nm, 9.6 yr
0.1−0.8 nm, 20 yr
0.05−0.4 nm, 3.6 yr
0.05−0.4 nm, 6.8 yr
0.05−0.4 nm, 9.6 yr
0.5−0.4 nm, 20 yr
100
10
1
0.1
0
0.02
0.04
0.06
0.08
Depth (mm)
0.1
0.12
0.14
10,000
(B)
Dose (krads)
1000
Electrons, 3.6 yr
Electrons, 6.8 yr
Electrons, 9.6 yr
Electrons, 20 yr
Protons, 3.6 yr
Protons, 6.8 yr
Protons, 9.6 yr
Protons, 20 yr
100
10
1
0.1
0
0.02
0.04
0.06
0.08
Depth (mm)
0.1
0.12
0.14
FIGURE 28.20 Dose-depth profiles for ionizing radiation in FEP for (A) solar flare X-ray exposure; (B) trapped electron and proton radiation exposure for the HST environment.
1995). GCR radiation consists of ions of all elements of the periodic table and is composed of approximately 83%
protons, 13% alpha particles (4He ions), 3% electrons, and 1% of heavier nuclei (Barth, 2003). Energies of GCR
particles range from about 108 2 1019 eV (Tribble, 1995). Low altitude/inclination orbits are protected from some
of the GCRs, because when a GCR approaches Earth in the plane of the equator, Earth’s magnetic field bends the
particle back to space or to the polar regions, depending on its initial direction and energy (Tribble, 1995).
Solar proton events (SPEs) result from coronal mass ejections producing significant proton flux over short duration periods of, on average, 15 days (Tribble, 1995). Some SPEs are heavy-ion-rich with energies ranging from
tens of MeV/per nucleon to hundreds of GeV/per nucleon (Barth, 2003). Spacecraft in low Earth orbits, such as the
HST, are generally protected by the magnetosphere from the majority of SPE proton flux, but SPEs are hazardous
to spacecraft in high inclination orbits and geosynchronous orbit (Dever et al., 2000). In the trapped radiation belts,
also called the Van Allen belts, energetic electrons and protons are confined to gyrate around the Earth’s magnetic
field lines. Trapped electrons have energies up to tens of MeV and trapped protons and heavier ions have energies
up to hundreds of MeV (Barth, 2003). Fluxes of protons and electrons in the trapped radiation belts are a function
of particle energy, altitude, inclination, and solar activity and can increase during solar storms (Tribble, 1995; Barth,
2003). Peak fluxes of both electrons and protons occur at around 3,000 km and a second peak of electron flux occurs
at around 25,000 km altitude (Tribble, 1995). Additional variables affecting trapped radiation flux include effects of
IV. CASE STUDIES
629
28.4 SPACE RADIATION EFFECTS
the Earth’s poles and the South Atlantic Anomaly. Spacecraft in polar orbits experience greater charged-particle
dose rates than spacecraft in equatorial orbits due to the Earth’s magnetic field funneling charged particles into the
polar regions (Tribble, 1995). An asymmetry in the Earth’s magnetic field lines causes charged particles to reach
lower altitudes (,1000 km) in the South Atlantic, so that spacecraft experience higher dose rates when passing
over this region (Tribble, 1995; Barth, 2003).
Estimates of the trapped proton and electron fluences have been obtained using NASA’s proton (AP-8) and
electron (AE-8) models (Sawyer and Vette, 1976; Vette, 1991). The trapped electrons and protons are considered
to be omnidirectional by the AP-8 and AE-8 models, although some degree of east-west asymmetry has been
observed by spacecraft measurements since these models were developed (Sawyer and Vette, 1976). Because of
the approximately omnidirectional nature of the trapped radiation, all exterior spacecraft surfaces are exposed to
this radiation.
28.4.2.3 Space Ionizing Radiation Effects on Materials
28.4.2.3.1 Polymers
Ionizing radiation interacts with atomic nuclei and their surrounding electron clouds without specificity to
particular chemical bonds, unlike UV radiation, which is absorbed by particular chemical bonds (Schnabel, 1981).
Polymers can become degraded in physical and mechanical properties due to accumulated total dose of ionizing
radiation, which includes electrons, protons and X-rays, and GCRs.
Polymer films, such as FEP and Kapton, are commonly used on exterior spacecraft surfaces for thermal control
blankets and adhesively bonded radiator surfaces and thus receive an unshielded dose of space radiation.
Embrittlement and loss of strength are not significant concerns for radiator surfaces rigidly adhered to metal
structures. However, for polymer films that are not rigidly supported, such as in multilayer insulation applications, degradation in mechanical properties can result in cracking of the exterior polymer film layers. Cracking
in thermal blanket layers can compromise the thermal protection of underlying equipment and may pose a contamination risk if pieces of the cracked film become dislodged. In addition to thermal insulation blankets, other
uses of polymers on spacecraft exterior surfaces include structural composites, adhesives, potting materials, and
wiring/cable insulation.
As a guide to the effects of radiation on spacecraft polymers, Table 28.5 shows the ionizing radiation doses
that have been found to produce mild-to-moderate and moderate-to-severe polymer degradation (Holmes-Siedle
and Adams, 2002).
Relative degradation data such as that in Table 28.5, which are found commonly in radiation effects references
and which vary slightly among references (see also Tribble, 1995; Stuckey and Meshishnek, 2001), can be used as
a screening tool for materials selection, but it is important to consider that many other factors may play a role in
overall materials degradation in a space radiation environment. For example, the uniquely low solar absorptance
TABLE 28.5
Effect of Gamma Radiation Dose on Some Common Spacecraft Polymers (Holmes-Siedle and Adams, 2002)
Gamma radiation dose (rad)
Material
Mild-to-moderate damage
Moderate-to-severe damage
Teflon fluorinated ethylene propylene (FEP
1E6-8E6
8E6-2E7
Teflon polytetrafluoroethylene (PTFE)
2E4-1E5
1E5-2E6
Kapton, polyimide
1E8-1E10
1E10-1E11
Mylar, polyethylene terephthalate (PET)
4E6-1E8
1E8-1E9
Polyethylene
1E7-8E7
8E7-2E8
Polyurethane
1E9-5E9
5E9-2E10
Silicone
1E8-1E9
1E9-5E9
Epoxy
2E8-8E8
8E8-5E9
Nylon, polyamide
3E5-2E6
2E6-2E7
Polyvinyl chloride (PVC)
1E7-7E7
7E7-2E8
IV. CASE STUDIES
630
28. DEGRADATION OF SPACECRAFT MATERIALS
and high thermal emittance of FEP have led to its wide use for spacecraft thermal control, despite its susceptibility to radiation degradation.
While the data in Table 28.5 are for polymers exposed to gamma radiation in air, it is generally found that irradiation in oxygen or air produces more severe degradation of polymers than irradiation in vacuum. For example,
air or oxygen is known to significantly affect the radiation-induced degradation of PTFE, which undergoes
scission at a much higher yield in air than in vacuum (Schnabel, 1981). A similar finding was reported for X-ray
irradiated Teflon FEP, which experienced an increase in the degradation of mechanical properties after air exposure (de Groh and Gummow, 2001). Also, many polymers that predominantly cross-link when irradiated in vacuum will undergo main chain-scission when irradiated in air, indicating that not just the extent of degradation is
affected by the exposure environment, but the mechanism of degradation as well (Schnabel, 1981).
Table 28.5 does not account for dose rate effects or synergistic effects of radiation and temperature. One report
summarized that long-duration low-dose-rate exposures produced greater degradation of FEP than shortduration high-dose-rate exposure, indicative of a possible dose-rate effect. However, this apparent dose-rate
effect was attributed to time spent at elevated temperatures rather than to dose rate alone (Jukeikis, 1979). This
report also stated that accelerated ground testing for ionizing radiation degradation to FEP may underpredict inspace damage by factors of 3 to 10, considered as a crude guideline. Further synergistic considerations for radiation and temperature are discussed in section 28.5.2.3.
Newer applications for polymer films on spacecraft require large, lightweight, deployable, and/or
inflatable spacecraft structures, requiring significantly thinner materials than used in thermal blanket applications. Decreasing the thickness of a polymer increases the risk for radiation damage effects. This is because space
ionizing radiation deposits a decreasing dose at increasing depths within a material, so that the greatest damage
to a polymer film is at the surface and the damage decreases with increasing depth. An example of dose versus
depth in FEP, a widely used and radiation-vulnerable material, is shown in Fig. 28.20 as indicated before, and
trapped for solar flare X-rays and trapped electron and proton radiation on HST (Dever et al., 2000). When ionizing radiation degradation is combined with UV degradation (another effect that diminishes with increased polymer thickness), it is evident that ultrathin polymer films, such as gossamer structures, are at great risk of radiationinduced degradation. Because of the greater radiation sensitivity for thinner polymer films, accurate modeling of
the space environment and associated dose-depth profiles for polymer films is critically important to predicting the
durability of thin polymer films for use on large spacecraft structures for long missions. Methods of modeling the
space environment are described in Holmes-Siedle and Adams (2002), Stuckey and Meshishnek (2001), and methods for developing appropriate dose-depth profiles and space radiation simulation tests are described in Stuckey
and Meshishnek (2001).
Some polymer film materials have been studied for their radiation durability for use in deployable and inflatable spacecraft structures. Two studies (Russell et al., 2002; Dever et al., 2002) examined the effects of electron
and proton radiation, in some cases with UV radiation as well, on aromatic polyimide films to simulate exposure
near the Earth-Sun Lagrangian points 1 and 2. Films tested included Kapton HN, Kapton E, Upilex S, CP-1, CP-2,
and either TOR-RC (Russell et al., 2002) or TOR-LM (Dever et al., 2002). All materials were found to undergo
increased solar absorptance due to the radiation exposure.
In addition to polymer mechanical degradation, electrical properties of polymers may also be compromised
due to radiation. This is a particularly important effect for polymers used for electrical insulation, such as wiring
or cable insulation. Ionizing radiation has been found to induce conductivity in polymer films (Laghari and
Hammoud, 1990). This induced conductivity has been found to be material-dependent and dose- rate-dependent,
but independent of the type of ionizing radiation (Laghari and Hammoud, 1990). One study examined radiationinduced conductivity in various polymers upon exposure to gamma rays up to 106 Gy (Harrison, 1962). This
study found increases in conductivity for polyethylene and polystyrene of 2 to 3 orders of magnitude, less than
an order of magnitude increase for Teflon, and negligible increases in conductivity for nylon, epoxy, and polyvinyl chloride. However, this induced conductivity has been found to be temporary. Once the radiation exposure
was discontinued, those materials that experienced radiation-induced conductivity recovered to near-initial conductivity within seconds to hours (Harrison, 1962).
One complication of accelerated rate charged-particle radiation testing of polymer materials is the possibility
for charging of insulating polymer surfaces. PTFE, e.g., is an excellent electrical insulator. On-orbit, the low current and dose rates permit continuous discharge of exposed PTFE materials. However, in ground laboratory
accelerated tests, electrostatic charging of the PTFE target in high current/dose-rate laboratory electron beams
can lead to deceleration and deflection of the incident electron beam, especially for lower energy electrons
(Koontz et al., 2003). This effect needs to be considered when evaluating accelerated rate space simulation test
results for insulating materials.
IV. CASE STUDIES
631
28.5 THERMAL AND THERMAL CYCLING EFFECTS
TABLE 28.6 Solar Absorptance of Original and Reformulated White Paints Before and After Exposure to 2,600 Equivalent Sun Hours
UV and 3 3 1016 electrons/m2 at 40 keV Energy (Calculated from Original Data in Meshishnek et al., 1995)
White-paint
type
Number of
samples
Avg. solar absorptance
(pristine)
Avg. solar absorptance
(posttest)
Avg. solar absorptance
Increase
Z-93
6
0.118 6 0.014
0.155 6 0.013
0.037
Z-93P
6
0.109 6 0.001
0.152 6 0.019
0.043
S13GLO-1
6
0.166 6 0.010
0.287 6 0.014
0.121
S13GPLO-1
7
0.150 6 0.009
0.377 6 0.087
0.227
YB-71
7
0.103 6 0.021
0.238 6 0.054
0.135
YB-71 P
9
0.091 6 0.013
0.345 6 0.037
0.254
28.4.2.3.2 White-Paint Thermal Control Coatings
Many satellite systems have used white-paint thermal control coatings Z-93, YB-71, and S13GLO-1. As a result of
a change in manufacturer of the potassium-silicate binder/encapsulant material used in all three paints, extensive
requalification testing examined the radiation stability of the newly reformulated versions of these paints, Z-93P, YB71P, and S13GPLO-1, and compared them to their predecessors (Harada and Deshpande, 1994; Meshishnek et al.,
1995). In one of the tests, (Meshishnek et al., 1995) samples were exposed to approximately 2,600 equivalent UV sun
hours along with 40 keV electrons at a flux of 6 3 109 electrons/m2s for a total fluence of 3 3 1016 electrons/m2.
Table 28.6 shows solar absorptance changes for original and reformulated white paints due to this testing.
This study proved a very important point regarding spacecraft materials degradation. Even seemingly minor
changes in a material can produce significantly different vulnerability to radiation degradation. Radiation degradation of the white-paint materials is very sensitive to impurity levels and processing conditions, so that even an
“equivalent” chemical substitute for the potassium-silicate binder did not produce a material with equivalent
radiation durability. In fact, the YB-71P was found to be inadequately radiation stable to be space qualified,
whereas Z-93P, using the same binder, was found to have adequate radiation stability to be considered space
qualified (Meshishnek et al., 1995).
28.4.3 Mitigation of Radiation Degradation
In general, mitigation of radiation degradation of spacecraft external materials is accomplished through careful
material selection based on understanding radiation conditions for the mission and space UV and ionizing radiation durability of materials being considered. Where UV-vulnerable materials are necessary, some materials may
be able to be protected through the use of protective coatings, such as the Optical Coating Laboratory Inc. (OCLI)
coating used on Tedlar described in section 28.4.1.2.1. Although the OCLI coating was successfully used on
Tedlar, adhesion issues were observed when an OCLI coating was used on FEP, which will be described in section 28.5.2.3, so protection for one material may not be a guarantee of successful protection application for
another. Metallic coatings, such as vapor-deposited aluminum (VDA), can generally protect vulnerable polymers
from UV degradation, but processing defects, especially scratches or other types of line-shaped defects, may lead
to local UV degradation. Additionally, metallic coatings and oxide coatings cannot provide significant protection
from high-energy ionizing radiation. In most cases, vulnerable materials, even with protective coatings, should
be tested to the most conservative conditions to estimate their lifetime at mission conditions. It must be considered that radiation effects are often synergistic with thermal effects, which will be described in section 28.5, and
ground laboratory evaluations should consider all aspects of the space environment together.
28.5 THERMAL AND THERMAL CYCLING EFFECTS
28.5.1 Environment Description
Earth orbital environments are capable of producing significant temperature variations as the spacecraft passes
from sunlight to shadow. The number of thermal cycles expected for a mission depends upon the orbit. For
example, spacecraft in LEO complete one orbit approximately every 90 minutes and spacecraft in geosynchronous orbit complete one orbit each Earth day (24 hours). The range of temperature a material experiences during
IV. CASE STUDIES
632
28. DEGRADATION OF SPACECRAFT MATERIALS
thermal cycling depends on its thermo-optical properties (solar absorptance and thermal emittance), its view of
the Sun, its view of the Earth, its view of other surfaces of the spacecraft, durations of time in sun and in shadow,
and the influence of equipment or components that produce heat.
28.5.2 Effects on Materials
On-orbit temperatures and thermal cycling pose a threat to materials durability for various reasons. First, for
inhomogeneous materials in intimate contact with one another, such as in composites or coated materials, a mismatch in coefficients of thermal expansion may lead to cracking or delamination when the material experiences
significant temperature excursions. Second, mechanical properties of polymer materials can be a strong function
of temperature (Reed et al., 1973). Therefore, during the course of on-orbit thermal cycling, a polymer may experience temperatures at which it has decreased strength or ductility, making it more vulnerable to damage.
Because radiation damage is generally more concentrated at the exposed surface of a material and diminishes
through the thickness, one can consider that a radiation-damaged polymer no longer has homogenous properties
through its thickness, making it vulnerable to effects of a mismatch in coefficient of thermal expansion (CTE).
28.5.2.1 Effects on Composites
One report described the results of an in-depth study of the effects of thermal cycling on composites for space
applications including a carbon fiber-reinforced epoxy (ERL1962) and carbon fiber-reinforced cyanate ester (RS3)
(Brown, 1998). Variables in the study included thermal cycling temperature range, composite layer thickness,
matrix type, and fiber type. It was found that cyanate ester matrix composites are more resistant to thermal
cycling-induced microcracking than epoxy matrix composites. For carbon-fiber epoxy matrix composites, the
number of microcracks induced by thermal cycling reaches a saturation level as early as several hundred cycles,
with many microcracks being observed after only a few thermal cycles.
However, cyanate ester matrix composites show a threshold effect such that for less severe temperatures, thermal cycling does not cause any significant microcracking, and, for temperature ranges that do cause microcracking, a saturation level may be in the thousands of cycles. Also, composites with thinner laminate layers (50.8 μm)
were found to be more susceptible to microcracking than composites with thicker layers (127 μm).
The LDEF provided opportunities to examine materials degradation due to space exposure in low Earth orbit
for 69 months. Fig. 28.21 shows an example of an apparent space thermal cycling-induced crack in an aluminumchromium coating on a graphite epoxy panel that was located on the leading edge of the LDEF (de Groh et al.,
1995). Atomic oxygen erosion of the graphite epoxy is also evident in the cracked area.
28.5.2.2 Effects on Spacecraft Paint Coatings
Z-93, a zinc-oxide pigment/potassium silicate binder white paint applied to aluminum substrates, has been
observed to microcrack upon thermal cycling (Daneman et al., 1994). This is an expected result of the CTE
FIGURE 28.21 Backscattered electron micrograph showing thermal cycle-induced microcracks in an Al/Cr-coated graphite epoxy sample
that was located on the leading edge of the LDEF.
IV. CASE STUDIES
28.5 THERMAL AND THERMAL CYCLING EFFECTS
633
mismatch between the coating and the substrate. However, with proper substrate surface preparations, delamination and spalling are prevented, so, despite a “mud-tiled” appearance due to microcracking, the painted surfaces
exhibit durability to orbital thermal cycling. Anodized aluminum substrates can provide a less severe CTE mismatch for Z-93 and minimize the degree of microcracking (Daneman et al., 1994).
28.5.2.3 Synergistic Thermal, Thermal Cycling, and Radiation Effects on Uncoated and Coated Teflon
FEP
Synergistic effects between radiation and thermal cycling have been observed for spacecraft materials in
ground testing and in-space. In one test, thermal cycling was found to cause delamination and spalling of protective oxide coatings on FEP (Dever and de Groh, 2002). In this study, coatings examined included SiOx (where
xB2) and a coating consisting of alternating layers of SiO2, TiO2, and Ta2O3 developed by OCLI. Samples were
exposed to 5 kGy of 1 MeV electron radiation followed by thermal cycling, nominally from 2115 C to 190 C.
While some OCLI-coated samples that were unexposed to radiation and thermal cycling showed signs of minor
coating adhesion problems, such as cracking and loss of coating in bent areas, spalling and delamination were
observed only for samples that had been exposed to thermal cycling following radiation. An example of the
effects of electron radiation and thermal cycling on the OCLI/FEP is shown in Fig. 28.22. Severity of delamination
and spalling was found to be worse for thicker coatings.
The VDA-coated 127-μm-thick FEP outermost layer of the multilayer insulation blankets on the HST has
became embrittled, resulting in severe on-orbit cracking as shown in Fig. 28.23 (Townsend et al., 1999a). A sample of FEP retrieved during the second servicing mission (SM2), after 6.8 years in space, was significantly more
embrittled than the same thickness FEP retrieved during the third servicing mission (SM3A), after 9.7 years in
space. One of the differences in the environmental exposures between these samples was the maximum temperature exposure during thermal cycling. The retrieved SM2 insulation section curled after cracking, exposing the
lower emittance back-surface aluminum to space. It was estimated that this extremely embrittled piece of insulation reached approximately 200 C on-orbit, 150 C higher than the nominal temperature extreme (2100 C to
150 C for solar facing FEP on HST). A review board that investigated the severe FEP degradation on HST concluded that electron and proton radiation combined with on-orbit thermal cycling was necessary to cause the
observed cracking of FEP on HST at areas of stress concentrations (Townsend et al., 1999a). Several studies have
been conducted to investigate the space environmental factors responsible for the degradation of FEP on the
HST, including examinations of the combined effects of radiation and temperature or temperature cycling.
One study examined the differences in degradation produced by the space environment and ground testing
intended to replicate exposure conditions for FEP on the HST (Townsend et al., 1998). Samples of 127-μm FEP
film were exposed to 0.5 MeV electrons and 1 MeV protons to provide fluences equivalent to those of various
HST exposure durations up to 20 years. This radiation exposure was followed by thermal cycling in the temperature range of 2100 C to 150 C, the nominal range for the FEP exterior layer of thermal insulation on HST.
Thermal cycling was conducted in a nitrogen-purged chamber at a nominal rate of four cycles per minute. Effects
FIGURE 28.22 Sample of 2-mil-thick FEP with OCLI oxide coating (total coating thickness in range of 700 2 1,400 nm) following electron
radiation exposure and thermal cycling. Dark regions indicate areas where coating is missing.
IV. CASE STUDIES
634
FIGURE 28.23
28. DEGRADATION OF SPACECRAFT MATERIALS
Large cracks in outer layer of solar facing MLI on HST as observed during SM2, after 6.8 years in space (Townsend et al., 1999a).
of these exposures compared to HST exposures on strength and elongation of FEP are shown in Fig. 28.24. It is
evident that laboratory testing severely underpredicted the on-orbit degradation caused by HST exposure of FEP
for even 3.6 years, where laboratory exposures equivalent to 20 2 40 years in space were required for similar degradation. The data point for SM2 (6.8 yr) in Fig. 28.24, near-zero elongation, is shown for reference; however, as
described above, this FEP sample was exposed to an upper temperature limit of B200 C, much higher than the
nominal FEP materials on the HST, which reach 50 C.
Research has been conducted to determine the effects of heating on irradiated FEP in order to better understand the effect of temperature on the rate of degradation and on the mechanism of degradation of FEP insulation
in the LEO environment. In one study, samples of pristine FEP, X-ray irradiated FEP, and FEP retrieved from the
HST were heated from 50 C to 200 C at 25 C intervals in a high-vacuum furnace and evaluated for changes in
tensile properties and density (de Groh and Martin, 2002). Even though the ground laboratory X-ray exposure
(conducted at room temperature) provided an areal dose (D, the total energy absorbed per unit area integrated
through the full thickness in FEP) that was orders of magnitude higher than the HST on-orbit areal dose, it did
not produce the extent of damage observed for the HST-exposed FEP. However, the laboratory exposure provided sufficient degradation to show the effects of subsequent heating on irradiated FEP. This study found that
heating did not embrittle nonirradiated FEP Teflon; however, there was significant dependence of the embrittlement of irradiated FEP on heating temperature, with near complete loss of elongation at failure at 100 C and
higher. These results are shown in Fig. 28.25.
A section of the retrieved HST bistem thermal shields (BSTS), which experienced 8.25 years of space exposure,
was analyzed by de Groh et al. (2008) for space environmental durability. The shields were comprised of 2 mil
(0.051 mm) aluminized-Teflon fluorinated ethylene propylene (Al-FEP) rings fused together into a circular bellows shape. Because the circular thermal shields had solar, antisolar, and solar-grazing surfaces and were
exposed to the space environment for a long duration, it provided a unique opportunity to study the solar effects
on the environmental degradation of Al-FEP. Tensile testing verified that near-solar-facing surfaces lost their
mechanical strength and elasticity, while the antisolar-facing surfaces maintained their ductility. A polar plot of
the elongation at failure versus solar angle for retrieved HST BSTS weld tensile samples is provided in Fig. 28.26.
The study shows the detrimental effect of solar exposure on FEP insulation in the space environment. Antisolarfacing surfaces experienced only marginal degradation in space, but solar-facing surfaces experienced severe
IV. CASE STUDIES
635
28.5 THERMAL AND THERMAL CYCLING EFFECTS
(A) 30
Pristine
Radiation
Radiation + thermal cycling
Ultimate tensile strength (MPa)
25
HST, 3.6 yr
HST, 6.8 yr
20
15
10
5
0
0
20
80
40
60
Equivalent exposure years
100
Pristine
Radiation
Radiation + thermal cycling
(B) 400
350
Elongation to failure (%)
HST, 3.6 yr
300
HST, 6.8 yr
250
200
150
100
50
0
0
20
40
60
80
Equivalent exposure years
100
FIGURE 28.24 Charged-particle radiation and radiation with sequential thermal cycling effects on mechanical properties of 5-mil Teflon
FEP with comparison to FEP retrieved from the HST during SM1 (3.6 years in space) and during SM2 (6.8 years in space): (A) ultimate tensile
strength; (B) elongation at failure.
degradation due to the additional radiation exposure (soft X-rays and UV radiation) combined with increased
on-orbit heating for Sun-facing surfaces.
Teflon FEP tensile samples showed similar environmental degradation after 1.5 years of space exposure as
part of the MISSE 7 mission (de Groh et al., 2016). Fig. 28.27 provides plots of the percent elongation at failure
versus the solar exposure in equivalent sun hours (ESH) for back-surface aluminized-Teflon FEP (Al-FEP, 2 mil
thick) samples. Samples sectioned both parallel to the roll direction and normal to the roll direction were flown
and tested. These graphs, like the HST data, highlight the impact of solar exposure on the embrittlement of
Teflon FEP in space, as there is a clear correlation of decreasing elongation, and hence embrittlement, with
increasing solar exposure. Again, this is attributed to both the increased total solar radiation dose and/or the
increased solar heating for the higher solar exposures (de Groh et al., 2016).
This and other studies (de Groh et al., 2001) support the conclusion that radiation- induced chain-scission
(solar, X-ray, particle radiation) is the primary mechanism of embrittlement of FEP on the HST, and indicate the
significant impact of the on-orbit temperature of FEP with respect to its degradation in the space environment.
IV. CASE STUDIES
636
28. DEGRADATION OF SPACECRAFT MATERIALS
350
Pristine
=115
Elongation at failure (%)
300
250
200
X-ray
=115.5
Pristine
150
X-ray
HST
=101.5
100
HST SM3A
50
0
20
40
60
80
100
120
140
160
Vacuum heat treatment temperature (°C)
180
200
FIGURE 28.25 Percent elongation at failure of pristine, ground laboratory X-ray exposed and retrieved HST FEP as a function of vacuum
heat-treatment temperature.
90
120
60
150
30
180
0
250
200
150
100
50
0
50
100
150
210
200
250
330
240
300
270
HST BSTS (Elongation (%) vs. Solar Angle)
Pristine 2 mil Al-FEP BSTS (216.2%)
FIGURE 28.26
Elongation at failure (%) of retrieved HST BSTS samples versus solar angle (de Groh et al., 2008).
IV. CASE STUDIES
637
28.5 THERMAL AND THERMAL CYCLING EFFECTS
300
Nadir
250
Parallel to roll direction
% Elongation at failure
Nadir
Normal to roll direction
200
150
Wake
100
Zenith
Ram
50
Zenith
0
0
1000
2000
3000
4000
5000
Equivalent sun hours (ESH)
FIGURE 28.27
Percent elongation at failure (%E) versus ESH for MISSE 7 Al-FEP samples sectioned parallel to the roll direction (dash
line) and normal to the roll direction (solid line) (de Groh et al., 2016).
TABLE 28.7
HST Thermal Control Candidate Replacement Materials
Material
Sample id.
Candidate material
1
B1.1 and M2.1
10 mil FEP/Ag/Inconel/adhesive/Nomex (polyphenylene isophthalate) scrim
2
B1.2 and M2.2
5 mil FEP/Ag/Inconel/adhesive/fiberglass scrim/adhesive/2 mil Kapton
3
B1.3 and M2.3
10 mil FEP/Al/adhesive/Nomex scrim
4
B1.4 and M2.4
5 mil FEP/Al/adhesive/fiberglass scrim/adhesive/2 mil Kapton
5
B1.5 and M2.5
5 mil FEP/Ag/Inconel/adhesive/Nomex scrim
6
B1.6 and M2.6
5 mil FEP/Al/adhesive/Nomex scrim
7
B1.7 and M2.7
OCLI multilayer oxide UV blocker/2 mil white Tedlar
8
B1.8 and M2.
5 mil FEP/Al (current HST material)
9
B1.9 and M2.9
SiO2/Al2O3/Ag/Al2O3/4 mil stainless steel
10
B1.0 and M2.0
Proprietary Teflon FEP/AZ93 White Paint/Kapton
One study examined candidate materials to replace the degrading outer layer of aluminized FEP on the HST.
Candidate materials that were considered are indicated in Table 28.7. Various sets of these candidate replacement
materials were exposed to combinations of electron/proton radiation, atomic oxygen, soft X-rays, thermal
cycling, and near-UV radiation at various facilities in order to evaluate their HST on-orbit durability (de Groh
et al., 1998; Townsend et al., 1999b). Two sets of samples (B1 and M2) previously exposed to charged-particle
radiation were exposed to soft X-rays, and one sample set (B1) was also thermal cycled under load. Thermal
cycling temperatures ranged between 2100 C to 150 C and spring loading provided stress on each sample of
approximately 12.4 MPa. Samples received 1,000 thermal cycles. Thermal cycled samples are shown in Fig. 28.28.
Metallized 5-mil-thick FEP samples B1.2 and B1.4, shown in Fig. 28.28A, with fiberglass scrims and Kapton substrates, tore in half during thermal cycling under load (de Groh et al., 1998). This may be attributed to the low
tear resistance of Kapton. These samples performed worse than Sample B1.8, shown in Fig. 28.28B, the HST MLI
material (5-mil-thick aluminized-FEP), which tore about 90% of the width during thermal cycling. Tear propagation of the B1 samples was attributed to thermal cycling under a high load. The prior radiation exposures did not
appear to have an additional effect on tearing, and no tearing occurred due to mechanical load cycling.
IV. CASE STUDIES
638
28. DEGRADATION OF SPACECRAFT MATERIALS
FIGURE 28.28 Candidate HST MLI under tension in the Rapid Thermal Cycling Facility after 1,000 thermal cycles. (A) samples
B1.1 2 B1.4 and (B) samples B1.5 2 B1.8.
Following the evaluation of all test results, 5-mil FEP/Al/adhesive/Nomex scrim was recommended as replacement material for the outer thermal blanket layer for the HST (Townsend et al., 1999b).
Ground-based environmental durability tests, such as those discussed in Townsend et al. (1998), indicate that
exposing materials in accelerated tests to environmental models predicted spacecraft mission exposures of UV
and/or ionizing radiation sources does not simulate the extent of damage that occurs in the space environment.
One approach to overcoming the difficulties in simulating the space environment using ground-based testing is
to calibrate the facility using data from actual space-exposed materials to determine the exposure levels required
to replicate the degraded properties observed in space. de Groh et al. (2003) describes a ground-to-space correlation method that uses a multistep process to determine the durability of expanded-polytetrafluoroethylene
(ePTFE) for International Space Station (ISS) applications, based on ground-based X-ray irradiation and heating
exposure that simulates bulk embrittlement as occurs in FEP thermal insulation covering the HST. This method
was designed to damage the back-surface of equivalent thickness ePTFE to the same amount of scission damage
as occurred in the HST FEP (based on elongation data), and then correct for differences in ground-test ionizing
radiation versus space-radiation effects, temperature variations, space ionizing radiation environment variations
(spacecraft altitude, inclination and duration), and thickness variations.
28.6 MICROMETEOROID AND ORBITAL DEBRIS EFFECTS
28.6.1 Environment Description
Micrometeoroids are of extraterrestrial origin and, as such, will have a flux that is reasonably constant with
time. Their velocity is typically in the 4 2 51 km/sec range, (Kinard and Martin, 1991; Zook, 1991) with an
average velocity near 20 km/sec. As shown in Fig. 28.29, as micrometeoroid particle size decreases, the flux of
particles increases. Orbital debris is manmade as a result of spent solid rocket booster exhaust, satellite breakups
or other man-caused origins. Orbital debris has an average velocity of 8.7 km/sec. Shipman and Williamson
(1997) Because of the manmade origin and atmospheric drag, orbital debris flux is highly dependent on the
IV. CASE STUDIES
639
28.6 MICROMETEOROID AND ORBITAL DEBRIS EFFECTS
1e+5
1e+4
Micrometeoroid flux model
Debris flux model (Kessler 1989)
Impact flux (impacts/yr-m2)
1e+3
1e+2
1e+1
1e+0
1e−1
1e−2
1e−3
1e−4
1e−5
1e−6
1e−7
1e−8
1e−6
FIGURE 28.29
1e−5
1e−4
1e−3 1e−2 1e−1 1e+0
Particle diameter (cm)
1e+1
1e+2
1e+3
Micrometeoroid and debris particle flux versus diameter for low Earth orbit (500 km).
world’s spacecraft launch frequency and occurrences of orbital breakups. The orbital debris size distribution can
be seen in Fig. 28.29. The Micrometeoroid Flux Model shown in Fig. 28.29 was developed based on the
CourPalais model, (Cour-Palais et al., 1969) and the Debris Flux model was developed based on the Kessler
model (Kessler et al., 1988). For a spacecraft with a fixed orientation relative to the ram velocity direction, the combined occurrence of micrometeoroid and orbital debris impacts is nonuniform around the spacecraft. This is the
case for the LDEF, which was in LEO for 69 months. Fig. 28.30 shows a polar plot of the number of impacts around
the LDEF (See et al., 1992). Row 9 was facing the ram direction (the leading edge), and row 3 was the trailing edge.
28.6.2 Interactions With Materials
The impact of micrometeoroid or orbital debris with spacecraft materials is usually sufficiently energetic to
cause vaporization of the impacting particle as well as produce an impact crater of volume an order of magnitude greater than the impacting particle. For example, the kinetic energy of an aluminum particle traveling at
6 km/sec is sufficient to vaporize aluminum to form a crater that is roughly five times the diameter of the incoming particle (Mandeville, 1992). Fig. 28.31 shows scanning electron microscope photos of typical hypervelocity
impact craters in aluminum and FEP Teflon from the LDEF spacecraft (Bunch et al., 1991). The violence of the
microscopic explosive vaporization can cause delamination in layered materials as shown in Fig. 28.32, which
shows a layered structure of FEP, silver, and Z306 black paint flown on the LDEF that was delaminated over a
diameter an order of magnitude greater than the impacting particle, followed by subsequent atomic oxygen oxidation of the underlying silver (Stein and Pippin, 1991). The ejection of impact crater material can be a source of
spacecraft self-contamination. Large particle impacts, although rare in occurrence, have the potential to penetrate
pressure vessels or cause structural damage. The largest impact crater on the LDEF spacecraft was 5.7 mm in
diameter (Zolensky et al., 1992).
28.6.3 Mitigation Techniques
Localized damage caused by micrometeoroid or orbital debris impacts on spacecraft surfaces can be missionthreatening if the impact occurs on electrical wires, power cables, or pressure vessels. Reduction in the probability
of catastrophic loss in an electrical power conductor can be achieved through use of a ladder-configured
conductor rather than a single conductor. The concept is that one would divide the current into two smaller conductors that are separated with occasional rung conductors that could carry current back and forth across the two
main long conductors in case one conductor was severed by hypervelocity impact. Only a short length of single
conductor between the rungs on either side of impact breakage site would have to carry twice the normal current.
IV. CASE STUDIES
Row 9
90
1.0
120
60
0.8
0.6
150
30
0.4
0.2
Row 6 180
0 Row 12
0
1.0
0.8
1.6
0.4
0.2
0
0.2
0.4
0.6
0.8
1.0
0.2
0.4
210
330
0.6
0.8
240
1.0
300
270
Row 3
FIGURE 28.30
Angular dependence of impacts around the LDEF spacecraft.
FIGURE 28.31
Scanning electron microscope photos of impact craters on the LDEF spacecraft: (A) in aluminum; (B) in FEP Teflon.
28.7 CONCLUDING REMARKS
FIGURE 28.32
641
Photo of impact site on FEP Teflon that was backed with silver and Z 306 paint.
Mitigation techniques for power cables can simply be the use of redundant separated cables so that the probability of all cables failing is acceptably small. The same technique is applicable to electrical signal conductors.
Such techniques are simple to accomplish in flat-ribbon conductors that are laminated with polymer insulation.
Because one micrometeoroid or debris impact can cause total loss in pressure of a tank or container that can
fail a mission or cause loss of life for manned missions, mitigation techniques are important. The technique used
to prevent penetration of pressure vessels has been to add an outer thin wall, often called a “Whipple shield,” to
cause the impacting particle to break up upon impact (Schwinghamer and Whitaker, 1993; Christiansen, 2003).
The debris from the impact then spreads over a large area on the surface of the pressure vessel, which is spaced
concentrically inside the shield material. Because the debris is spread over a large area, its ability to penetrate the
inner critical chamber is greatly reduced. Variants of this concept for structured protection of tubular structures
include filling the space between two sides of a tubular structure with low-density ceramic fiber fill, which also
acts to absorb the energy of the broken up primary impact debris (Christiansen, 2003). Damage to the back wall
of the structure is prevented by the energy-absorbing fill material.
28.7 CONCLUDING REMARKS
Spacecraft materials exposed to LEO environments have been found to undergo degradation or damage due
to environmental threats including atomic oxygen, contamination, radiation, temperature effects and temperature
cycling, and micrometeoroids and orbital debris. The degree to which the space environment degrades or
damages materials depends on the unique conditions of an individual spacecraft environment and the susceptibility of the material to being altered by these environmental exposures. As has been shown in this chapter, space
environment interactions with materials are complex, often producing combined or synergistic effects. Accurately
predicting the behavior of materials in these complex space environments is important to the success of space
missions, but can be difficult to accomplish.
Since the first Space Shuttle flights, significant advancements have been made in understanding atomic oxygen
effects on materials, namely: (1) in quantifying atomic oxygen degradation for many polymers; (2) in development of coatings to prevent atomic oxygen degradation; (3) in development of calibrated laboratory simulation
methods; and (4) in development of modeling tools to predict in-space degradation. However, for radiation
effects and synergistic effects with temperature, accelerated testing in the laboratory using individual or
combined environments to estimated mission exposure fluences or doses has been generally unable to accurately
replicate damage observed in space. Furthermore, complex differences between the space and laboratory environments require cautious interpretation of results. These complications of predicting durability of spacecraft materials have led to efforts in developing ground-to-space correlation methods for radiation durability testing (e.g.,
exposing Teflon materials to exposures based on the dose of radiation that produces reduction in FEP mechanical
properties equivalent to that observed in space, rather than to actual mission fluence exposures).
As space missions are becoming longer in duration and require lighter-weight materials with specially tailored
properties, it is critical that advancements continue in understanding material degradation mechanisms and in
the development of simulation laboratory durability testing and modeling capabilities. This is crucially dependent on the retrieval and subsequent testing of long duration space-exposed materials, such as the MISSE flight
experiments. These advancements will ensure accuracy in predicting long-term material durability in the space
environment for future missions.
IV. CASE STUDIES
642
28. DEGRADATION OF SPACECRAFT MATERIALS
References
Adams, M.R., 1993. The Degradation of Polymeric Spacecraft Materials by Far-UV Radiation and Atomic Oxygen. UMI Dissertation Services,
Ann Arbor, MI, p. 138.
American Society for Testing and Materials. Solar constant and air mass zero solar spectral irradiance tables. ASTM-E 490-73a; (Reapproved
1992).
Banks B., Stueber T., Norris M. Monte carlo computational modeling of the energy dependence of atomic oxygen undercutting of protected
polymers. In: 4th International Space Conference ICPMSE-4. Toronto, Canada; April 23 2 24, 1998. NASA TM-1998- 207423.
Banks, B., Miller, S., de Groh, K., Chan, A., Sahota, M. The development of surface roughness and implications for cellular attachment in
biomedical applications. In: Materials Research Society 2001 Fall Meeting. Boston, MA; November 2630, 2001. NASA TM- 2001-211288.
Banks, B.A., Rutledge, S.K. Low earth orbital atomic oxygen simulation for materials durability evaluation. In: Proceedings of the 4th
European Symposium on Spacecraft Materials in Space Environment. Toulouse, France; September 6 2 9, 1988. pp. 372 2 92.
Banks, B.A., Stueber, T.J., 1997. Monte Carlo computational techniques for prediction of atomic oxygen erosion of materialsWroclaw, Poland,
May 20 2 23 In: Tennyson, R.C., Kiv, A.E. (Eds.), Proceedings of the NATO Advanced Research Workshop on Computer Modeling of
Electronic and Atomic Processes in Solids, 1996. Kluwer Academic Publishing, The Netherlands.
Banks B.A., Mirtich M.J., Rutledge S.K., Swec, D.M. Sputtered coatings for protection of spacecraft polymers. In: 11th International Conference
on Metallurgical Coatings (AVS). San Diego, CA; April 9 2 13, 1984. NASA TM-83706.
Banks B.A., Mirtich M.J., Rutledge S.K., Swec D.M. Nahra H.K. Ion beam sputter-deposited coatings for protection of spacecraft polymers in
low earth orbit. In: 23rd Aerospace Sciences Meeting (AIAA). Reno, NV; January 14 2 17, 1985. NASA TM-87051.
Banks B.A., Auer B.M., Rutledge S.K., Gebauer L., Sechkar E.A. Monte Carlo modeling of atomic oxygen interaction with protected polymers
for projection of materials durability in low earth orbit. In: Proceedings of the MRS 1992 Spring Meeting. San Francisco, CA; April
27 2 May 1, 1992.
Banks B.A., de Groh K., Rutledge S., DiFilippo F. Prediction of in-space durability of protected polymers based on ground laboratory thermal
energy atomic oxygen. In: Presented at the 3rd International Conference for Protection of Materials and Structures From Low Earth Orbit
Space Environment, University of Toronto Institute for Aerospace Studies. Toronto, Canada; April 25 2 26, 1996. NASA TM-107209.
Banks, B.A., de Groh, K.K., Rutledge, S.K., DiFilippo, F.J., 1999. Prediction of in-space durability of protected polymers based on ground
laboratory thermal energy atomic oxygen. In: Kleiman, J.I., Tennyson, R.C. (Eds.), Protection of Materials and Structures Form the Low
Earth Orbit Space Environment. Kluwer Academic Publishers, The Netherlands, pp. 89100.
Banks B.A., de Groh K.K., Rutledge S.K., Haytas C.A. Consequences of atomic oxygen interaction with silicone and silicone contamination on
surfaces in low earth orbit. In: 44th Annual Meeting, Sponsored by the International Society for Optical Engineering. Denver, CO; July 21,
1999. NASA TM-1999-209179.
Banks, B.A., Rutledge, S.K., Auer, B.M., DiFilippo, F.J., 1990. Atomic oxygen undercutting of defects on SiO2 protected polyimide solar array
blankets. In: Srinivasan, V., Banks, B. (Eds.), Materials Degradation in Low Earth Orbit. TMS, pp. 1533.
Banks, B.A., Rutledge, S.K., de Groh, K.K., 1991. Low earth orbital atomic oxygen, micrometeoroid, and debris interactions with photovoltaic
arrays. Presented at the 11th Space Photovoltaic Research and Technical Conference (SPRAT XI). NASA Lewis Research Center, Cleveland,
OH.
Banks, B.A., Rutledge, S.K., de Groh, K.K., Mirtich, M.M., Gebauer, L., Olle, R., et al. The implications of the LDEF results on space station
freedom power system materials. In: 5th International Symposium on Materials in a Space Environment. Cannes-Mandelieu, France;
September 16-0, 1991.
Banks B.A., Snyder A., Miller S.K., Demko R. Issues and consequences of atomic oxygen undercutting of protected polymers in low earth
orbit. In: 6th International Conference on Protection of Materials and Structures From Space Environment. Toronto, Canada; May 1 2 3,
2002. NASA TM-2002-211577.
Banks B.A., Miller S.K.R., de Groh K.K., Demko R. Scattered atomic oxygen effects on spacecraft materials. In: Proceedings of the 9th international symposium on materials in a space environment. Noordwijk, The Netherlands; June 16 2 20, 2003. pp. 145 2 52. ESA SP-540.
Barth J.L. Space and atmospheric environments: from low earth orbits to deep space. In: Proceedings of the 9th international symposium on
materials in a space environment, European Space Agency, ESA SP-540; September 2003. pp. 17 2 30.
Brown T.L. Microcracking of materials for space. Ph.D. Dissertation, Virginia Polytechnic Institute and State University; October 1998. Also,
Final Report for Grant NAG-1-1912, NASA Langley Research Center, October 1998.
Bunch T.E., diBrozolo F.R., Fleming R.H., Harris D.W., Brownlee D., Reilly T.W. LDEF impact craters formed by carbon-rich impactors: a preliminary report. In: Proceedings of the LDEF—69 months in space first post-retrieval symposium. Kissimmee, FL; June 2 2 8, 1991. NASA
CP-3134 Part 1. pp. 549 2 64.
Caledonia, G.E., 1989. Laboratory simulations of energetic atom interactions occuring in low earth orbit. In: Munts, E.O., Weaver, D.P., Campbell,
D.H. (Eds.), Rarefied Gas Dynamics: Space Related Studies, vol. 116. Progress in Astronautics and Aeronautics, AIAA, pp. 129142.
Chen P.T. Contamination effects due to space environmental interactions. In: Am Institute Aeronautics Astronautics January 2001; AIAA
Paper 2001-0095.
Christiansen, E.L., 2003. Meteoroid/Debris Shielding. NASA TP-2003-210788.
Cour-Palais, B.G., et al., 1969. Meteoroid Environment Model—1969 (Near Earth to Lunar Surface). NASA SP-8013.
Daneman, S.A., Babel, H.W., Hasegawa, M.M. Selection rationale, application, optical properties, and life verification of Z-93 for the space station. McDonnell Douglas Aerospace, Report 94H0632; August 1994.
de Groh, K.K., 2010. Materials Spaceflight Experiments. In: Blockley, R., Shyy, W. (Eds.), Encyclopedia of Aerospace Engineering. John Wiley
& Sons Ltd, Chichester, UK, pp. 25352552.
de Groh, K.K., Banks, B.A., August 1994. Atomic oxygen undercutting of long duration exposure facility aluminized-kapton multilayer insulation. J. Spacecraft Rockets (AIAA) 31 (4), 656664.
de Groh, K., McCollum, T., 1995. Low earth orbital durability of protected silicones for refractive photovoltaic concentrator arrays.
J. Spacecraft Rockets 32 (1).
de Groh, K.K., Dever, T.M., Quinn, W.F., April 1990. The Effect of Leveling Coatings on the Atomic Oxygen Durability of Solar Concentrator
Surfaces. NASA TM 102557.
IV. CASE STUDIES
REFERENCES
643
de Groh, K.K., Banks, B.A., Smith, D.C., Environmental durability issues for solar power systems in low earth orbit. solar engineering 1995,
vol. 2; March 1995. pp. 939 2 950. NASA TM 106775.
de Groh, K.K., Banks, B.A., Sechkar, E.A., Scheiman, D.A. Simulated solar flare X-Ray and thermal cycling durability evaluation of hubble
space telescope thermal control candidate replacement materials. In: Presented at the 4th ICPMSE Conference. Toronto, Canada; April
23 2 24, 1998. NASA TM-1998-207426. December 1998.
de Groh, K.K., Gummow, J.D., September 2001. Effect of air and vacuum on the tensile properties of X-ray Exposed aluminized-FEP. HPP 13
(3), S421S431.
de Groh, K.K., Martin, M., July 2002. The effect of heating on the degradation of ground laboratory and space irradiated teflon FEP. National
Aeronautics and Space Administration, TM-2002-211704.
de Groh, K.K., Dever, J.A., Sutter, J.K., Gaier, J.R., Gummow, J.D., Scheiman, D.A., et al., 2001. Thermal Contributions to the Degradation of
Teflon FEP on the Hubble Space Telescope. High Performance Polymers 13, S401S420.
de Groh K.K., Banks B.A., Dever J.A. Hodermarsky J.C. Use of hubble space telescope degradation data for ground-based durability projection
of ePTFE on ISS. In: Proceedings of the 9th international symposium on materials in a space environment. Noordwijk, The Netherlands;
June 16 2 20, 2003. pp. 281 2 90. ESA SP-540, September 2003.
de Groh, K.K., Snyder, A., Finlay, K.A., 2008. Degradation of Hubble Space Telescope Aluminized-Teflon Bi-Stem Thermal Shields. High
Performance Polymers 20, 410428.
de Groh, K.K., Perry, B.A., Banks, B.A., 2016. Effect of 1.5 Years of Space Exposure on Tensile Properties of Teflon. Journal of Spacecraft and
Rockets 53 (6), 10021011. November-December.
de Groh, K.K., Banks, B.A., Asmar, O.C., Yi, G.T., Mitchell, G.G., Guo, A. and Sechkar, E.A., February 2017. Erosion Results of the MISSE 8
Polymers Experiment After 2 Years of Space Exposure on the International Space Station. NASA-TM-2017-219445 (February 2017).
de Groh, K.K., Banks, B.B., Mitchell, G.G., Yi, G.T., Guo, A., Ashmead, C.C., Roberts, L.M., McCarthy, C.E., Sechkar, E.A., 2013. MISSE 6
stressed polymers experiment atomic oxygen erosion data. In: Proceedings of the “12th International Symposium on Materials in the Space
Environment (ISMSE 12).” Noordwijk, The Netherlands (ESA SP-705, February 2013); also NASA TM-2013-217847.
de Groh, K. K., Banks, B. A. and McCarthy, C. E. NASA Technical Handbook, Spacecraft Polymers Atomic Oxygen Durability Handbook
(NASA-HDBK-6024), June 2014.
de Groh, K.K., Banks, B.A., Yi, G.T., Haloua, A., Imka, E.C., Mitchell, G.G., Asmar, O.C., Leneghan, H.A., Sechkar, E.A., December 2016.
“Erosion Results of the MISSE 7 Polymers Experiment and Zenith Polymers Experiment After 1.5 Years of Space Exposure.” NASA-TM2016-219167.
Dever, J., de Groh, K., February 2002. Vacuum ultraviolet radiation and atomic oxygen durability evaluation of HST Bi-Stem thermal shield
materials. NASA/TM-2002-211364.
Dever, J., Semmel, C., Edwards, D., Messer, R., Peters, W., Carter, A., et al., April 2002. Radiation durability of candidate polymer films for the
next generation space telescope sunshield. AIAA 2002 2 1564.
Dever, J.A., de Groh, K.K., Banks, B.A., Townsend, J.A., Barth, J., Thomson, L., et al., March 2000. Environmental exposure conditions for teflon FEP on the hubble space telescope. High Perform Polym 12 (1), 125139.
Dever J.A., McCracken C.A. Effects of various wavelength ranges of vacuum ultraviolet radiation on teflon FEP Film. In: Proceedings of the
9th international symposium on materials in a space environment. European Space Agency; September 2003. pp. 367 2 73. ESA SP-540.
Dever J.A., Rutledge S.K., Hambourger P.D., Bruckner E., Ferrante R., Pal A.M., et al. Indium tin oxide-magnesium fluoride co-deposited films
for spacecraft applications. International conference on metallurgical coatings (AVS). San Diego, CA; April 24 2 26, 1996. NASA TM-1988208499.
Dever, J.A., Messer, R., Powers, C., Townsend, J., Wooldridge, E., September 2001. Effects of vacuum ultraviolet radiation on thin polyimide
films. High Perform Polym 13 (3), S391S399.
Dever, J.A., Pietromica, A.J., Stueber, T.J., Sechkar, E.A., Messer, R.K., January 2001. Simulated space vacuum ultraviolet (VUV) exposure testing for polymer films. American Institute of Aeronautics and Astronautics, AIAA Paper 2001-1054.
Dever, J.A., Miller, S.K., Sechkar, E.A., Wittberg, T.N., October 2008. Space environment exposure of polymer films on the materials international space station experiment: results from MISSE 1 and MISSE 2. High Perform. Polym. 20 (4/5), 371387.
Dever, J.A., Waters, D.L., de Groh, K.K., Nguyen, Q.-V., 2010. Evaluation of optical properties and atomic oxygen erosion yields of polymer
film materials exposed to the space environment on MISSE 3 & 4, In: Proceedings of the 2010 National Space and Missile Materials
Symposium (NSMMS), Scottsdale, AZ.
Dickerson, R.E., Gray, H.B., Haight, G.P., 1979. Chemical Principles, 3rd ed Benjamin Cummings Publishing Co. Inc., Menlo Park, CA, p. 457.
DuPont, Technical Information on Tedlar (Polyvinyl fluoride film). ,http://www.dupont.com/tedlar/. ; 2003.
Gatto, A., Escoubas, L., Roche, P., Comandre, M., 1998. Simulation of the degradation of optical glass substrates caused by UV irradiation
while coating. Optics Communications 148, 347354.
Gonzalez R.I., Tomczak S.J., Minton T.K., Brunsvold A.L., Hoflund, G.B. Synthesis and atomic oxygen erosion testing of space-survivable
POSS (Polyhedral Oligomeric Silsesquioxane) polyimides. In: Proceedings of the 9th international symposium on materials in a space environment. Noordwijk, The Netherlands; June 16 2 20, 2003. pp. 113 2 20. ESA SP-540.
Goode D.C., Williams A.W., Wood N.J., Binzakaria A. Photothermal imaging of gold and vermiculite coated kapton exposed to atomic oxygen.
In: ESA Proceedings of the 6th international symposium on materials in a space environment; Nov. 1994. pp. 201 2 6. SEE N95-27568 09-23.
Grossman E., Noter Y., Lifshitz Y. Oxygen and VUV irradiation of polymers: atomic force microscopy (AFM) and complimentary studies.
Proceedings of the 7th International Symposium on Materials in Space Environment, Toulouse, France, June 16 2 20, 1997. Noordwijk,
Netherlands: ESA; 1997. pp. 217 2 22.
Grossman E., Noter Y., Lifshitz Y. Oxygen and VUV irradiation of polymers—atomic force microscopy (AFM) and complementary studies.
Proceedings of the 7th International Symposium on Materials in Space Environment, Toulouse, France, June 16 2 20, 1997. Noordwijk,
Netherlands: ESA; 1997. pp. 217 2 23.
Harada Y., Deshpande M. Requalification of White Thermal Control Coatings. Wright Laboratory Report WL-TR-94-4126; October 1994.
Harrison S.E. A study of gamma ray photoconductivity in organic dielectric materials. Sandia Corporation, Report SCDC-2580; 1962.
IV. CASE STUDIES
644
28. DEGRADATION OF SPACECRAFT MATERIALS
Hastings, D., Garrett, H., 1996. Spacecraft-Environment interactions. Cambridge University Press, New York.
Hedin, A.E., 1987. MSIS-86 thermospheric model. J Geophys Res 92 (4), 649.
Holmes-Siedle, A., Adams, L., 2002. Handbook of radiation effects. Oxford University Press, New York, pp. 568570.
Illingsworth, M.L., Betancourt, L., He, L., Chen, Y., Terschak, J.A., Banks, B.A., et al., 2001. Zr- Containing 4,4’-ODA/PMDA polyimide composites. NASA TM-2001-211099.
Iskanderova, Z.A., Kleiman, J.I., Gedimenko, Y., Morrison, W.D., Tennyson, R.C., 1999. Surface modification of polymer-based materials by
ion implantation—a new approach for pro- tection in LEO. In: Kleiman, J.I., Tennyson, R.C. (Eds.), Protection of materials and structures
for the low earth orbit space environment. Kluwer Academic Publishers, The Netherlands, pp. 225234.
Jaworske, D.A., de Groh, K.K., Podojil, G., McCollum, T., Anzic, J., June 1994. Leveling coatings for reducing the atomic oxygen defect density
in protected graphite fiber epoxy composites. J IES (Inst Environ Sci) XXXVII (3), 2631.
Jukeikis H.S. Space radiation effects on teflon films. Report SAMSO-TR-79-070, The Aerospace Corporation; September 1979.
Kessler, D.J., Reynolds, R.C., Anz-Meador, P.D., 1988. Orbital debris environment for spacecraft designed to operate in low earth orbit. NASA
TM-100471.
Kinard W.H., Martin G.D. Long duration exposure facility (LDEF) space environments overview. In: Proceedings of the LDEF—69 months in
space first post-retrieval sym- posium. Kissimmee, FL; June 2 2 8, 1991. NASA CP-3134 Part 1. pp. 49 2 60.
Koller, L.R., 1965. Ultraviolet radiation. John Wiley & Sons, Inc, NewYork.
Koontz S.L., Golden J.L., Lorenz M.J., Pedley M.D. Ionizing radiation effects on ISS ePTFE jacketed cable assembly. In: Proceedings of the 9th
international symposium on materials in a space environment, Noordwijk, The Netherlands; June 16 2 20, 2003. ESA SP-540, September
2003. pp. 297 2 301.
Laghari, J.R., Hammoud, A.N.A., April 1990. Brief survey of radiation effects on polymer dielectrics. IEEE Trans Nucl Sci 37 (2), 1,0761,083.
Leger, L.J., 1983. Oxygen atom reaction with shuttle materials at orbital altitudes—data and experiment status. AIAA-83-0073.
Liau, W.B., Tseng, F.P., 1998. Effect of Long-term ultraviolet light irradiation on polymer matrix composites. Polym Compos 19 (4), 440445.
Luoma, G.A., Rowland, R.D., April 1989. The effects of sunlight and humidity on CF-188 composite resin. Defence research establishment
pacific technical memorandum 89-07.
Mandeville J.-C. Micrometeoroids, debris on LDEF. In: Proceedings of the LDEF—69 months in space second post-retrieval symposium. San
Diego, CA; June 1 5, 1992. NASA CP-3194 Part 2. pp. 303 2 12.
Mata, A., Su, X., Fleischman, A., Banks, B., Miller, S., Midura, R., December 2003. Osteoblast attachment to a textured surface in the absence of
exogenous adhesion proteins. IEEE Trans Nanobiosci 2 (4).
McKellar, J.F., Allen, N.S., 1979. Photochemistry of man-made polymers. Applied Science Publishers, Ltd, London, p. 31.
Mende, S.B., Swenson, G.R., Clifton, K.S., 1984. Science 225, 191.
Meshishnek M.J., Hemminger C.S., Gyetvay S.R. Space environmental exposure of IITRI white thermal control paints. Aerospace corporation,
Report TR-95(5904)-2; April 27, 1995.
N.O.A.A., N.A.S.A. and U.S.A.F. U.S. Standard Atmosphere, 1976. NASA Tech. Memo TMX-74335; 1976.
Noter Y., Grossman E., Genkin L., Haruvy Y., Murat M., Ross I., et al. Variations in the telemetry of an offeq satellite’s sun sensors. Presented
at the 3rd international space forum on protection of materials and structures from low earth orbit space environment, University of
Toronto institute for aerospace studies. Toronto, Canada; April 25 2 26, 1996.
Reed, R.P., Schramm, R.E., Clark, A.F., February 1973. Mechanical, thermal, and electrical properties of selected polymers. Cryogenics 13,
6782.
Rice, C.J., Russell, R.W., 1992. Infrared spectral measurement of space shuttle glow. Geophys Res Lett 22 19 (10), 989992.
Russell, D.A., Connell, J.W., Fogdall, L.B., 2002. Electron, proton, and ultraviolet radiation effects on thermophyisical properties of polymeric
films. J Spacecraft Rockets 39 (6), 833838.
Rutledge S.K., Cooper J.M., Olle R.M. The effect of atomic oxygen on polysiloxane-polyimide for spacecraft applications in low earth orbit. In:
Proceedings of the 4th annual workshop on space operations applications and research (SOAR ‘90). Albuquerque, NM; June 26 2 28, 1990.
vol. II, pp. 755 2 62. NASA CP 3103.
Rutledge, S.K., Mihelcic, J.A., 1990. The effect of atomic oxygen on altered and coated kapton surfaces for spacecraft applications in low earth
orbit. In: Srinivasan, V., Banks, B. (Eds.), Materials degradation in low earth orbit. TMS, pp. 3548.
Rutledge S.K., Olle R.M. Space station freedom solar array blanket coverlay atomic oxygen durability testing results. In: Proceedings of the
38th international SAMPE sympo- sium; May 10 2 13, 1993. pp. 679 2 93.
Sawyer, D.M., Vette, J.I., December 1976. AP-8 trapped proton environment for solar maximum and solar minimum. NASA/Goddard Space
Flight Center, Greenbelt, MD, NSSDC/ WDC-A-R&S, 76-06.
Schnabel, W., 1981. Polymer degradation: principles and practical applications. Macmillan Publishing Co., Inc, New York, Chapter 5.
Schwinghamer, R.J., Whitaker, A., 1993. ISSN 0145-319X/December 1 Shield design for protection against hypervelocity par- ticles. NASA
Tech Briefs, p. 76.
See T.H., Mack K.S., Warren J.L., Zolensky M.E., Zook H.A. Continued investigation of LDEF’s structural frame and thermal blankets by the
meteoroid and debris special investigation group. In: Proceedings of the LDEF—69 months in space second post-retrieval symposium. San
Diego, CA; June 1 2 5, 1992. NASA CP-3194 Part 2. pp. 313 2 24.
Seki, K., Tanaka, H., Ohta, T., 1990. Electronic structure of poly(tetrafluoroethylene) studied by UPS, VUV absorption, and band calculations.
Physica Scripta 41, 167171.
Shepp, A., Haghighat, R., Lennhoff, J., Schuler, P., Connell St., J., Clair, T., et al., 1999. TOR and COR AO-VUV resistant polymers for space.
In: Kleiman, J.I., Tennyson, R.C. (Eds.), Protection of materials and structures for the low earth orbit space environment. Kluwer Academic
Publishers, The Netherlands, pp. 235254.
Shipman J., Williamson J. AIAA survivability technical committee Draft. Aerospace Am December 1997.
Silverman E.M. Space environmental effects on spacecraft: LEO materials selection guide. Prepared for Langley Research Center under
Contract NAS1-19291. NASA Contractor Report 4661 Part1; August 1995a.
Silverman E.M. Space environmental effects on spacecraft: LEO materials selection guide. NASA Contractor Report 4661, Part 2; August,
1995b. Sections 10 2 38.
IV. CASE STUDIES
FURTHER READING
645
Skurat V.E., Barbashev E.A., Budashov I.A., Dorofeev YuI, Nikiforov A.P., Ternovoy A.I., et al. The separate and combined effects of VUV
Radiation and fast atomic oxygen on teflon FEP and silicon carbide. Proceedings of the 7th international symposium on materials in space
environment, Toulouse, France, June 16 2 20, 1997. Noordwijk, Netherlands: ESA; 1997. pp. 267 2 79.
Stein B.A., Pippin H.G. Preliminary findings of the LDEF materials special investigation group. In: Proceedings of the LDEF—69 months in
space first post-retrieval sympo- sium. Kissimmee, FL; June 2 2 8, 1991. NASA CP-3134 Part 2. pp. 617 2 41.
Stuckey W.K., Meshishnek M.J. Space environmental stability of tedlar with multi-layer coatings: space simulation testing results. Aerospace
Corp., Report AD-A387139; TR-2000(85665)-8; SMC-TR-00-03 under Contract F04701-93-C-0094. El Segundo, CA; Aug. 20, 2000.
Stuckey, W.K., Meshishnek, M.J., 2001. Solar ultraviolet and space radiation effects on inflatable materials. In: Jenkins, C.H.M. (Ed.), Gossamer
spacecraft: membrane and inflatable structures technology for space applications, vol. 191. American Institute of Aeronautics and
Astronautics, Reston, VA, pp. 303320. , Progress in Astronautics and Astronautics.
Toupikov V.I., Khatipov S.A., Charniavsky A.I., Stepanov V.F. Degradation of mechanical and electrophysical properties of teflon FEP films
under combined action of far ultra- violet and thermal cycling. Proceedings of the 7th International Symposium on Materials in Space
Environment, Toulouse, France, June 16 2 20, 1997. Noordwijk, Netherlands: ESA; 1997. pp. 77 2 85.
Townsend J., Powers C., Viens M., Ayres-Treusdell M., Munoz B. Degradation of teflon FEP following charged particle radiation and rapid
thermal cycling. Proceedings of the 20th space simulation conference: the changing testing paradigm. National Aeronautics and Space
Administration; October 1998 CR-1998-208598. pp. 201 2 9.
Townsend, J.A., Hansen, P.A., Dever, J.A., de Groh, K.K., Banks, B.A., Wang, L., et al., March 1999a. Hubble space telescope metallized teflon
FEP thermal control materials: on-orbit degradation and post-retrieval analysis. High Perform Polym 11 (1), 81100.
Townsend, J.A., Hansen, P.A., McClendon, M.W., de Groh, K.K., Banks, B.A., 1999b. Ground-based testing of replacement thermal control
materials for the hubble space telescope, special issue. High Perform Polym 11, 6379.
Tribble, A.C., 1995. The space environment: implications for spacecraft design. Princeton University Press, Princeton, NJ.
Tribble A.C., Boyadjian B., Davis J., Haffner J., McCullough E. Contamination control engineering design guidelines for the aerospace community. National Aeronautics and Space Administration, NASA Contractor Report 4740; May 1996.
Vette, J.I., November 1991. The AE-8 trapped electron model environment. NASA/ Goddard Space Flight Center, Greenbelt, MD, NSSDC/
WDC-A-R&S 91-24.
Walter, N.A., Scialdone, J.J., June 1997. Outgassing data for selecting spacecraft materials. National Aeronautics and Space Administration.
NASA Reference Publication 1124, Rev. 4.
Waters, D. L., de Groh, K. K., Banks, B. A. and Cameron, K. C., 2009. Changes in optical and thermal properties of the MISSE 2 PEACE polymers and spacecraft silicones. In: Proceedings of the IISMSE-11, September 15-18, 2009, Aix-en-Provence, France.
Website of the National Oceanic and Atmospheric Administration’s Space Environment Center, Primer on Space Weather. ,http://www.sec.
noaa.gov/primer/primer.html..
Website of the National Oceanic and Atmospheric Administration’s National Geophysical Data Center, Space Physics Interactive Data
Resource. ,http://spidr.ngdc.noaa.gov/spidr..
Zolensky M.E., Zook H.A., Horz F., Atkinson D.R., Coombs C.R., Watts A.J., et al. Interim report of the meteoroid and debris special investigation group. In: Proceedings of the LDEF—69 months in space second post-retrieval symposium. San Diego, CA; June 1 2 5, 1992. NASA
CP-3194 Part 2. pp. 277 2 302.
Zook H.A. Deriving the velocity distribution of meteoroids from the measured meteoroid impact directionality on the various LDEF surfaces.
In: Proceedings of the LDEF—69 months in space first post-retrieval symposium. Kissimmee, FL; June 2 2 8, 1991, NASA CP-3134 Part 1.
pp. 569 2 79.
Further Reading
Banks, B.A., 1997. The use of fluoropolymers in space applications. In: Scheirs, J. (Ed.), Modern fluoropolymers. John Wiley & Sons, Ltd, New
York, pp. 103113.
Banks B.A., Rutledge, S.K., Paulsen P.E., Steuber T.J. Simulation of the low earth orbital atomic oxygen interaction with materials by means of
an oxygen ion beam. Presented at the 18th annual symposium on applied vacuum science and techcnology. Clearwater Beach, FL;
February 6 2 8, 1989. NASA TM 101971, 1989.
de Groh K.K., Banks B.A., Hammerstrom A.M., Youngstrom E.E., Kaminski C., Marx L.M., et al. MISSE PEACE polymers: an international
space station environmental exposure experiment. In: Proceedings of the conference on ISS utilization—2001. Cape Canaveral, FL; October
15 2 18, 2001. AIAA #2001-4923; also NASA TM-2001- 211311, November 2001.
Integrity testing laboratory, inc. Prediction of erosion of polymer-based materials by atomic oxygen in LEO. Final Report, GRC Contract #C72917-G; 1998.
Kleiman J., Iskanderova Z., Banks B.A., de Groh K.K., Sechkar E.A. In: Proceedings of the 8th International Symposium on Materials in a
Space Environment. Arcachon, France; June 5 2 9, 2000.
IV. CASE STUDIES
Progress Report
Spectral Characterization of Novel Spacecraft Materials at LEO Environment
Elena A Plis, Georgia Tech Research Institute
Milestone – 002
11/27/2020
1
Progress Report
I. Project Summary
We propose to fly several novel and well-understood materials on the MISSE-FF and monitor the changes
in spectral reflectivity that occur as a result of exposure to different components of the low Earth orbit
(LEO) environment. Energy deposition from the environment leads to chemical changes in the material
that in turn alter the optical properties. The same chemical damage that manifests as changes in optical
reflectance and absorptance also leads to changes in myriad physical properties such as mechanical
strength, electrical conductivity, and chemical reactivity. Correlation of the changes in each of these
properties as function of radiation type, flux and fluence will allow us to infer a host of material properties
based on one experimentally convenient measurement: color change. By flying identical test fixtures on
the ram, zenith, and wake positions of the MISSE-FF and collecting spectrally resolved images of the
materials throughout the mission along with concomitant measurement of the ambient space
environment, we may measure the optical changes that occur in each material as a function of exposure
to neutral atomic oxygen (AO), unfiltered solar ultraviolet (UV) radiation, and electrons, respectively.
These time-resolved data will function as a ground-truth reference for our team’s ongoing laboratorybased space weather-material interaction experiments. Correlation of the MISSE data with extensive
testing of duplicate samples under simulated space weather conditions will enable development of
fundamental chemical models for material degradation. This knowledge will be of use to material
scientists in academia and industry for understanding and developing novel spacecraft materials.
Aerospace engineers also will benefit from the insight into material behavior changes throughout a
mission lifetime. Further, a detailed predictive knowledge of space weather-induced material color
change will enable more robust and accurate space domain awareness by allowing remote observers to
glean knowledge about a spacecraft by examining the spectral signature of unresolved images. Use of
spacecraft optical properties in this way will increase space domain awareness and enable remote
diagnosis and anomaly resolution.
The specific objectives of the proposed project include (1) measurement of changes in spectral reflectivity
for a variety of novel and heritage spacecraft materials as a function of LEO environmental exposure; (2)
thorough physical and chemical characterization of duplicate materials before and after ground-based
simulated space weather experiments; (3) establishment of correlation factors between true spaceexposure and simulated (accelerated) space weather experiments at ground facilities, enabling accurate
prediction of on-orbit material performance based on laboratory-based testing; and (4) development of
techniques for remote spacecraft/debris diagnosis.
Expected project outcomes include (1) vetting of newly developed materials for aerospace/defense
applications under the LEO environment together with the development of accurate models describing
material weathering on LEO; (2) insight into the development of new materials with improved AO- and
UV-resistance; (3) more accurate material property inputs to be provided to the NASA/Air Force
Spacecraft Charging Analyzer Program (NASCAP), NASA’s Orbital Debris Spacecraft Materials Spectral
Database (ODSMSD), and the NASA Orbital Debris Engineering Model (OrDEM); (4) improvement of
characterization of resident space objects (RSOs) by measuring material- and space weather exposuredependent spectral changes. Knowledge of the evolution of a material’s spectral response across several
important ground-based astronomical observation bands, (U-I, 400-950 nm) will enable observers to
extract more information from remote color photometry data.
2
II. Estimated Flight Readiness
A. Estimated flight hardware readiness or target launch vehicle
Mark Shumbera, Director of the Business Development at the Alpha Space (AS), has communicated to us
on August 2020 that there is a delay with current and subsequent MISSE launces. The current MISSE
mission (MISSE-14) is delayed till the first quarter of 2021. Then next planned launches are:
MISSE-15 launch in September/October 2021; experiment h/w needed at AS by about April 2021
MISSE-16 launch in March/April 2022; experiment h/w needed at AS by about October/November 2021
MISSE-17 launch in August/Sept/Oct 2022; experiment h/w needed at AS by about March/April 2022
Considering several factors, in particular, possible slowdowns due to current pandemic and
November/December 2020 holidays, the MISSE-16 launch was chosen as the most realistic goal to launch
our experiment. This additional time will allow us to upgrade the hardware of the experiment, in order to
incorporate better light source and camera at MISSE Flight-Facility for the improved data acquisition
procedure. We want to allocate enough time for all of these hardware upgrades and come up not with
just good but with the best experiment possible. New target launch dates (March/April 2022) were
approved by AS. The ground-based test schedule is not affected, since we will independently characterize
pristine and even flight-simulated materials in according to our own timeline and resources.
III. Project Milestones and Schedule Updates
The planned major project milestones are, as stated in USER AGREEMENT #UA-2020-1116:
Table 1. Milestone Schedule
Milestone Schedule
Milestone
#
001
002
003
004
005
006
Milestone Description
Due Date
Status
Authorization to proceed (“ATP”)
8/31/2020
Progress report 1: Provide project status report detailing
preflight activities completed to date, baselined schedule to
11/27/2020
flight readiness, baselined payload turnover date, and baselined
payload launch date.
Progress report 2: Provide project status report detailing
preflight activities completed to date. Please include results of
3/26/2021
any ground testing completed to date along with status of NASA
Phase 0/I/II/III safety certifications, status of COFR, payload
turnover date and payload launch date.
Progress Report 3: Provide project status report detailing
7/23/2021
payload flight readiness, payload turnover, payload launch, and
any initial in-orbit operations status updates.
30 Day Post-flight Report 4: Provide report detailing in-orbit
operations and initial findings and/or results of the investigation. 2/18/2022
Describe how initial results align with project objectives and next
steps in post flight analysis.
Final Report: Provide a final report which includes detailed 9/17/2022
results of the in-orbit operations and how investigation results

3

007
aligned with project objectives. Please also include next steps
towards applying investigation results to a commercialization
plan, publications, and/ or other applications.
Closeout Documentation– Due 90 days after Final Report.
12/17/2022
(Closeout documents: final patent/inventions report, GFE/GFI
report, final value-impact survey)
Project timeline, accommodating the delay with current and subsequent MISSE launches, is shown in
Table 2.
4
Table 2. Project timeline.
5
IV. Preflight activities completed to date
A. Test Materials
Table 3 below presents overview of the test materials intended for the MISSE-16 experiment. PI, LCP,
and PET films were provided by DuPont, Inc, and CGRP, GFRP, and POSS materials were purchased by
Kirtland AFRL.
Table 3. MISSE-16 Test Materials.
1
2
3
4
5
6
7
8
9
10
11
12
13
14
15
Material
Abbrev.
Trade name
Polyimide (PMDA)
Polyimide (PMDA)
Polyimide (PMDA)
Polyimide (PMDA)
Polyimide (PMDA)
Polyimide (PMDA)
Polyimide (PMDA)
Carbon fiber reinforced
polymer
Glass
fiber
reinforced
polymer
Liquid crystal polymer
Polyethyleneterephthalate
Polyethyleneterephthalate
PI/Polyhedral Oligomeric
Silsesquioxane
PI/Polyhedral Oligomeric
Silsesquioxane
Aluminum oxide
PI
PI
PI
PI
PI
PI
PI
LCP
PET
PET
Kapton CR
Kapton CS
Kapton WS
Kapton XC
Kapton TF
200DR9
Kapton HN
Economyplate™
Carbon Fiber
G-10/FR4 Glass
Epoxy
Zenite®
Melinex® 454
Mylar® MO21
POSS
CFRP
GFRP
Density
(g/cm3)
1.54
1.42
1.42
1.41
1.42
1.42
1.4273
1.4393
1.9
Thickness
(mils)
2
2
1
2
4.0
2
3
29
66
1.75
1.4
1.39
3, 4, 5
9, 10, 14
CORIN®XLS
1.40
0.6
POSS
Thermalbright®N
2.08
0.8
ALUM
Alumina
3.95
3.8
B. MISSE Flight Facility hardware upgrade
Several weeks of technical research and discussion with Alpha Space resulted in a satisfying technical
solution that will yield scientifically useful data in a cost-effective and low failure-probability way. The
following passage, drawn primarily from written communications between our teams, sums up the
discussions that resulted in our decision on the best technical solution available within the constraints of
existing MISSE-FF hardware and experimental budget:
When considering the spectroscopy that we want to do on orbit, we need to design the experiment to
yield the greatest amount of useful data. With that in mind, several factors must be taken into
consideration. In order of importance:
1. Spectral bandwidth: The sensitivity of our imager must span the widest possible spectral range.
Initial investigations into polyimide and polyethylene terephthalate (several varieties of which
we plan to fly on this mission) show that electron damage results primarily in optical changes
between 450-650 nm, while for thermal control paints, extensive changes in reflectivity were
6
2.
3.
4.
5.
observed out to 1800 nm and beyond. The effects of UV exposure are expected to be similar to
that of electrons while AO damage may presumably result in darkening across all spectral bands.
The wider the spectral sensitivity, the more useful the experiment will be for both the proposed
materials and any other material varieties that will be flown in the future.
Spectral resolution: our laboratory experiments measure reflectance data with nm resolution.
While that resolution may not be feasible within the framework of the proposed experiment,
better wavelength resolution will enable us to distinguish similar materials from each other and
to identify specific chemical damage pathways.
Time resolution: a quick cadence of image capture on each face, especially during the initial
radiation exposure, i.e. first month, will allow for validation of chemical kinetic models for
radiation damage that we are currently developing in our laboratories.
Feasibility: The experiment must fit within the mass, volume, power/data usage, interoperability
tolerances of the mission.
Cost: The experiment must provide scientifically rigorous data using the most cost efficient
equipment and procedures.
Keeping in mind these experimental design factors, we envision a few potential configuration of the
orbital experiment:
A. Add a second very small camera (e.g. visible blind Basler Dart with IR bandpass filter) to the
existing MISSE-FF test fixture.
This experimental configuration will allow to Illuminate samples simultaneously with B, G, R, and IR lights.
Light capture in B, G, R channels will be performed with existing Basler Dart camera, and IR channel will
be captured with new Basler Dart camera. The Darts are quite small: 29 x 29 mm wide, 16 mm deep (w/o
optics) and reasonably priced. Ideally, the camera would have sufficient quantum efficiency at 850-940
nm, resulting in 4 broad color bands between the two cameras. This solution would require the addition
of one camera per face, for a total of 2 and therefore would require some skilled Tetris with the current
measurement trolley and presumably additional labor costs from Alpha Space.
B. Single camera with spectral sensitivity in the B,G,R, and IR bands utilizing separate image capture
procedure for R,G,B and IR channels.
Samples will be illuminated sequentially. First, B, G, and R LEDs shine and an image is captured. Then,
samples will be illuminated with only IR LED(s) followed by a second image capture. For this option
resolution would be 4 broad color bands with sufficient sensitivity at B,G,R, IR (850 nm) and very low
sensitivity at 940 nm. This solution would eliminate the need for a second camera, but would require
some additional wiring/circuitry to power the visible and IR LEDs separately. This work would increase the
cost of the project due to increased labor commitment by Alpha Space.
C. Replacement of current camera with multispectral/hyperspectral camera.
This experimental configuration will provide the desired spectral range and resolution (470-900 nm best
case with 150 spectral bands and ~3nm resolution), however, he cost may be prohibitive (~150k for 3
7
cameras). We have approached our respective institutions and there is no obvious source on our side for
an extra 100k
Solution A (multiple cameras) was deemed unfeasible due to limited space on the existing MISSE-FF
camera trolley and the lack of a second USB connection to trigger the second camera independently. The
installation additional USB feedthroughs to enable solution A was not possible due to budgetary
constraints. Solution C (hyperspectral cameras) was not feasible due to the cost of hyperspectral cameras
(>$50k/camera vs. ~$300 per RGB camera) and to the necessity of building motion stages for each camera
as they multispectral capabilities are designed in a line-scan setup. Further, although the spectral
resolution would have increased using a hyperspectral camera, the spectral range would have been
significantly reduced and we would have lost spectral bands that we know to be chemically active under
electron and UV irradiation in PI.
Our teams decided to press forward with a modified solution B. We are going to use the single existing
USB terminated power cable to power a circuit board with a microprocessor to control
individual/sequential LED illumination. This allows for novel 4 color photometry using a $300 COTS
camera. Presently, the details of experimental hardware at MISS-FF are being in final review by Alpha
Space experts. The figure illustrating the illumination/detection scheme that we will utilize on the MISSE
16 mission is shown in Appendix A.
C. Modification of the Bi-directional Reflectance Distribution function (BRDF) measurement setup
To accommodate the BRDF measurements of control (pristine) and flight-duplicate (i.e. atomic oxygen,
electron, and UV exposed) samples, the Orbital Calibration (OrCa) experimental set up developed for the
characterization of directional reflectance properties of the Orca satellite (fig. 1) was significantly
modified.
Figure 1. The OrCa laboratory setup developed by GTRI’s team for the characterization of directional
reflectance properties of the Orca satellite. (a) Orca is rotated on the (b) platform while a (c) light
source remains fixed relative to the face of the Orca satellite being measured. The (d) spectrometer
fore-optic is oriented at fixed intervals to measure the radiance leaving the target from a given viewing
direction.
8
Unlike OrCa satellite, a 12U small sat, the test samples intended for the MISSE-16 mission are flat polymer
sheets with area of each sample is 1 in2 and total thickness of each sample is 10 mil (0.01 inch). The
schematic of experimental fixture designed for the characterization of directional reflectance properties
of control (pristine) and flight duplicate samples is presented in figure 2. It features the 3D printed housing
designed to clamp the polymer samples with (width x length x thickness) = 2” x 2” x 1/8”. Motorized
rotation stage will allow automated and precise sample positioning during the BRDF data acquisition
procedure. Data collection will be performed using the ASD FieldSpec 4 spectrometer designed to collect
radiance measurements over the 350 nm – 2500 nm spectral range. The fore optics of the spectrometer
will be precisely oriented with respect to the sample using the machined holes in one of shoulders of the
experimental fixture. The fore-optic will be positioned 7” away from the sample center, which
corresponds to a sampled area subtended by the fore-optic of approximately 2.15” x 2.15”, which will
reduce the influence of background scene reflections.
Figure 2. (a) Schematic of BRDF measurements experimental fixture including (b) ASD FieldSpec 4
spectrometer and (c) motorized rotation stage.
Special care was taken to create a light source which will approximate the spectral energy density (SED)
by solar SED and collimate the light into a tight beam. Resulting light source optical setup is shown in
figure 3. In addition, improvements in housing being made to reduce stray light.
9
Figure 3. (a) Schematic and (b) experimental light source optical setup for BRDF measurements.
D. Modification of the Surface Potential Decay (SPD) measurement setup
We intend to use the surface potential discharge (SPD) method to evaluate the electrical conductivity of
control (pristine) and flight-duplicate samples. The SPD method uses a low energy electron gun to dust
electrons onto the surface of an insulating material with a grounded backing. Next, the charging beam is
extinguished and the resultant surface potential is then measured as a function of time using a noncontact electric field probe (TREK model 370). Figure 4 demonstrates the exterior and interior of vacuum
chamber intended for SPD measurements. The initial configuration of SPD setup presented in figure 4 was
sufficient for measurements of one sample with area of 2 in2 and total thickness of several mils. However,
to increase the measurement throughput and optimize the samples handling procedure, we modified the
SPD setup as illustrated in figure 5. New configuration of grounded plate shown in figure 5 may
accommodate 12 samples with 1 in2 area allowing for their simultaneous SPD measurements.
Figure 4. The exterior and interior of vacuum chamber intended for SPD measurements.
10
Figure 5. The modifed grounded plate of SPD setup.
E. Development of MISSE-16 Unified Test Plan for Pre-flight Characterization
Unified test plan for pre-flight characterization of MISSE-16 materials describing specific tests that will be
carried out by each institution/agency on control (pristine) MISSE-16 materials was developed. These tests
include surface morphology, optical properties, and charge transport properties characterization using
scanning electron microscopy (SEM), atomic force microscopy (AFM), optical transmission and reflectance
measurements, bi- directional reflectance distribution function (BRDF) measurements, and electrical
conductivity measurements. The developed MISSE-16 Unified Test Plan for Pre-flight Characterization is
presented in Appendix B.
V. Progress made since last report
N/A
VI. Near-term forward work
During the time period between now and the next progress report we plan to achieve the following goals:
A. MISSE Flight Facility hardware upgrade
Hardware design of each carrier housing accommodating a Basler daA1600-60uc camera with improved
spectral response in IR region along with additional IR illumination source (IR LED) to provide a broader
illumination range in IR region will be finished. We plan on including the timed lights illumination scheme
for the visible and IR light sources, in order to avoid installation of second camera or camera rebuild.
B. Payload turnover/launch dates
11
Based on Alpha Space schedule, payload turnover date and payload launch date will be finalized.
C. Characterization of control (pristine) materials
We plan to finish characterization of control (pristine) materials, including their surface morphology,
optical properties, and charge transport properties, in according to the developed MISSE-16 Unified Test
Plan for Pre-flight Characterization (see Appendix B).
VII. Issues/concerns/risks
N/A
12
Appendix A: Illumination/detection scheme proposed for the MISSE 16 mission.
13
Appendix B: MISSE-16 Unified Test Plan for Pre-flight Characterization.
MISSE-16 Unified Test Plan for Pre-flight Characterization
Dr. Elena Plis (PI, GTRI), Drs. D. Engelhart and V. Murray (co-PIs, AFRL)
Dr. G. Badura (co-I, GTRI), Dr. H. Cowardin (co-I, NASA) Drs. Cone (Co-I, UTEP) and Roberson (Co-I,
UTEP); Mr. R. Hoffmann (co-I, AFRL), Mr. T Scott (Co-I)
Acronyms
AFM
AFRL
AO
ASD
BRDF
CFRP
DSLR
GFRP
GTRI
LCP
LED
nm
ODSMSD
OMC
PET
POSS ®
SCICL
SEM
UV
VIS
Atomic Force Microscopy
Air Force Research Laboratory
Atomic oxygen
Analytical Spectral Device
Bidirectional reflectance distribution
carbon fiber reinforced polymers
digital single-lens reflex (camera)
Glass Fiber Reinforced Polymer
Georgia Tech Research Institute
Liquid Crystal Polymer
Light Emitting Diode
Nanometer
Orbital Debris Spacecraft Materials Spectral Database
Optical Measurement Center
Polyethylene Teraphthelate Polymer
Polyhedral Oligomeric Silsesquioxane
Spacecraft Charging and Instrument Calibration Laboratory
Scanning Electron Microscope
Ultraviolet
Visible
1. Introduction
The following document will outline the specific tests that will be carried out by each institution/agency
on control (pristine) MISSE-16 materials. These tests include surface morphology, optical properties, and
charge transport properties characterization using scanning electron microscopy (SEM), atomic force
microscopy (AFM), optical transmission and reflectance measurements, bi-directional reflectance
distribution function (BRDF) measurements, and electrical conductivity measurements. In addition, color
photo-images of control (pristine) samples will be taken using a comparable DSLR camera.
14
2. Test Materials
2.1. Polyimides (PIs)
We will use the well-studied Kapton-HN® film as a reference material for our experiment. In addition,
several new Kapton varieties will be flown, in particular, CR (corona-resistant; important for shielding of
sensitive equipment on-board of a spacecraft), CS (clear, smooth, thermoformable; possible replacement
of traditional PI film in MLI blankets), WS (possible alternative for traditional PI film in MLI blankets), XC
(electrically conductive; important for space charge mitigation on-board of a spacecraft), and TF
(thermoformable; possible candidate for small satellite parts manufacturing). Finally, novel 200DR9
polymer material which represents an untested variety of high performance organic polymer for
spacecraft applications is included into PI selection list.
2.2. Polyethylene terephthalate (PET) films
PET film (Mylar) is an essential part of MLI blankets which are used on the exterior surfaces of spacecraft
for the purpose of passive thermal control. We will use two novel types of Mylar, Melinex® 454 and Mylar®
MO21, intended for the improved MLI design and durability.
2.3. Liquid crystal polymers (LCPs)
LCPs present a special category of materials which have excellent dimensional stability (fatigue and creep
resistance), as well as high dielectric strength over wide temperature range. These unique properties in
conjunction with light weight make LCP materials very attractive for space applications. Before
recommendation of LCP material to spacecraft designers, evaluation of its durability under the harsh
space conditions of LEO need to be performed, which motivated us to include one variety of DuPont
Zenite® LCP into the study.
2.4. PI/Polyhedral Oligomeric Silsesquioxane (POSS) hybrid materials
Many applications that could benefit from using a thin polymer film are restricted from use due to the
fact that many currently available materials do not meet durability or packaging requirements. A
modification of NeXolve's AO-resistant CORIN® XLS polyimide is proposed that would improve orbit
lifetime and provide lightweight and high packaging efficiency alternatives for use in Low Earth Orbit (LEO)
applications. We included POSS Corin XLS and doped-POSS materials Thermalbright™ into material
selection list since whereas PI/POSS hybrid materials have already been identified as excellent AO
resistant materials, a detailed study of their optical properties will be of great utility to the remote sensing
community for the characterization of orbital debris and operational RSOs.
2.5. Carbon Fiber Reinforced Polymer (CFRP) and Glass Fiber Reinforced Polymer (GFRP)
CFRP and GFRP are two examples of modern materials used in the construction of present-day LEO
satellites. Unfortunately, minimal tests have been conducted in laboratory settings to better characterize
their optical properties and potential degradation in the LEO environment. The objective for studying
these materials, specifically, would provide a baseline of how these materials respond to the harsh LEO
environment, utilizing optical monitoring for better space domain awareness. CFRP has a very low albedo
15
and during laboratory impact tests has shown to fragment into small needle-like shapes, but it is unclear
whether these fragments also change their albedo which makes ground-based observations of these
bullets with low albedo a clear risk in the orbital debris environment. Secondly, GFRP tends to be
translucent in nature and laboratory studies have not determined how or if they degrade with time.
Summary of the different materials utilized for the ground-based measurements in support of the MISSE16 experiment including material abbreviation, trade name(s), and thicknesses of the tested films is given
in Table 1.
Table 1. MISSE-16 Test Materials
Material
1
2
3
4
5
6
7
Trade name
PI
PI
PI
PI
PI
PI
PI
Density
(g/cm3)
1.54
1.42
1.42
1.41
1.42
1.42
1.4273
Thickness
(mils)
2
2
1
2
4.0
2
3
1.4393
29
1.9
66
1.75
1.4
1.39
3, 4, 5
9, 10, 14
9
Glass fiber reinforced polymer
GFRP
10
11
12
Liquid crystal polymer
Polyethyleneterephthalate
Polyethyleneterephthalate
PI/Polyhedral Oligomeric
Silsesquioxane
PI/Polyhedral Oligomeric
Silsesquioxane
Alumina
LCP
PET
PET
Kapton CR
Kapton CS
Kapton WS
Kapton XC
Kapton TF
200DR9
Kapton HN
Economyplate™
Carbon Fiber
G-10/FR4 Glass
Epoxy
Zenite®
Melinex® 454
Mylar® MO21
POSS
CORIN®XLS
1.40
0.6
POSS
Thermalbright®N
2.08
0.8
ALUM
Alumina
3.95
3.8
8
13
14
15
Polyimide (PMDA)
Polyimide (PMDA)
Polyimide (PMDA)
Polyimide (PMDA)
Polyimide (PMDA)
Polyimide (PMDA)
Polyimide (PMDA)
Carbon fiber reinforced
polymer
Abbrev.
CFRP
3. Testing Procedure
3.1. NASA JSC Optical Measurement Center
To obtain pre-flight spectral measurements of the control (pristine) MISSE-16 materials to provide base
line characteristics in comparison to laboratory exposed and in situ measurements. The data collected will
utilize the instrumentation outlined below in the NASA JSC OMC. All spectral data will be acquired and
analyzed by Dr. H. Cowardin (NASA) and Ms. Jackie Reyes (UTEP). The data will be shared with the MISSE
team in the form of spectral plots and associated spectral reflectance text documents. Lastly, data
collected will be stored in the Orbital Debris Spacecraft Materials Spectral Database (ODSMSD).
16
Figure 1. Left: Spectrometer being used to measure coating of Hubble radiator,
Right: Spectrometer setup for laboratory measurements.
3.1.1. Instrumentation
An Analytical Spectral Device (ASD) field spectrometer, ASD FieldSpec4, with a range from 300-2500 nm
and a full width half maximum of 3 nm at 700 nm, and 10 nm at 1400/2100 nm, respectively. The ASD
FieldSpec 4 has a field of view of ~ 25°. Measurements are acquired by placing the target under the quartz
lamp illuminator and orienting the spectrometer’s fiber feed (mounted in a pistol grip) approximately
perpendicular to the target surface, as shown in Figure 1. A Spectralon panel from Labsphere is used as a
reference standard, scattering light uniformly and is ~ 100% reflective when properly maintained, and is
required calculate the absolute reflectance.
3.1.2. Methodology and Measurement Requirements



The measurements will require a dark room (i.e., no overhead florescent lights) and standard 120
V power outlets.
The measurements will take ~2 hours depending on the number of samples and calibration
needed for various spectral properties. Spectral measurements shall only take one-two days to
measure all materials.
A list of all materials and their properties would be needed prior to set-up for familiarization and
planning.
3.2. SCICL team/AFRL
3.2.1. Electrical Conductivity Measurements
3.2.1.1. ASTM D-257
Due to the low electrical conductivity of the pristine test materials, the constant voltage method
performed with a Keithley 6517B low-conductivity meter and a Keithley model 8009 test fixture
conforming to ASTM D-257 at different values of applied bias will be utilized to measure the electrical
conductivity of the control (pristine) materials. ASTM D-257 dictates that bulk resistivity be determined
by measuring the current passing through the material while sandwiched between electrodes across
which a bias is applied. Thus, volume electrical resistivity values (cm) will be determined directly from
the measurements using the following equation
17
𝜌=
22.9 ∙ 𝑉
𝑡𝑐 ∙ 𝐼
(1)
Where  is the volume resistivity of the specimen, V is the applied voltage for the electrometer, tc is the
average thickness of the specimen in centimeters, I is the current reading from the electrometer. The
electrical conductivity of the test sample will be determined as reciprocal to the electrical resistivity value.
All electrical conductivity data will be acquired and analyzed by Drs. D. Engelhart and V. Murray. The data
will be shared with the MISSE team in the form of conductivity vs applied bias plots and associated text
documents.
Figure 2. The Keithley model 8009 test fixture to be utilized by SCICL team/AFRL for evaluation of
electrical conductivity of control (pristine) test materials.
3.2.1.1.1. Methodology and Measurement Requirements




At least 2 hours of vacuum dehydration bake-out at 60C will need to be performed on each
sample prior measurements;
The measurements will be performed at least three times at each value of the applied bias (100
V, 500 V, and 1000V);
Thickness (in cm) of each sample will need to be measured prior or after measurements for the
resistivity values calculation suing equation (1);
Measurements shall only take one-two days to measure all materials.
3.2.1.2. Surface Potential Decay (SPD) Method
The resistivity derived from the ASTM D-257 reflects the behavior of the least resistive domains in a
heterogeneous material. To reflect the contribution of the most resistive domains, the SPD method will
be utilized. It uses a low energy electron gun to dust electrons onto the surface of an insulating material
with a grounded backing. Next, the charging beam is extinguished and the resultant surface potential is
then measured as a function of time using a non-contact electric field probe (TREK model 370). As the
electrons are transported through the material, the surface potential will decay towards zero. The rate of
this decay is directly related to the electrical conductivity of the material. Figure 3 demonstrates the
exterior and interior of vacuum chamber intended for SPD measurements. All electrical conductivity data
18
will be acquired by Drs. D. Engelhart and V. Murray. SPD data will be analyzed by Dr. E. Plis. Then data
will be shared with the MISSE team in the form of text documents.
Figure 3. (Left) The exterior and (right) interior of vacuum chamber for SPD measurements.
3.2.1.2.1. Methodology and Measurement Requirements
 At least 2 hours of vacuum dehydration bake-out at 60C will need to be performed on each
sample prior measurements;
 The SPD setup is capable of simultaneous measurements of 12 samples with individual area of 1
in2;
 Due to low conductivity of control (pristine) materials, SPD measurements will take at least five
days to complete. The measurement is finished once surface potential decays below 5% of its
initial value;
 Due to the low electrical conductivity of control (pristine) materials, at least two weeks are
required to measure all the MISSE-16 materials.
3.2.2. Surface Morphology Evaluation
Surface morphology of control (pristine) materials will be accessed with the VIC-3D Stereo Microscope
Measurement System located at Kirtland AFB. All surface morphology data will be acquired and analyzed
by Drs. D. Engelhart and V. Murray. The data will be shared with the MISSE team in the form of jpeg
images.
3.2.2.1. Methodology and Measurement Requirements
 Samples will need to be cleaned with sequential Acetone and IPA solutions for 1 minute each and
then baked-out at least for 2 hours under vacuum at 60C prior measurements;
 2D and 3D measurements will be performed at least on three different areas of 2” x 2” sample
area;
19


Each measurements will be performed with four different field of view (zoom range): 0.8mm, 2
mm, 5 mm, and 7mm;
Optical measurements shall take three days to measure all materials.
3.2.3. Optical Transmission/Reflectance Measurements
Optical characterization of control (pristine) materials will be performed using Perkin-Elmer Lambda 950
UV/VIS and Surface Optics Corporation SOC-400T portable FTIR spectrometers in ambient atmosphere.
All data will be acquired by Drs. D. Engelhart and V. Murray, then transferred Dr. E. Plis for analysis. Finally,
data will be shared with the rest of the MISSE team in the form of optical reflectance/transmittance vs
wavelength plots and associated text documents.
3.2.3.1. Methodology and Measurement Requirements
 Samples will need to be cleaned with sequential Acetone and IPA solutions for 1 minute each and
then baked-out at least for 2 hours under vacuum at 60C prior measurements;
 Prior measurements, spectrometers will need to be stabilized (warmed up) during at least 2 hours;
 Optical measurements shall take one day to measure all materials.
3.3. GTRI team
3.3.1. Surface Morphology Measurements
3.3.1.1. DSLR camera measurements
Each sample will be measured with DSLR camera individually. Measurements will be performed by Dr. E.
Plis and shared with the rest of the MISSE team in a form of the jpeg images.
3.2.3.1. Methodology and Measurement Requirements
 Samples will need to be cleaned air-brush prior measurements;
 Measurements shall be taken using white tech-wipe as a background.
3.1.1.2. Atomic Force Microscopy (AFM) measurements
Surface roughness measurements will be performed using AFM located at Georgia Tech IEN/IMAT
Materials Characterization facility. Its proprietary sensor design achieves closed-loop performance with
open-loop noise levels for previously unseen resolution on a large-sample, tip-scanning AFM. Significantly
reduced noise floor enables imaging at atomic level in contact mode, with less than 30pm in
TappingMode™. Drift rates less than 200pm per minute render distortion-free images. It allows to
measure surface roughness up to 5 µm on areas as large as 200 µm x 200 µm. AFM measurements will be
performed by Dr. E. Plis. Results will be shared with the rest of the MISSE team in a form of 2D and 3D
jpeg images as well as RMS roughness values.
3.1.1.2.1. Methodology and Measurement Requirements
 Samples will need to be cleaned with N2-brush each prior measurements;
20


Five 5 µm x 5 µm scans will be taken per each sample. Surface roughness will be calculated for
each scan taken, and the Ra surface roughness determined as an average roughness of individual
measurements;
AFM measurements shall take at least one week to measure all materials.
3.1.1.3. Scanning Electron Microscopy (SEM) measurements
Surface morphology of test samples will be accessed using a high resolution (2 nm) Field emission Hitachi
S4100 SEM is a part of the Georgia Tech IEN/IMAT Materials Characterization facility, located at the
Materials Analysis Center (MAC). It is equipped with a NORAN thin window light element x-ray detector,
image analysis (compositional mapping and line scans) and quantitative software, and full color display.
The Hitachi SEM also contains a TV speed four quadrant backscatter detector (operable at 500 volts) and
houses a custom built conductivity probe. SEM measurements will be performed by Dr. E. Plis. Results will
be shared with the rest of the MISSE team in a form of 2D images.
3.1.1.3.1. Methodology and Measurement Requirements





Samples will need to be cleaned with air-brush each prior measurements;
Measurements will be performed at least on three different areas of 2” x 2” sample area;
Each measurements will be performed with three different field of view (zoom range): 400 µm,
100 µm, and 40 µm;
If material is non-conductive, no further attempts shall be undertaken to acquire an image since
it may destroy the material due to electron bombardment. In addition, no coating with conductive
substance (gold or silver) shall be applied to the material. Instead, conclusion about surface
morphology will be based using optical microscope and AFM images;
SEM measurements shall take two-three days to measure all (conductive) materials.
3.3.2. Bi-directional Reflectance Distribution function (BRDF) Measurements
The schematic of experimental fixture to be utilized for the characterization of directional reflectance
properties of control (pristine) samples is presented in figure 4. It features the 3D printed housing
designed to clamp the polymer samples with (width x length x thickness) = 2” x 2” x 1/8”. Motorized
rotation stage will allow automated and precise sample positioning during the BRDF data acquisition
procedure. Data collection will be performed using the ASD FieldSpec 4 spectrometer designed to collect
radiance measurements over the 350 nm – 2500 nm spectral range. The fore optics of the spectrometer
will be precisely oriented with respect to the sample using the machined holes in one of shoulders of the
experimental fixture. The fore-optic will be positioned 7” away from the sample center, which
corresponds to a sampled area subtended by the fore-optic of approximately 2.15” x 2.15”, which will
reduce the influence of background scene reflections.
21
Figure 2. (a) Schematic of BRDF measurements experimental fixture including (b) ASD FieldSpec 4
spectrometer and (c) motorized rotation stage.
The BRDF measurements will be performed by Dr. G. Badura. He will analyze the data, and then share
them to the rest of the MISSE team in a form of charts and text documents.
3.3.2.1. Methodology and Measurement Requirements
 Samples will need to be cleaned with air-brush each prior measurements;
 The measurements will require a dark room (i.e., no overhead florescent lights) and standard 120
V power outlets;
 BRDF measurements shall take two-three days to measure all materials.
3.4. UTEP team
At the moment, the laboratory access of the UTEP team is pending due to the pandemic. The
measurements to be decided once access to the facilities is granted.
22
Characterization of novel spacecraft materials under high
energy electron and atomic oxygen exposure
Elena A. Plis∗
Georgia Tech Research Institute (GTRI), Atlanta, GA, 30318, USA; Assurance Technology Corporation, Carlisle, MA,
01741, USA
Miles T. Bengtson†
National Research Council Research Associateship Program, 3550 Aberdeen Ave., Kirtland AFB, New Mexico 87117,
USA
Daniel P. Engelhart‡
University of New Mexico, Albuquerque, NM, 87131, USA
Gregory P. Badura§
Georgia Tech Research Institute (GTRI), 925 Dalney St NW, Atlanta, GA 30318
Heather M. Cowardin¶
NASA Johnson Space Center, Orbital Debris Program Office, 2101 NASA Pkwy., Houston, TX 77058, USA
Jacqueline A. Reyes‖
University of Texas at El Paso, 500 W. University Ave., El Paso, TX, 79968, USA
Ryan C. Hoffmann, Alexey Sokolovskiy, Dale C. Ferguson∗∗
Air Force Research Laboratory, Space Vehicles Directorate, USA, Kirtland AFB, Albuquerque, NM, 87117, USA
Jainisha R. Shah and Sydney Horne, ††
Assurance Technology Corporation, Carlisle, MA, 01741, USA
Timothy R. Scott‡‡
DuPont de Nemours, Inc, Durham, NC 27703, USA
The harsh space environment imposes stringent requirements upon spacecraft materials,
especially those located on exterior surfaces of space objects in low Earth orbit (LEO). As
humankind moves from space exploration to space commercialization, these polymers may
have to last for 15–20 years without considerable degradation of their material properties.
Hence, we must understand the effects of the space environment on materials currently in use
as well as on untested materials. In the presented study authors exposed several space-relevant
polymers to the simulated space weather comprised by electrons and atomic oxygen irradiation.
First, materials were thoroughly characterized in their pristine state to create a baseline for
the ground- and space-based experiments. Next, alteration of optical, surface, and charge
transport properties of the same materials were studied under space-simulated environment.
∗ Senior Research Engineer, Georgia Tech Research Institute (GTRI), 925 Dalney St NW, Atlanta, GA 30318; Assurance Technology Corporation,
Carlisle, MA, 01741, USA
† Postdoctoral Fellow, National Research Council Research Associateship Program, 3550 Aberdeen Ave., Kirtland AFB, New Mexico 87117, USA
‡ Research Professor, Department of Chemistry and Chemical Biology, University of New Mexico, 300 Terrace St NE, Albuquerque, NM, 87131.
§ Research Scientist, Georgia Tech Research Institute (GTRI), 925 Dalney St NW, Atlanta, GA 30318.
¶ Laboratory and In Situ Lead Orbital Debris Business Unit Manager, NASA Johnson Space Center, Orbital Debris Program Office, 2101 NASA
Pkwy., Houston, TX 77058, USA
‖ University of Texas at El Paso, 500 W. University Ave., El Paso, TX, 79968, USA
∗∗ Air Force Research Laboratory, Space Vehicles Directorate, USA, Kirtland AFB, Albuquerque, NM, 87117, USA
†† Assurance Technology Corporation, Carlisle, MA, 01741, USA
‡‡ Business Development Leader, Aerospace & Defense, DuPont de Nemours, Inc, Durham, NC 27703.
1
I. Introduction
s humankind’s forays into space become simultaneously more commonplace and ambitious, the need for a
fundamental understanding of the interactions between the harsh space environment and spacecraft materials
becomes ever more important. Research has shown that the physical, chemical, and optical properties of matter change
continuously as a result of exposure to solar radiation and aggressive chemical species produced in Earth’s upper
atmosphere. A thorough knowledge of the evolution of material properties throughout a planned mission lifetime is of
primary importance when designing long-term space missions. Further, as the requirements for new space missions
become more stringent and extensive, novel lightweight materials must be developed with improved long-term radiation
shielding and mechanical properties for use in internal and external spacecraft systems. Newly developed materials
must be thoroughly vetted before implementation. Characterization of heritage materials in the space environment is
also necessary for identification and tracking of orbital debris, a crucial first step toward debris remediation [1, 2].
Ground-based simulation of the space environment is experimentally challenging because the space environment
varies wildly depending on orbit, solar conditions, and many other factors [3]. Accurate space weather simulation must
include several different species of radiation, charged and uncharged chemical species, and temperature fluctuations [4].
Precise material characterization is also demanding as post-irradiation measurements must be carried out in vacuum or
under inert gas flow, as atmospheric gases can participate in post-irradiation material chemistry, rendering any results
non-representative of a material’s actual behavior on orbit. Therefore, ground testing requires thorough validation in
order to develop trustworthy models of material degradation.
One major barrier to utilizing novel materials on spacecraft is the immense cost of flying to space. With exact
knowledge of material properties throughout an expected mission lifetime, the best material can be chosen to ensure
optimal performance for the entirety a spacecraft’s mission. By developing thorough and accurate ground-based space
weather simulation capabilities, the barrier to vetting new materials will be reduced considerably. A newfound freedom
in material selection could very well lead to dramatic advances in design that would otherwise be unfeasible.
The the Materials International Space Station Experiment Flight Facility (MISSE-FF) at International Space Station
(ISS) is the perfect test bed to generate benchmark data with which to validate the efficacy of ground-based space weather
simulation experiments [5, 6]. Our ultimate goal is to fly several novel and well-understood materials on the MISSE-FF
and monitor the changes in spectral reflectivity that occur as a result of exposure to different components of the low
Earth orbit (LEO) environment. Energy deposition from the environment leads to chemical changes in the material that
in turn alter the optical properties. The same chemical damage that manifests as changes in optical reflectance and
absorptance also leads to changes in myriad physical properties such as mechanical strength, electrical conductivity,
and chemical reactivity [7–10]. Correlation of the changes in each of these properties as function of radiation flux
and fluence will allow us to infer a host of material properties based on one experimentally convenient measurement:
color change. By flying identical test fixtures on the ram, zenith, and wake positions of the MISSE-FF and collecting
spectrally resolved images of the materials throughout the mission along with concomitant measurement of the ambient
space environment, we may measure the optical changes that occur in each material as a function of exposure to neutral
atomic oxygen (AO), unfiltered solar ultraviolet (UV) radiation, and electrons, respectively. Correlation of the MISSE
data with extensive testing of flight-duplicate (AO-, UV-, and electron-exposed) and control (pristine) samples will
enable development of fundamental chemical models for material degradation.
In presented study, several spacecraft polymers were first thoroughly characterized in their pristine state to create
a baseline for the ground- and space-based experiments. Next, alteration of optical, surface, and charge transport
properties of the same materials were studied under space-simulated environment comprised by, separately, high-energy
electron and AO exposure. The same material selection is planned to be flown during the MISSE-16 mission; however,
the space experiment data as well as their comparison with the results of ground experiments is the subject of future
publication.
A
II. Experimental Details
A. Materials
This study focuses on several visually white materials from the different classes of polymers, namely, polyimides
(PIs) (Kapton®WS Glossy), Polyethyleneterephthalate or PET (Melinex®454, M021), liquid crystal polymer (Zenite®),
and PI/Polyhedral Oligomeric Silsesquioxanes (POSSs, Thermalbright™N), as shown in Table 1.
Kapton®has been used for more than 50 years to protect nearly every satellite, spacecraft, and astronaut since Apollo
11. The Kapton®WS Glossy is an opaque film with high thermal stability; possible alternative for traditional PI film in
2
Table 1
Material
Kapton®WS Glossy
Melinex®454
Mylar®M021
Zenite®
Thermalbright™N
Studied spacecraft polymers.
Abbrev.
PI
PET
PET
LCP
POSS
Thickness (mil)
1
5
10
3
0.8
Potential application
Thermal control foils
Improved reflectivity of MLI
Improved reflectivity of MLI
Flexible antennas and circuits
AO-resistive insulation
*MLI stands for the multi-layer insulation
**Trade names and trademarks are used in this report for identification only. Their usage does not constitute an
official endorsement, either expressed or implied, by the National Aeronautics and Space Administration.
MLI blankets.
PET film, more often known by its trade name Mylar, is an essential part of MLI blankets that are used on the
exterior surfaces of spacecraft for the purpose of passive thermal control. These composite materials work by limiting
the amount of radiative heat transfer through multiple layers of thin reflectors and spacer materials. An ideal thermal
blanket reflects 100% of the incident radiation. In practice, this ideal value has not been obtainable using traditional
MLI blankets. Thus, new materials, such as Melinex®454 and Mylar®M021 need to be developed for improved MLI
reflectivity and durability.
LCPs present a special category of materials, which straddles the boundary between an ordinary solid and a liquid.
LCPs have excellent dimensional stability (fatigue and creep resistance), as well as high dielectric strength over wide
temperature range[11]. Moreover, LCPs show a high resistance to UV radiation and very good electrical insulation
properties [12]. These unique properties in conjunction with light weight make LCP materials very attractive for space
applications. Flexible LCP antennas and LCP-based circuits molded to available spacecraft areas could eliminate heavy
metal boxes that currently house rigid circuit boards [13].
PI/Polyhedral Oligomeric Silsesquioxane (POSS) hybrid materials were proposed to increase the AO-resistivity of
PI films by incorporation of the monomer POSS into the PI polymer matrix, either as a chemically bound co-polymer or
an additive blended into the matrix [14]. The unique benefit of this method over traditional methods, such as coatings,
is that when POSS-PI is initially exposed to AO, it forms a layer of SiOx over the surface. This passivating layer
prevents further material ablation, thus lowering the overall erosion yield. Moreover, if the surface is damaged in anyway
and cracks appear, the newly exposed surface will simply form another layer of SiOx. Hence, POSS-PI essentially a
self-healing polymer. As PI/POSS hybrid materials have already been identified as excellent AO resistant materials, a
detailed study of Thermalbright™PI’s optical properties will be of great utility to the remote sensing community for the
characterization of orbital debris.
Figure 1 demonstrates the photographs of studied materials in their pristine state.
Kapton
Fig. 1
WS
Melinex
454
Mylar
MO21
Thermalbright
N
Zenite
Photographs of materials under study in their pristine state
B. Irradiation procedure
High energy electron irradiation was performed with high energy (100 keV) mono-energetic electron radiation from
a Kimball Physics EG8105-UD electron flood gun in the Spacecraft Charging and Instrument Calibration Laboratory
(SCICL) at Kirtland Air Force Base in New Mexico, USA [15]. Samples were mounted over a reflective aluminum
surface of the same area, then samples were adhered to a carousel that rotated through the hot spot of the electron
beam to ensure uniform irradiation. Prior to the radiation exposure, the 100 keV electron beam was characterized via a
3
Faraday cup (FC) housed in the center of the carousel. The relative geometry of the hot spot of the electron beam, the
FC, and the sample positions on the rotating carousel are shown in Figure 2. More details of the electron irradiation
procedure are reported elsewhere [16]. Materials were irradiated to the maximum fluence of 4 x 1013 electrons/cm2 ,
which corresponds to approximately one year of equivalent LEO exposure. AO exposure was performed using the FAST
source at the Physical Sciences Inc, as shown in Figure 3 in according to ASTM-E2089-15. The effective peak atomic
oxygen fluence during the exposure was 3.1 x 1020 O/cm2 . The 8 km/s O-atom beam was generated in high vacuum
chamber with pulsed laser discharge. AO exposure fluence was determined using the reference Kapton®H and the
published erosion rate for Kapton®H by 5 eV AO, 3 x 1024 cm3 /AO. During the exposure, temperature was controlled
(25C) across the 11.4 cm x 11.4 cm sample holder to accommodate simultaneous and uniform exposure of materials.
-8
-6
-4
-2
0
0
-0.2
Y Axis / Inches
-2
-0.4
-4
-0.6
-0.8
-6
-1.0
-8
-1.2
Electron Flux x 10-7 / A cm-2
0
X Axis / Inches
Fig. 2 Flux map for the 100 keV electron beam. Relative positions of the sample carousel (white dashed line),
the FC (white square), and the mounted samples (black dashed circles) are shown. A sample remains in the hot
spot of the beam (solid black circle) for a total of 60 seconds before the carousel is rotated clockwise (direction
indicated by the white arrow).
Fig. 3
FAST source utilized for the AO exposure of materials. Image credit: Physical Sciences Inc.
C. Materials characterization
1. Surface properties
Surface morphology and roughness of studied materials was examined using Bruker Dimension ICON atomic force
microscopy (AFM) allowing measurement of surface roughness up to 5 𝜇m on areas as large as 200 𝜇m x 200 𝜇m.
4
2. Optical properties
Transmittance and reflectance of pristine materials were characterized using Perkin-Elmer UV/Vis spectrometer
within 200 – 2500 nm spectral range. In addition, Fourier-Transform Infrared (FTIR) spectroscopy measurements are
utilized to identify the vibrational bands of different chemical moieties of pristine and irradiated materials. The usual
practice of plotting transmission (in %), as a function of wavenumber (in 1/cm), has been followed. The minima or
dips in the transmission spectra refer to absorption maxima corresponding to characteristic molecules and groups. The
direct hemispherical reflectance (DHR) measurements of pristine and irradiated films will be also measured during
the irradiation procedure in according to the optical data acquisition procedure reported elsewhere [17]. Finally, the
Hemispherical Conical Reflectance Factor (HCRF), the laboratory approximation of true Bidirectional Reflectance
Distribution Function (BRDF) measurements, of pristine samples with fixed viewing angles (0◦ , 30◦ , 60◦ , -30◦ , and
-60◦ ) and variable illumination angles (0◦ - 70◦ degrees range, exact values of illumination angles are specified separately
for each measurement), was evaluated in the principal plane of illumination. The samples were illuminated with
a collimated 50 Watt tungsten halogen that approximated a blackbody source over the range of 350-2500 nm. An
Analytical Spectral Devices Spectrometer was used to collect radiance measurements over this spectral range.
3. Charge transport properties
To assess the charge transport properties of pristine and irradiated materials the surface potential decay (SPD)
method was utilized [18]. In short, the SPD method uses a low energy electron gun to dust electrons onto the surface
of an insulating material with a grounded backing. Next, the charging beam is extinguished and the resultant surface
potential is then measured as a function of time using a non-contact electric field probe (TREK model 370). As the
electrons are transported through the material, the surface potential will decay towards zero. The rate of this decay is
directly related to the electrical conductivity of the material.
4. Mass loss
Mass loss of AO-exposed materials was measured with 10 microgram accuracy using the microgram balance.
Prior weight measurements, all samples were stored in the vacuum chamber for 24 hours to remove water. They were
then removed and weighed over time to monitor water absorption and allow for the dry mass to be calculated. Next,
AO-erosion rate was estimated.
III. Results
A. Electron exposure
In according to [19], mean annual electron flux (>100 keV, electrons, orbit averaged) experienced by the International
Space Station (ISS) is 1013 electrons/cm2 /second. The maximum electron fluence the materials under investigation
were exposed to is 2.2 x 1014 electrons/cm2 , which corresponds to approximately ten years at 600-800 km (LEO) orbit,
delivered during 5.5 hours. Visual inspection did not reveal significant discoloration of the irradiated materials after
irradiation with maximum electron fluence.
Table 2 summarizes the surface roughness of pristine and electron-irradiated samples accessed by AFM. Average
surface roughness (𝑅 𝑎 ) values are average of several 5𝜇m x 5𝜇m scans taken at different parts of the pristine and
electron-irradiated samples. Figure 4 shows representative AFM scans for each studied sample. Surface roughness of
all irradiated materials, except Zenite®, improved in the range of 4.3 - 43.6 nm.
Table 2
Surface roughness of pristine and electron-irradiated polymer samples
Material
Kapton®WS Glossy
Zenite®
Melinex®454
Mylar®M021
Thermalbright®N
Pristine (𝑅 𝑎 , nm)
72.3
14.9
8.2
7.8
52.5
5
Irradiated (𝑅 𝑎 , nm)
28.7
22.6
4.9
3.5
18.1
Fig. 4
Representative 5𝜇m x 5𝜇m AFM scans of pristine and electron-irradiated materials.
Values of volume resistivities of pristine and electron-irradiated with maximum fluence materials measured by the
SPD method are summarized in Table 3. Charge transport properties of irradiated materials were affected by the electron
irradiation; Kapton®WS demonstrated one order of magnitude reduction of its resistivity. Resistivity of Melinex®454
was also reduced but only by factor of 1.3. Another PET material, Mylar®M021, became essentially an insulator, and
resistivity of Zenite®as well as Thermalbright®N materials was increased by an order of magnitude.
Table 3 Volume resistivities of studied materials in their pristine and electron-irradiated states measured by
the SPD method.
Material
Kapton®WS Glossy
Zenite®
Melinex®454
Mylar®M021
Thermalbright®N
Pristine (Ω·cm)
4.5 ·1019
1.2 · 1019
9.6 · 1018
4.3 · 1018
1.6 · 1019
Irradiated (Ω·cm)
6.9 ·1018
7.9 · 1020
7.5 · 1018
2.0 · 1022
1.2 · 1022
Figure 5 displays the absolute reflectance spectra of pristine and irradiated with high-energy electrons polymer
materials. Irradiation with maximum electron fluence degraded the reflectance of each material in 400 - 800 nm range
except Thermalbright®N. Degree of the change depends on the material, with maximum degradation demonstrated by
the Melinex®454 at 350-420 nm range.
Figure 6 demonstrates the FTIR spectra of pristine and electron-irradiated materials in 400 - 1800 1/cm range. It
should be noted that no new peaks are generated after electron exposure, suggesting that no new chemical moieties are
formed.
Figure 7 demonstrates the results of HCRF measurements of for pristine and irradiated materials performed at fixed
viewing angle (30◦ ) and variable illumination angles. It should be noted that all studied materials showed a significant
reduction in specular reflection – both in terms of the peak specular signature and the width of the specular lobe.
6
1.0
1.0
Kapton WS
Thermalbright N
0.8
0.8
0.6
0.6
DHR
DHR
0.8
DHR
1.0
Melinex 454
M021
0.4
0.4
0.4
Pristine
1.0 x 1014 e/cm2
2.2 x 1014 e/cm2
0.2
400
800
1200
Wavelength (nm)
1600
Pristine
1.0 x 1014 e/cm2
2.2 x 1014 e/cm2
0.2
0.0
0.6
400
800
1200
Wavelength (nm)
1600
Zenite
Pristine
1.0 x 1014 e/cm2
2.2 x 1014 e/cm2
0.2
0.0
400
800
1200
1600
Wavelength (nm)
Fig. 5 DHR values of pristine and irradiated with high-energy electrons materials: (left panel) Kapton®WS,
(central panel) PET samples, Melinex®454 and Mylar®M021, and (right panel) PI/POSS material, Thermalbright®N and liquid crystal polymer, Zenite®
Fig. 6 FTIR spectra of pristine (black curves) and irradiated (red curves) materials showing transmittance as
a function of wavenumber for the range (400 - 2000) 1/cm. (a) Kapton®WS, (b) Zenite®, (c) Mylar®M021, (d)
Melinex®454, and (d) Thermalbright®N.
B. AO exposure
Exposure of polymer materials to AO at 8 km/s is sufficient to break the polymer bonds and induce oxidative
decomposition, resulting in substantial erosion of polymer surfaces which manifests itself as the mass loss, thinning, and
texture roughening. The rough texture results in degradation of optical properties, as illustrated by Fig. 8. Post-irradiated
PET samples (Mylar®M021 and Melinex®454) as well as LCP sample (Zenite®) show hazy white surface which is
result of O-atom erosion leading to the microscopic roughness of these samples. Kapton®WS and Thermalbright®N
materials retained their pristine-like visual appearance suggesting the less AO-damaged, if any, surface.
Mass loss and the relative (to Kapton®H) erosion rate were evaluated and are presented in Table 4. The erosion
rates for the samples ranged from .17 to 1.70, relative to the erosion rate of the Kapton®H witness samples.
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Fig. 7 HCRF measured for pristine and irradiated with maximum electron fluence (a) Kapton®WS, (b)
Zenite®, (c) Mylar®M021, (d) Melinex®454, and (d) Thermalbright®N materials with fixed viewing angle (30◦ )
and variable illumination angles.
Kapton
WS
Thermalbright
Melinex
Mylar
454
N
MO21
Zenite
Fig. 8 Photographs of pristine (left sample on each picture) and exposed to AO (right sample on each picture)
materials: (top left panel) Kapton®WS, (top central panel) Melinex®454, (top right panel) Mylar®M021, and
(right panel) PI/POSS material, (bottom left panel) Thermalbright®N, and (bottom right panel) liquid crystal
polymer, Zenite®
.
IV. Discussion
A. PI (Kapton®WS)
Visual inspection of post electron-irradiated material revealed no visible color change. Similarly, no surface color
degradation after AO-exposure was observed. Due to the small mass of impinging electrons, no significant roughening
of surface was expected after electron irradiation. However, electron-irradiated material appeared smoother compared
to its pristine state (R𝑎 of 28.7 nm and 72.3 nm, respectively). Similarly, Mishra et al. [20] reported that irradiation of
8
Table 4
Mass loss and the relative (to Kapton®H) erosion rate of AO-irradiated materials.
Material
Kapton®WS Glossy
Zenite®
Melinex®454
Mylar®M021
Thermalbright™N
Mass loss (mg)
1.71
2.12
8.01
7.89
1.23
Erosion rate
0.22
0.47
1.64
1.70
0.17
PI with high energy (2 MeV) electrons resulted in reduced mean surface roughness of irradiated material. Increased
level of cross-linking in irradiated PI [8] may be responsible for the overall surface flattening.
Alike other PI films [19], conductivity of Kapton®WS increased after electron irradiation. The observed bulk
conductivity increase in irradiated PI may be attributed to the radiation-induced generation of electron hopping sites and
the formation of extended pi-bonded chemical structures which are not present in the pristine material.
Overall reflectance of Kapton®WS material reduces with high energy electrons irradiation in 380 - 860 nm range
and stays unchanged afterwards. Sharp absorption edge around 380 nm demonstrated by the material correlates with
its visibly white appearance. The nearly zero reflectance in shorter wavelength region (below 380 nm) exhibited by
both pristine and damaged PIs is associated with absorption due to the 𝜋 - 𝜋 ∗ transitions in the aromatic groups [21].
Absorption feature at 1650 nm may be attributed to CH3 in the first overtone [22]
The FTIR spectrum of irradiated PI did not reveal any additional bands; however, increase of absorption intensity
of overall spectrum was observed. Since every minima position in transmission FTIR spectrum is fundamental to
molecular bonding structure or functional group existing in the infrared active material, no new functional groups as
well as no change in bond strength or bond angle was observed after irradiation. Variation in intensity correlates to
the proportion of that functional group present in the material, so increased overall absorption intensity of irradiated
spectrum associates with the increased degree of radiation-induced cross linking. As was reported earlier [8], increased
degree of crosslinking in electron-irradiated PI is likely due to the formation of new inter-chain C–N bonds.
HCRF measurements of irradiated material demonstrated significant reduction of signal intensity, especially at -20◦
and -40◦ illumination angles (the viewing angle was fixed and corresponds to 30◦ ).
B. LCP (Zenite®)
No electron-induced optical degradation was observed; AO-exposure caused the hazy white finish of the treated
material, attributed to the surface morphology modification due to the O-erosion. Surface roughness of electron-irradiated
Zenite®material was increased compared to its pristine state (22.6 nm and 14.9 nm, respectively). Electron bombardment
induced the increased electrical resistivity of the Zenite®. Reflectance values of the irradiated Zenite®were overall
decreased in (300 - 1600) nm region. FTIR spectrum demonstrated significantly increased overall transmission in (400 2000) 1/cm region suggesting, along with the increased surface roughness, increased disorder of polymer chains in the
LCP structure. HCRF measurements of irradiated material demonstrated significant reduction of signal intensity at all
illumination angles (the viewing angle is fixed and corresponds to 30◦ ).
C. PET (Melinex®454, Mylar®M021)
Whereas electron bombardment produced no visible color change of both PET samples, AO-exposure significantly
degraded surface quality of of Mylar®M021 and Melinex®454, resulting in their hazy white appearance. This is
attributed to the formation of so-called "rag" structure caused by the O-erosion of the polymers [23]. Surface roughness
of both PET materials was reduced after electron irradiation. Earlier, Chinaglia et al. [24] and Lee et al. [25]
demonstrated that after low-energy (0-20 keV) electron irradiation the PET material acquires the enhanced surface
smoothness. Similar observations were reported by Mishra et al [20] after irradiation of PET samples with high energy
(2 MeV) electrons.
It is interesting to note that resistivity of the electron-irradiated PET samples either increased (Mylar®M021) or did
not change significantly (Melinex®454). Similar phenomenon was observed by Chaudhary et al. [26] and Oproiu et al.
[27]. They attributed this phenomenon to the increased crosslinking of the PET chains due to the electron irradiation,
which in turn may obstruct the charge carrier from hopping from one chain to another chain resulting in decrease of
9
electrical conductivity.
Melinex®454 demonstrated the drastic degradation of electron-irradiated material reflectance in the region below
835 nm. The same trend, whereas not so prominent, was demonstrated by the irradiated Mylar®M021 sample. Another
important feature of the Melinex®454 material caused by irradiation is the appearance of intense absorption band at
1650 nm. The Mylar®M021, irradiated to the maximum fluence, revealed a slight (3%) increase of the reflectance
beyond 850 nm. Finally, additional radiation -induced absorption features appeared in both materials; at 500 nm and
600 nm for Melinex®454, and at 500 nm for Mylar®M021.
FTIR transmission spectrum of electron-irradiated Melinex®454 did not show any changes compared to the prsitine
material. The irradiated FTIR spectrum of Mylar®M021 demonstrated slightly increased overall transmission, with
increased intensity of absorption at 631 1/cm, characteristic for the methylene group [28].
HCRF measurements of irradiated material demonstrated reduction of signal intensity at all illumination angles for
Mylar®M021 sample. Similar trend was observed for Melinex®454, except the HCRF signal at -40◦ illumination angle,
which revealed approximately factor of 10 higher signal. The viewing angle was fixed and corresponds to 30◦ .
D. PI/POSS (Thermalbright®N)
The Thermalbright®N material demonstrated no visible degradation of optical surface quality after electron
irradiation and a pristine-like appearance after the AO-exposure. The enhanced durability of PI/POSS under AO
exposure is a result of a silica passivation layer formation on top of the degraded organic material (PI) during the
AO-bombardment [29]. This silica layer protects the underlying polymer from further degradation resulting in the lowest
erosion rate of the Thermalbright®N material among the studied samples. Electron irradiation enhanced smoothness of
the material; average surface roughness was 52.5 nm for the pristine and 18.1 nm for the irradiated films, respectively.
While Thermalbright®N is referred as a white polyimide, unlike polyimides from the Kapton ®family it became
significantly more resistive after electron irradiation, with volume resistivity of irradiated material being a factor of
1000 higher compared to the pristine one.
DHR spectrum of the electron-irradiated POSS material showed no change of reflectance with irradiation to the
maximum electron fluence. Comparison of FTIR spectra of pristine and electron-irradiated Thermalbright®N revealed
radiation-induced increased absorption in (670 - 700) 1/cm region and at 1900 1/cm. HCRF measurements of irradiated
material demonstrated the overall reduction of signal intensity at all illumination angles .
V. Conclusions
Thorough physical and chemical characterization of novel and heritage spacecraft materials during the simulated
space weather experiments is important for establishment of correlation factors between true space-exposure and
accelerated space weather experiments at ground facilities as well as enabling accurate prediction of on-orbit material
performance based on laboratory-based testing. This work focuses on characterization of material properties of selected
spacecraft materials under independent LEO-simulated electron irradiation and AO exposure.
VI. Acknowledgements
Authors gratefully acknowledge Physical Sciences, Inc., especially Drs. Daniel M. Hewett and David B. Oakes,
for conducting AO-exposure experiments. Authors also would like to thank DuPont de Nemours, Inc, for providing
polyimide materials for this research. This work was partially supported by Air Force Office of Scientific Research,
Remote Sensing and Imaging Physics Portfolio (Dr. Michael Yakes) Grant 20RVCOR024 and Georgia Tech Research
Institute Independent Research and Development (IRAD) program.
References
[1] McKnight, D., Witner, R., Letizia, F., Lemmens, S., Anselmo, L., Pardini, C., Rossi, A., Kunstadter, C., Kawamoto, S., Aslanov,
V., Perez, J.-C. D., Ruch, V., Lewis, H., Nicolls, M., Jing, L., Dan, S., Dongfang, W., Barranov, A., and Grishko, D., “Identifying
the 50 statistically-most-concerning derelict objects in LEO,” Acta Astronautica, Vol. 181, 2021, pp. 282–291.
[2] Reyes, J. A., and Cowardin, H. M., “Spectral characterization of spacecraft materials used in hypervelocity impact testing,”
Algorithms, Technologies, and Applications for Multispectral and Hyperspectral Imaging XXVII, Vol. 11727, International
Society for Optics and Photonics, 2021, p. 117271G.
10
[3] Gordo, P., Frederico, T., Melício, R., Duzellier, S., and Amorim, A., “System for space materials evaluation in LEO environment,”
Advances in Space Research, Vol. 66, No. 2, 2020, pp. 307–320.
[4] Lu, Y., Shao, Q., Yue, H., and Yang, F., “A review of the space environment effects on spacecraft in different orbits,” IEEE
Access, Vol. 7, 2019, pp. 93473–93488.
[5] deGroh, K. K., Dever, J. A., Jaworske, D. A., Miller, S. K., Sechkar, E. A., Panko, S. R., et al., “NASA Glenn research center’s
materials international space station experiments (MISSE 1-7),” 2008.
[6] de Groh, K. K., and Banks, B. A., “Atomic Oxygen Interactions and the Design of MISSE-9 and MISSE-10 Experiments,” 2018.
[7] Plis, E., Engelhart, D. P., Barton, D., Cooper, R., Ferguson, D., and Hoffmann, R., “Degradation of polyimide under exposure to
90 keV electrons,” physica status solidi (b), Vol. 254, No. 7, 2017, p. 1600819.
[8] Rahnamoun, A., Engelhart, D. P., Humagain, S., Koerner, H., Plis, E., Kennedy, W. J., Cooper, R., Greenbaum, S. G., Hoffmann,
R., and van Duin, A. C., “Chemical dynamics characteristics of Kapton polyimide damaged by electron beam irradiation,”
Polymer, Vol. 176, 2019, pp. 135–145.
[9] Alegaonkar, P., Bhoraskar, V., Balaya, P., and Goyal, P., “Dielectric properties of 1 MeV electron-irradiated polyimide,” Applied
physics letters, Vol. 80, No. 4, 2002, pp. 640–642.
[10] Mishra, R., Tripathy, S., Dwivedi, K., Khathing, D., Ghosh, S., Müller, M., and Fink, D., “Spectroscopic and thermal studies of
electron irradiated polyimide,” Radiation measurements, Vol. 36, No. 1-6, 2003, pp. 621–624.
[11] White, T. J., and Broer, D. J., “Programmable and adaptive mechanics with liquid crystal polymer networks and elastomers,”
Nature materials, Vol. 14, No. 11, 2015, pp. 1087–1098.
[12] Campo, E. A., Complete Part Design Handbook, Hanser Germany, 2006.
[13] Gupta, S., “Application of Liquid Crystals in Space Activities,” Frontiers of Polymer Research, Springer, 1991, pp. 189–193.
[14] Brunsvold, A. L., Minton, T. K., Gouzman, I., Grossman, E., and Gonzalez, R., “An investigation of the resistance of polyhedral
oligomeric silsesquioxane polyimide to atomic-oxygen attack,” High Performance Polymers, Vol. 16, No. 2, 2004, pp. 303–318.
[15] Cooper, R., and Hoffman, R., “Jumbo space environment simulation and spacecraft charging chamber characterization,” Tech.
rep., Air Force Research Laboratory, Space Vehicles Directorate Kirtland AFB . . . , 2015.
[16] Engelhart, D. P., Plis, E., Humagain, S., Greenbaum, S., Ferguson, D., Cooper, R., and Hoffmann, R., “Chemical and electrical
dynamics of polyimide film damaged by electron radiation,” IEEE Transactions on Plasma Science, Vol. 45, No. 9, 2017, pp.
2573–2577.
[17] Bengtson, M., Maxwell, J., Hoffmann, R., Cooper, R., Schieffer, S., Ferguson, D., Johnston, W. R., Cowardin, H., Plis, E.,
and Engelhart, D., “Optical characterization of commonly used thermal control paints in a simulated GEO environment,” The
Advanced Maui Optical and Space Surveillance Technologies Conference, 2018, p. 33.
[18] Frederickson, A. R., and Dennison, J., “Measurement of conductivity and charge storage in insulators related to spacecraft
charging,” IEEE Transactions on Nuclear Science, Vol. 50, No. 6, 2003, pp. 2284–2291.
[19] Plis, E. A., Engelhart, D. P., Cooper, R., Johnston, W. R., Ferguson, D., and Hoffmann, R., “Review of radiation-induced effects
in polyimide,” Applied Sciences, Vol. 9, No. 10, 2019, p. 1999.
[20] Mishra, R., Tripathy, S., Dwivedi, K., Khathing, D., Ghosh, S., Müller, M., and Fink, D., “Modification in etching characteristics
and surface topography of some electron irradiated polymers,” Radiation measurements, Vol. 34, No. 1-6, 2001, pp. 95–98.
[21] Ishida, H., Wellinghoff, S. T., Baer, E., and Koenig, J. L., “Spectroscopic studies of poly [N, N’-bis (phenoxyphenyl)
pyromellitimide]. 1. Structures of the polyimide and three model compounds,” Macromolecules, Vol. 13, No. 4, 1980, pp.
826–834.
[22] Burke, M., Dawson, C., Allen, C. S., Brum, J., Roberts, J., and Krekeler, M. P., “Reflective spectroscopy investigations of
clothing items to support law enforcement, search and rescue, and war crime investigations,” Forensic science international,
Vol. 304, 2019, p. 109945.
[23] Tennyson, R., “Atomic oxygen effects on space inflatable materials,” Progress in Astronautics and Aeronautics., Vol. 191, 2001,
pp. 281–302.
11
[24] Chinaglia, D., Constantino, C., Aroca, R., and Oliveira Jr, O., “Surface modifications on Teflon FEP and Mylar C induced by a
low energy electron beam: A Raman and FTIR spectroscopic study,” Molecular Crystals and Liquid Crystals, Vol. 374, No. 1,
2002, pp. 577–582.
[25] Lee, E., Lewis, M., Blau, P., and Mansur, L., “Improved surface properties of polymer materials by multiple ion beam treatment,”
Journal of Materials Research, Vol. 6, No. 3, 1991, pp. 610–628.
[26] Chaudhary, N., Koiry, S., Singh, A., Tillu, A., Jha, P., Samanta, S., Debnath, A., Aswal, D., Mondal, R., Acharya, S., et al.,
“Electron beam induced modifications in flexible biaxially oriented polyethylene terephthalate sheets: Improved mechanical and
electrical properties,” Materials Chemistry and Physics, Vol. 189, 2017, pp. 237–244.
[27] Oproiu, C., Martin, D., Toma, M., Marghitu, S., and Jianu, A., “Transitory and permanent effects of electron beam irradiation
on insulating materials,” Nuclear Instruments and Methods in Physics Research Section B: Beam Interactions with Materials
and Atoms, Vol. 166, 2000, pp. 669–675.
[28] Dworecki, K., Hasegawa, T., Sudlitz, K., and Wasik,
˛
S., “Modification of electrical properties of polymer membranes by ion
implantation,” Nuclear Instruments and Methods in Physics Research Section B: Beam Interactions with Materials and Atoms,
Vol. 166, 2000, pp. 646–649.
[29] Minton, T. K., Wright, M. E., Tomczak, S. J., Marquez, S. A., Shen, L., Brunsvold, A. L., Cooper, R., Zhang, J., Vij, V.,
Guenthner, A. J., et al., “Atomic oxygen effects on POSS polyimides in low earth orbit,” ACS applied materials & interfaces,
Vol. 4, No. 2, 2012, pp. 492–502.
12
JOURNAL OF SPACECRAFT AND ROCKETS
Vol. 60, No. 2, March–April 2023
Downloaded by GA Inst of Technology on May 9, 2023 | http://arc.aiaa.org | DOI: 10.2514/1.A35502
Ground Testing of the 16th Materials
International Space Station Experiment Materials
Elena A. Plis∗
Georgia Tech Research Institute, Atlanta, Georgia 30318
Miles T. Bengtson†
Kirtland Air Force Base, Albuquerque, New Mexico 87117
Daniel P. Engelhart‡
University of New Mexico, Albuquerque, New Mexico 87131
Gregory P. Badura§
Georgia Tech Research Institute, Atlanta, Georgia 30318
Heather M. Cowardin¶
NASA Johnson Space Center, Houston, Texas 77058
Jacqueline A. Reyes**
University of Texas at El Paso, El Paso, Texas 79968
Ryan C. Hoffmann,†† Alexey Sokolovskiy,†† and Dale C. Ferguson††
U.S. Air Force Research Laboratory, Kirtland Air Force Base, Albuquerque, New Mexico 87117
Jainisha R. Shah‡‡ and Sydney Collman§§
Assurance Technology Corporation, Carlisle, Massachusetts 01741
and
Timothy R. Scott¶¶
DuPont de Nemours, Inc., Durham, North Carolina 27703
https://doi.org/10.2514/1.A35502
External spacecraft materials play an important role in satellite protection from the harsh space environment.
Research has shown that the physical, chemical, and optical properties of matter change continuously as a result of
exposure to solar radiation and aggressive chemical species produced in Earth’s upper atmosphere. Thorough
knowledge of the material properties’ evolution throughout a planned mission lifetime helps to improve the reliability
of spacecraft. Moreover, the establishment of correlation factors between true space exposure and accelerated space
weather experiments at ground facilities enables accurate prediction of on-orbit material performance based on
laboratory-based testing. The presented work aims to evaluate the radiation effects of a low-Earth-orbit environment,
namely, exposure to the high-energy electrons and atomic oxygen (AO) of heritage and novel spacecraft material
selection. The studied materials represent the “flight duplicates” of samples that are launched as a part of the 16th
Materials International Space Station Experiment Flight Facility (MISSE-FF) mission in 2022.
I.
aggressive chemical species produced in Earth’s upper atmosphere
[1–4]. The motivation of the research reported here is the natural
curiosity of a team of scientists for fundamental understanding of the
interactions between the harsh space environment and spacecraft
materials. As humankind’s forays into space become simultaneously
more commonplace and ambitious, the need for this understanding
becomes ever more important. Furthermore, novel lightweight materials must be developed with improved long-term radiation shielding
and mechanical properties for use in internal and external spacecraft
systems. At low Earth orbit (LEO), the most destructive agents are
atomic oxygen (AO) and unfiltered vacuum ultraviolet (VUV) radiation. Combined with charged particles (electrons) irradiation, AO and
VUV radiation may significantly degrade the outer spacecraft materials, thus drastically reducing the lifespan of the spacecraft.
To date, the effect of LEO space weather exposure on the durability and erosion of different materials was studied in both laboratory
space-simulated [5–9] and true space environments [10]. In particular,
the interaction of LEO space weather, comprising AO/VUV radiation
and some high-energy electron exposure, with a wide variety of
spacecraft materials (including thermal control paints and organic
polymers) has been reported. However, through vetting of novel
materials under space environment is necessary and lacking. Independently, the observation of spectral changes in materials exposed
to the space environment, including a “reddening” (i.e., decrease in
optical reflectance with decreasing wavelength) of the spectra, has
also been reported, as well as potential schemes to use this feature for
spacecraft material identification (e.g., Ref. [11]). Still, no dynamic
Introduction
E
XTERIOR spacecraft materials play an important role in satellite protection from the strident space environment. Research has
shown that the physical, chemical, and optical properties of matter
change continuously as a result of exposure to solar radiation and
Received 30 June 2022; revision received 7 November 2022; accepted for
publication 27 November 2022; published online 10 January 2023. Copyright
© 2022 by the American Institute of Aeronautics and Astronautics, Inc. All
rights reserved. All requests for copying and permission to reprint should be
submitted to CCC at www.copyright.com; employ the eISSN 1533-6794 to
initiate your request. See also AIAA Rights and Permissions www.
aiaa.org/randp.
*Senior Research Engineer, 925 Dalney Street NW; also Assurance
Technology Corporation, Carlisle, Massachusetts 01741
†
Postdoctoral Fellow, National Research Council Research Associateship
Program, U.S. Air Force Research Laboratory.
‡
Research Professor, Department of Chemistry and Chemical 300 Terrace
Street NE.
§
Research Scientist, 925 Dalney Street NW.
¶
Laboratory and In Situ Lead Orbital Debris Business Unit Manager,
Orbital Debris Program Office, 2101 NASA Parkway.
**500 W. University Avenue.
††
Space Vehicles Directorate.
‡‡
Research Scientist, Assurance Technology Corporation, Carlisle,
Massachusetts 01741; shah@assurtech.com.
§§
Research Scientist, Assurance Technology Corporation, Carlisle,
Massachusetts 01741; collman@assurtech.com.
¶¶
Business Development Leader, Aerospace and Defense.
385
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386
PLIS ET AL.
changes in material properties under exposure to the complex LEO
environment were studied as well as the utilization of these changes
for remote spacecraft material diagnosis.
The presented work reports on the progress our team made so far
by studying the effect of space-simulated low-Earth-orbital environment on the optical, charge transport, and surface morphology properties of the selected spacecraft polymers. A ground-based simulation
of the space environment is experimentally challenging because the
space environment varies wildly depending on the orbit, solar conditions, and many other factors [12]. Simultaneous exposure of the
materials to the charge particles, VUV radiation, and/or temperature
may be hindered by different acceleration factors for particles and
VUV flux, leading to significantly longer VUV irradiation times [13].
Oppositely, once a sequential exposure approach is used, the materials are exposed to charging particles (protons, electrons), AO, and
VUV radiation in successive steps. Thus, degradation induced by
each environmental component may be monitored separately. In this
work, we use a space-simulated LEO environment comprising, separately, high-energy electrons and AO exposure. Furthermore, to
correlate the evolution of material properties to those throughout a
mission lifetime, we will employ the spectral reflectivity changes of
the same materials induced by exposure to different components of
the LEO environment during the six-month-long 16th Materials
International Space Station Experiment Flight Facility (MISSE-FF)
mission.
II.
Experimental Details
A. Materials
The studied sample collection comprises several different class of
polymers, including polyimides (PIs), polyethylene terephthalate
(PET) materials, liquid crystal polymers (LCPs), PI/polyhedral oligomeric silsesquioxane (POSS) materials, carbon-fiber-reinforced
polymers (CFRPs), and glass-reinforced polymers (GFRPs). Table 1
lists the materials and the material classes that they belong to, and
Fig. 1 demonstrates the visual appearance of the materials in their
pristine states.
5.5 h. Details of the electron irradiation procedure were reported
elsewhere [16].
AO exposure was performed using the FAST source in according
to ASTM-E2089-15 [17]. The effective peak atomic oxygen fluence
during the exposure was 3.1 × 1020 O∕cm2 , which corresponds to
six weeks of LEO exposure. Details of the AO-irradiation procedure
may be found in Ref. [18].
C. Characterization Methods
Each material in its pristine and irradiated state was subjected to a
characterization protocol; this is schematically illustrated in Fig. 2
using an electron-irradiated Mylar® MO21 sample as an example.
In particular, surface properties were accessed by the atomic force
microscopy (AFM) and scanning electron microscopy techniques;
optical properties were characterized using the hemispherical conical
reflectance factor (HCRF), the laboratory approximation of true
bidirectional reflectance distribution function measurements, UV/
visible transmission and reflectance measurements, directional hemispherical reflectance (DHR) measurements, and Fourier-transform
infrared (FTIR) spectroscopic measurements. Furthermore, the effect
of space-simulated exposure on charge transport properties was
evaluated with direct-current ASTM D-257 [19] and the surface
potential decay (SPD) [20] methods. In addition, mass loss of the
AO-exposed materials was estimated. This work will focus on spaceweather-induced changes of surface morphology and the optical
behavior of all studied materials measured by DHR (for electronirradiated materials) and UV/visible spectroscopy (for AO-exposed
materials).
The surface morphology and roughness of the studied materials
were examined using Bruker’s Dimension Icon AFM, allowing
measurement of the surface roughness up to 5 μm on areas as large
as 200 × 200 μm. The DHR measurements of pristine and electronirradiated samples were performed before and during the electron
irradiation procedure according to the optical data acquisition procedure reported elsewhere [21]. Transmission measurements of the
AO-exposed materials were performed with a Cary 5000 UV-visible/
near-infrared spectrometer.
B. Irradiation Procedure
High-energy electron and AO exposure were performed sequentially in the jumbo space irradiation chamber at the Spacecraft
Charging and Instrument Calibration Laboratory (SCICL) [14] and
at Physical Sciences, Inc. Materials were bombarded with highenergy (100 keV) electrons produced by a monoenergetic Kimball
Physics EG8105-UD electron flood gun. According to Ref. [15], the
mean annual electron flux (greater than 100 keV, electrons with
energies greater than 100 keV, orbit averaged) experienced by the
International Space Station is ∼1013 electrons∕cm2 ∕s. The maximum electron fluence the materials under investigation were exposed
to was 8.5 × 1013 electrons∕cm2 , which corresponds to approximately 10 years at a 600–800 km (LEO) orbit delivered during
Table 1
No.
1
2
3
4
5
6
7
8
9
10
11
12
13
14
a
D. Results and Discussion
The average surface roughness (Ra ) values of pristine, irradiated
with high-energy electrons, and AO-exposed materials are summarized in Table 1. Ra values are the average of several 5 × 5 μm scans
taken at different parts of the sample. Note that the roughness of the
CFRP material was too large, even in its pristine form, to be measured
reliably with the AFM technique.
Most of the electron-irradiated materials from the Kapton® family
demonstrated a significant reduction of Ra roughness (HN, CS, WS,
XC, DR9) after electron irradiation. Kapton CR did not change, and
only one PI material (Kapton TF) showed surface roughening after
electron bombardment. The surface roughness of PET materials also
Roughness values of pristine, irradiated with high-energy electrons, and AO-exposed materialsa
Trade name
Kapton CR
Kapton CS
Kapton WS
Kapton TF
200DR9
Kapton HN
Economyplate™
G-10/FR4 glass epoxy
Zenite
Melinex 454
Mylar MO21
CORIN XLS
Thermalbright N
Kapton XC
N/A denotes “not applicable.”
Category
PI
PI
PI
PI
PI
PI
CFRP
GFRP
LCP
PET
PET
POSS
POSS
PI
Pristine Ra , nm
26.7
5.5
72.3
5.5
26.7
11.4
N/A
12.2
14.9
8.2
7.8
9.4
52.5
56.2
E-irradiated Ra , nm
25.6
2.0
32.4
12.3
4.6
2.5
N/A
4.6
25.1
5.0
3.6
2.2
47.7
34.5
AO-exposed Ra , nm
62.9
15.4
20.1
136.3
111.0
91.0
N/A
218.0
71.9
116.7
187.2
14.3
74.6
161.0
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PLIS ET AL.
Fig. 1
387
Studied sample collection comprising several different classes of polymers, as shown in Table 1.
Fig. 2 Illustration of material properties characterization protocol used for pristine and irradiated materials using electron-irradiated Mylar MO21
sample as an example. Here, GTRI stands for the Georgia Tech Research Institute.
Downloaded by GA Inst of Technology on May 9, 2023 | http://arc.aiaa.org | DOI: 10.2514/1.A35502
388
PLIS ET AL.
Fig. 3
DHR measurements of transparent and opaque materials from the studied material selection irradiated with high-energy electrons.
reduced after irradiation. The surface of the LCP (Zenite®), similar to
Kapton TF, revealed increased Ra after irradiation. Both POSS materials became smoother after electron exposure; however, CORIN®
XLS demonstrated a roughness reduction factor of 4.3, whereas the
surface roughness of Thermalbright® N changed slightly. Finally, the
GFRP material was factor of three smoother after electron irradiation.
Exposure of the polymer materials to AO at 8 km/s is sufficient to
break the polymer bonds and induce oxidative decomposition, resulting
in substantial erosion of polymer surfaces, which manifests itself as
mass loss, thinning, and texture roughening. AO exposure clearly
affected the surface roughness of the studied materials. All polymers
from the Kapton family, excluding Kapton WS, revealed a matte surface, which is the result of O-atom erosion, leading to the microscopic
roughness of these samples. Postexposed Kapton WS retained the
pristinelike visual appearance with reduced Ra roughness as compared
to its pristine state. Both PET samples (Mylar MO21 and Melinex®
454) were severely degraded after AO exposure, with Ra roughness
increasing by more than a factor of 10 as a result of AO erosion. The
AO-induced damage of the LCP, GFRP, and CFRP materials was also
easily observed. The AO-exposed surfaces of these samples became
less glossy. Finally, the Ra roughness of two POSS materials, CORIN
XLS and Thermalbright N, increased insignificantly, suggesting good
resistance of these materials to AO-induced erosion.
The results of in situ DHR measurements on transparent and opaque
materials are summarized in Fig. 3. All measured materials demonstrated that the reflectance changes with electron exposure. It is
interesting that degradation of both studied POSS materials was
minimal, suggesting good resistance of the POSS films from electron
bombardment. In Fig. 4, the observed DHR data trends are summarized for every material in more visual form. In the shorter-wavelength
region, materials apart from the POSS revealed the reduction of the
signal. Both POSS materials, as mentioned earlier, did not change
considerably. Generally, no signal degradation was observed beyond
12 nm. It is interesting that one of the PET materials, Mylar MO21,
became more reflective in the longer-wavelength range.
The results of the UV/visible transmission measurements of
pristine and AO-exposed transparent and opaque materials from
the studied material selection are presented in Fig. 5.
Fig. 4 Observed DHR data trends of electron-irradiated transparent
and opaque materials from the studied material selection.
The light transmission characteristics of all studied materials
considerably degraded after AO exposure. The measurable (i.e.,
transparent enough) materials from the Kapton family (HN, TF,
CR, and CS) demonstrated a reduction of transmission. The Kapton WS, Thermalbright N, GFRP, and Zenite materials have low
light transmittance in their pristine states (0.1–0.2%); after AO
exposure, their appearances changed to completely opaque. Transmission of both PET samples degraded considerably after AO
exposure. CORIN XLS, as expected, was not affected be the AO
exposure. The transmission curve of this material did not show
significant changes in the 400–2000 nm wavelength region.
However, transmission of the AO-exposed material increased
around to 300 nm. It should be noted that all measured curves
possessed a kink at 800 nm, which was an instrumental artifact
corresponding to the change of detectors in Cary 5000 during the
measurements.
E. MISSE-FF Materials Integration
The identical material set will be installed on each face of the
MISSE-FF to learn about the separate effects of LEO environmental
Downloaded by GA Inst of Technology on May 9, 2023 | http://arc.aiaa.org | DOI: 10.2514/1.A35502
PLIS ET AL.
389
Fig. 5 UV/visible transmission measurements of pristine and AO-exposed transparent and opaque materials from the studied material selection.
components on these spacecraft materials: in particular, ram (direction
of travel) with dominant AO exposure, zenith (direction away from
Earth) with dominant VUVexposure, and wake (opposite to the travel
direction) with high-energy electron exposure. To closely monitor
the optical changes of materials during the mission, we proposed to
upgrade the MISSE-FF with different illumination schemes, including
white and IR light-emitting-diode illumination sources as well as a
high-resolution camera to record Red Green Blue (RGB) and IR color
data. This upgraded hardware will stay on board the MISSE-FF and be
available for future scientific work.
Fig. 6 MISSE-FF camera board assembly (left) and sample image delivered by the Basler daA1600 camera (right). Image credit: U.S. Air Force Research
Laboratory.
390
PLIS ET AL.
The 16th Materials International Space Station Experiment’s
(MISSE-16’s) flight duration will be approximately six months;
a daily image of each sample will be taken for the first 30 days,
a weekly image of each sample will be taken for the following two
months, and monthly images will be taken for the remaining duration
of the mission. Thus, we will monitor the aging of the materials,
which will manifest themselves as the reflectivity changes; this will
occur almost in real time during the initial stage of the LEO exposure,
and then very closely during the rest of the mission.
The MISSE-FF materials integration along with camera board
integration for the MISSE-16 mission were performed on September
2021. Identical MISSE-FF carriers were prepared for the ram, zenith,
and wake faces of the MISSE-FF. Many of our materials were
mounted in multiple layers to get the average thickness of 10 mils.
Figure 6 shows the camera board assembly and the sample image
delivered by the Basler daA1600 camera.
Downloaded by GA Inst of Technology on May 9, 2023 | http://arc.aiaa.org | DOI: 10.2514/1.A35502
III.
Conclusions
Thorough characterization of the material properties of different
spacecraft materials during the simulated space-weather experiments
is important for the establishment of correlation factors between true
space exposure and accelerated space-weather conditions at ground
facilities, as well as enabling accurate prediction of on-orbit material
performance based on laboratory-based testing. This work focuses on
the characterization of the surface morphology and optical properties
of selected spacecraft materials under independent LEO-simulated
electron irradiation and AO exposure.
Acknowledgments
This work was partially supported by the U.S. Air Force Office of
Scientific Research, Remote Sensing and Imaging Physics Portfolio
(Michael Yakes) grant 20RVCOR024 and the Georgia Tech Research
Institute Independent Research and Development program. The
authors gratefully acknowledge Physical Sciences, Inc., especially
Daniel M. Hewett and David B. Oakes, for conducting atomic
oxygen-exposure experiments. The authors also would like to thank
DuPont de Nemours, Inc., for providing polyimide materials for this
research. Part of this work will be presented at the 2023 AIAA
SciTech Forum, 23–27 January 2023. The views expressed are those
of the authors and do not necessarily reflect the official policy or
position of the U.S. Department of the U.S. Air Force, the U.S.
Department of Defense, or the U.S. Government. The appearance of
external hyperlinks does not constitute endorsement by the U.S.
Department of Defense (DOD) of the linked websites or the information, products, or services contained therein. The DOD does not
exercise any editorial, security, or other control over the information
you may find at these locations. Trade names and trademarks are used
in this paper for identification only. Their usage does not constitute an
official endorsement, either expressed or implied, by the National
Aeronautics and Space Administration.
References
[1] Dooling, D., Material Selection Guidelines to Limit Atomic Oxygen Effects
on Spacecraft Surfaces, NASA Marshall Space Flight Center, 1999.
[2] de Groh, K. K., and Banks, B. A., “MISSE 2 PEACE Polymers Erosion
Morphology Studies,” Proceedings of the International Symposium on
Materials in a Space Environment (ISMSE-11), Aix-en-Provence,
France, Sept. 2009.
[3] Miller, S. K., and Dever, J. A., “Materials International Space Station
Experiment 5 Polymer Film Thermal Control Experiment,” Journal of
Spacecraft and Rockets, Vol. 48, No. 2, 2011, pp. 240–245.
https://doi.org/10.2514/1.49482
[4] Ouchen, F., Aga, R., Harvey, M., and Heckman, E., “Space Survivability
for Printed Electronics Applications,” Flexible and Printed Electronics,
Vol. 6, No. 1, 2021, Paper 015012.
https://doi.org/10.1088/2058-8585/abe654
[5] Tagawa, M., and Yokota, K., “Atomic Oxygen-Induced Polymer Degradation Phenomena in Simulated LEO Space Environments: How Do
Polymers React in a Complicated Space Environment?” Acta Astronautica, Vol. 62, Nos. 2–3, 2008, pp. 203–211.
https://doi.org/10.1016/j.actaastro.2006.12.043
[6] Awaja, F., Moon, J. B., Zhang, S., Gilbert, M., Kim, C. G., and Pigram,
P. J., “Surface Molecular Degradation of 3D Glass Polymer Composite
Under Low Earth Orbit Simulated Space Environment,” Polymer
Degradation and Stability, Vol. 95, No. 6, 2010, pp. 987–996.
https://doi.org/10.1016/j.polymdegradstab.2010.03.013
[7] Jin, S., Son, G., Kim, Y., and Kim, C.-G., “Enhanced Durability of
Silanized Multi-Walled Carbon Nanotube/Epoxy Nanocomposites
Under Simulated Low Earth Orbit Space Environment,” Composites
Science and Technology, Vol. 87, Oct. 2013, pp. 224–231.
https://doi.org/10.1016/j.compscitech.2013.08.017
[8] Tonon, C., Duvignacq, C., Teyssedre, G., and Dinguirard, M., “Degradation of the Optical Properties of ZnO-Based Thermal Control Coatings in Simulated Space Environment,” Journal of Physics D: Applied
Physics, Vol. 34, No. 1, 2001, Paper 124.
https://doi.org/10.1088/0022-3727/34/1/319
[9] Dworak, D. P., Banks, B. A., Karniotis, C. A., and Soucek, M. D.,
“Evaluation of Protective Silicone/Siloxane Coatings in Simulated
Low-Earth-Orbit Environment,” Journal of Spacecraft and Rockets,
Vol. 43, No. 2, 2006, pp. 393–401.
https://doi.org/10.2514/1.15787
[10] Dever, J., Banks, B., de Groh, K., and Miller, S., “Degradation of
Spacecraft Materials,” Handbook of Environmental Degradation of
Spacecraft Materials, edited by M. Kutz, William Andrew Publ., Norwich, NY, 2005, pp. 465–501.
https://doi.org/10.1016/B978-081551500-5.50025-2
[11] Abercromby, K., Okada, J., Guyote, M., Hamada, K., and Barker, E.,
“Comparisons of Ground Truth and Remote Spectral Measurements of
the Formosat and Ande Spacecraft,” Proceedings of the 2006 AMOS
Technical Conference, Maui Economic Development Board Inc Kihei,
Maui, HI, Sept. 2006.
[12] Gordo, P., Frederico, T., Melcio, R., Duzellier, S., and Amorim, A.,
“System for Space Materials Evaluation in LEO Environment,”
Advances in Space Research, Vol. 66, No. 2, 2020, pp. 307–320.
https://doi.org/10.1016/j.asr.2020.03.024
[13] Duzellier, S., Gordo, P., Melicio, R., Valério, D., Millinger, M., and
Amorim, A., “Space Debris Generation in GEO: Space Materials
Testing and Evaluation,” Acta Astronautica, Vol. 192, March 2022,
pp. 258–275.
https://doi.org/10.1016/j.actaastro.2021.12.036
[14] Cooper, R., and Hoffman, R., “Jumbo Space Environment Simulation
and Spacecraft Charging Chamber Characterization,” U.S. Air Force
Research Lab., Space Vehicles Directorate Kirtland AFB, AFRL-RVPSTP-2015-0012, Albuquerque, NM, 2015.
[15] Plis, E. A., Engelhart, D. P., Cooper, R., Johnston, W. R., Ferguson, D.,
and Hoffmann, R., “Review of Radiation-Induced Effects in Polyimide,” Applied Sciences, Vol. 9, No. 10, 2019, Paper 1999.
https://doi.org/10.3390/app9101999
[16] Engelhart, D. P., Plis, E., Humagain, S., Greenbaum, S., Ferguson, D.,
Cooper, R., and Hoffmann, R., “Chemical and Electrical Dynamics of
Polyimide Film Damaged by Electron Radiation,” IEEE Transactions
on Plasma Science, Vol. 45, No. 9, 2017, pp. 2573–2577.
https://doi.org/10.1109/TPS.2017.2729516
[17] American Society for Testing and Materials (ASTM) Book of Standards, Vol. 15.03 Developed by Subcommittee, E21.04, p. 5.
https://doi.org/10.1520/E2089-15R20
[18] Plis, E., Bengtson, M., Engelhart, D. P., Badura, G., Scott, T., Cowardin,
H., Reyes, J., Hoffmann, R., Sokolovskiy, A., Ferguson, D. C., Shah, J.,
and Horne, S., “Characterization of Novel Spacecraft Materials Under
High Energy Electron and Atomic Oxygen Exposure,” AIAA SciTech
2022 Forum, AIAA Paper 2022-0797, 2022.
https://doi.org/10.2514/6.2022-0797
[19] American Society for Testing and Materials (ASTM) Book of Standards, Vol. 10.01 Developed by Subcommittee, D09.12, p. 18.
https://doi.org/10.1520/D0257-14R21E01
[20] Frederickson, A. R., and Dennison, J., “Measurement of Conductivity
and Charge Storage in Insulators Related to Spacecraft Charging,” IEEE
Transactions on Nuclear Science, Vol. 50, No. 6, 2003, pp. 2284–2291.
https://doi.org/10.1109/TNS.2003.821397
[21] Bengtson, M., Maxwell, J., Hoffmann, R., Cooper, R., Schieffer, S.,
Ferguson, D., Johnston, W. R., Cowardin, H., Plis, E., and Engelhart, D.,
“Optical Characterization of Commonly Used Thermal Control Paints
in a Simulated GEO Environment,” The Advanced Maui Optical and
Space Surveillance Technologies Conference, Proceedings of the
Advanced Maui Optical and Space Surveillance Technologies
Conference, edited by S. Ryan, The Maui Economic Development
Board, Sept. 2018, p. 33.
T. K. Minton
Associate Editor
The Spectral Characterization of Novel Spacecraft Materials in the
Low Earth Orbit Environment
Elena A Plis*,** and Gregory Badura*
The protection of satellites from the harsh space environment is heavily
reliant on external spacecraft materials. These materials undergo continuous changes in their physical, chemical, and optical properties due to exposure to solar radiation and aggressive chemical species present in
Earth's upper atmosphere. By gaining a thorough understanding of how
these material properties evolve over the planned lifetime of a mission, the
reliability of spacecraft can be enhanced. Additionally, the establishment of
correlation factors between actual space exposure and accelerated space
weather experiments performed in ground facilities allows for the precise
prediction of on-orbit material performance based on laboratory-based testing. This paper presents an assessment of the radiation effects of the low
Earth orbit (LEO) environment, including exposure to high-energy electrons, atomic oxygen (AO), and vacuum ultraviolet (VUV), on the material
properties of a prospective spacecraft polymer, Mylar M021. The study is
part of the 16th Materials International Space Station Experiment (MISSE),
which our team conducted from 08/31/22 to 02/23/23. An overview of the
MISSE project is provided, including details on the materials used, results
of ground experiments, and preliminary findings from the orbital phase of
the study.
* Georgia Tech Research Institute, 925 Dalney St NW, Atlanta, GA 30318
**Assurance Technology Corporation, Carlisle, MA, 01741 USA
1
INTRODUCTION
The low Earth orbit (LEO) is a valuable asset of humankind for many different reasons.
LEO satellites have been utilized for voice and data communication since 1998 (Iridium Communications Inc.’s network) due to the low delay that occurs in round-trip data transmission.
Since then, the number of communication satellites in LEO has increased significantly, with
over 3,717 Starlink satellites launched since 2018 to provide high-speed internet service worldwide [1]. These satellite surveillance and communication capabilities are vital for USSF military
and intelligence operations. Next, LEO is a critical staging point for space exploration and the
development of technologies required for future missions to more distant destinations such as
the Moon, Mars, and beyond. Finally, humankind’s only permanent residence in space, the
International Space Station (ISS), is operating in LEO.
During a space mission, surface spacecraft materials are exposed to various environmental factors such as different types of ionizing radiation, including high-energy electrons and
protons, AO ions, micrometeoroids, vacuum, and large temperature fluctuations [2, 3, 4]. The
resulting change in the material properties can significantly impact materials’ performance and
durability [e.g., 5, 6] which could degrade the functionality of the spacecraft elements. Even
though all space environmental components can lead to the deterioration of spacecraft components, the threat posed by AO, the most abundant specie at LEO, is especially severe in
terms of structural, thermal, and optical damage, particularly to exterior spacecraft components
that are susceptible to oxidation [7].
A comprehensive understanding of how the properties of materials change over the
course of a long-term space mission is crucial for ensuring that the materials used can withstand the harsh conditions of space and function as intended throughout the mission. Moreover, as space exploration becomes simultaneously more commonplace and ambitious, new
and advanced materials are required to meet the increasingly demanding needs of space missions. Novel materials born on the expanding boundaries of organic polymer chemistry must
be lightweight, have improved long-term radiation shielding properties, and possess improved
mechanical properties such as enhanced elasticity and mechanical strength for use in internal
and external spacecraft systems.
One major barrier to utilizing novel materials on spacecraft is the immense cost of flying
to space. Testing novel materials intended for the spacecraft industry on Earth may be difficult,
as it is challenging to recreate the exact space environment in a laboratory setting. LEO is a
particularly harsh environment for organic polymers because AO is present along with all other
2
environmental components [8]. Efforts to reproduce the complex LEO environment in groundbased facilities have been reported by US and international research teams [e.g., 9, 10, 11].
However, significant differences exist between actual LEO and ground-based exposures (i.e.,
variations in AO quantum state and AO energies and fluxes, temperature cycling, and simultaneous UV and ionizing radiation). In addition, the effects of the space environment on spacecraft components vary based on specific mission details such as mission duration, orbital parameters, view angle of spacecraft surfaces to the sun, and orientation of spacecraft surfaces
with respect to the spacecraft velocity vector in LEO [12]. Such integrated testing with replication of actual space conditions is virtually impossible to realize in Earth-based testing facilities.
Therefore, ground-based experiments must be benchmarked against actual LEO exposure.
The MISSE Flight Facility (MISSE-FF) provides unique capabilities for testing neoteric materials under exposure to the integrated space environment. Establishing correlation factors between actual space exposure and accelerated space weather experiments is a key benefit of
using the MISSE-FF in conjunction with space-simulated facilities. This enables accurate prediction of on-orbit material performance based on laboratory-based testing, which can improve
the reliability and safety of space missions.
The MISSE project is a NASA initiative that was launched in 2001 with the aim of studying
the effects of the LEO space environment on a variety of materials. Since its inception, the MISSE
project has tested approximately 4,000 different material samples and specimens, evaluating their
durability and performance in the harsh and challenging conditions of space. From 2001 to 2013,
individual flight experiments were carried out using Passive Experiment Containers (PECs), which
were suitcase-like containers placed outside the ISS either in a Ram/Wake or a Zenith/Nadir orientation by an astronaut during a spacewalk [13]. Following the 1- 4 years exposure period, the
PECs were retrieved during a spacewalk and brought back to Earth on the shuttle for post-flight
analysis. In 2018, the MISSE-FF, a modular robotically serviceable external platform that can be
configured to carry out a wide range of passive and active experiments, was launched [14]. Since
then, several ~ 6 months long MISSE missions were successfully completed.
The project discussed here was a component of the MISSE-16 space mission, which was
launched aboard the SpaceX-25 spacecraft from Kennedy Space Center on July 15th, 2022, and
subsequently docked at the Harmony Forward (Node 2 Forward) on July 16th, 2022. MISSE-16
Science Carriers (MSCs) were installed on the MISSE-FF on 31 July - 2 August 2022 with support
from the Canada Space Agency ISS Robotic Systems Team.
3
This project aimed to gather high-fidelity time-resolved spectral measurements along with
environmental data collected at the same cadence in order to clearly delineate the effect of total
irradiation dose as well as any flux-dependent material response. During the expected mission
duration (approx. 6 months), we intend to investigate short- and long-term degradation effects of
LEO on selected materials, by tracking their spectral properties as a function of time-on-orbit.
Previous experiments have shown that radiation-induced color change occurs rapidly upon first
exposure to radiation and that the changes become more gradual with increasing fluence. By
correlating color change to a host of other material properties determined using ground-based
characterization, we will enable Earth-bound observers to remotely diagnose spacecraft health
based on color shifts detected via unresolved remote observations. Lastly, laboratory-based irradiation experiments are nearly always performed at an accelerated dose rate with imperfect radiation sources. Therefore, truth data gathered in the real space weather environment is invaluable
for developing accurate ground testing protocols.
This paper offers a high-level overview of the project, including details on the materials
used, results of ground experiments, and preliminary findings from the orbital phase of the study.
EXPERIMENTAL PROCEDURE
Materials
Three identical sets of samples were placed on the Ram, Zenith, and Wake faces of the
ISS. Table 1 lists the flight samples that included several materials from the Kapton® family, polyethylene terephthalate (PET), liquid crystal polymer (LCP), carbon and glass fiber-reinforced
polymers (CFRP and GFRP, respectively), polyhedral oligomeric silsesquioxanes (POSS), and
alumina slide. The area of each sample was 1 inch2. The samples were constructed by stacking
multiple layers to achieve a minimum total thickness of 10 mils (0.01 inch). In some cases, more
than one layer was required to reach the desired thickness. For space-simulated experiments
including exposure to AO, VUV, and high-energy electrons, single-layer samples were utilized.
4
TABLE 1. List of materials for flight samples.
8
9
10
11
12
13
14
15
Density
(g/cm3)
1.54
1.42
1.42
1.41
1.42
1.42
1.43
Single layer
thickness (mil)
2
2
1
2
4
2
3
CFRP
1.44
29
1
GFRP
1.90
66
1
LCP
1.75
1.20
1.39
1.40
2.08
3.95
3
5
9
0.6
0.8
3.8
4
2
1
17
13
3
Polyimide
(PI)
PET
POSS
ALUM
# of layers
5
5
10
5
3
5
4
MSC
side
Mount
Abbreviation
Swing
1
2
3
4
5
6
7
Material/Tradename
Kapton CR
Kapton CS
Kapton WS
Kapton XC
Kapton TF
DR9
Kapton HN
Economyplate™
Carbon Fiber
G-10/FR4 Glass
Epoxy
Zenite®
Melinex 454
Mylar M021
CORIN® XLS
Thermalbright® N
Alumina
Kapton® has been a trusted material for safeguarding satellites, spacecraft, and astronauts for over five decades, dating back to the Apollo 11 mission [15]. We used Kapton-HN® film
as a well-established reference material for our experiments. Furthermore, we incorporated various new Kapton formulations, including CR (corona-resistant, crucial for shielding sensitive equipment on spacecraft), CS (clear, smooth, and thermoformable, with potential to replace traditional
PI film in MLI blankets and enhance next-generation space solar cells), WS (opaque film with high
thermal stability, a viable alternative to traditional PI film in MLI blankets), XC (electrically conductive, valuable for mitigating space charge on spacecraft), and TF (thermoformable, a promising
material for small satellite components). Additionally, we flew the DR9 polymer, a novel highperformance organic polymer material that has not been previously tested for spacecraft applications.
CFRP and GFRP are widely used in the construction of present-day LEO satellites [16,
17]. Unfortunately, little is known about the evolution of the optical properties of these materials
in the harsh LEO environment [18, 19]. Laboratory impact tests have shown that CRFP shatters
into small needle-like shapes, which are a clear risk in the orbital debris environment. It is unclear
whether the optical properties of these fragments change following impact, complicating ground-
5
based observations of these objects. GFRP tends to be translucent in nature and laboratory studies have not determined how or if it degrades with time-on-orbit. Studying the optical properties
of these materials would provide a baseline of how they respond to the harsh LEO environment
and allow better monitoring for improved space domain awareness.
LCPs present a special category of materials, which straddles the boundary between an
ordinary solid and a liquid. LCPs have excellent dimensional stability (fatigue and creep resistance), as well as high dielectric strength over wide temperature range [20]. Moreover, LCPs
show a high resistance to UV radiation and very good electrical insulation properties [21]. These
unique properties in conjunction with light weight make LCP materials very attractive for space
applications. Flexible LCP antennas and LCP-based circuits molded to available spacecraft areas
could eliminate heavy metal boxes that currently house rigid circuit boards [22]. Before recommendation of LCP material to spacecraft designers, evaluation of its durability under the harsh
space conditions of LEO must be performed. We will test effect of LEO environment on one
variety of DuPont Zenite® LCP.
PET films are an essential part of multi-layer insulation (MLI) blankets that are used on
the exterior surfaces of spacecraft for the purpose of passive thermal control. MLI composite
materials work by limiting the amount of radiative heat transfer through multiple layers of thin
reflectors (shields) and spacer materials. The shields are generally PET or polyimide (PI) metalcoated films with high mechanical strength and low thermal conductivity. An ideal thermal blanket
completely reflects the incident radiation. Although these perfect reflectance characteristics have
not been obtainable in practice yet [23] and research is being conducted on alternatives to traditional MLI blankets, such as those based on silica aerogels [24, 25, 26], which offer improved
reflectivity and durability, PET-based shields are still an essential part of spacecraft thermal control system. Thus, new materials need to be developed for improved MLI reflectivity and durability.
We included two novel types of PET, Melinex® 454 and Mylar® MO21, into the experimental
materials selection.
One promising method to increase the AO-resistivity of PI films is incorporation of the
monomer POSS into the PI polymer matrix, either as a chemically bound co-polymer or an additive blended into the matrix [5, 6]. The unique benefit of this method over traditional methods,
such as coatings, is that when POSS-PI is initially exposed to AO, it forms a layer of SiOx over
the surface. This passivating layer prevents further material ablation, thus lowering the overall
erosion yield. Moreover, if the surface is damaged in anyway and cracks appear, the newly exposed surface will simply form another layer of SiOx. Hence, POSS-PI essentially a self-healing
polymer. We included two POSS/PI samples, Thermalbright™ (a colorless polyimide with POSS
6
additives) and CORIN (a transparent, sprayable PI with POSS additives). As PI/POSS hybrid
materials have already been identified as excellent AO resistant materials, a detailed study of
their optical properties will be of great utility to the remote sensing community for the characterization of orbital debris and operational RSOs.
Finally, to ensure the accurate comparison between images captured after different space
exposure times, we have chosen pure aluminum oxide (alumina, Al2O3) as our reflectance standard as it is resistant to AO attack and has a flat white spectral profile.
Space-simulated Irradiation Procedure
Material irradiation with high-energy electrons as well as the VUV exposure was performed in the Jumbo space irradiation chamber at the Spacecraft Charging and Instrument Calibration Laboratory (SCICL) at Kirtland AFRL [27]. Materials were bombarded with high-energy
(100 keV) electrons produced by a mono-energetic Kimball Physics EG8105-UD electron flood
gun. The maximum electron fluence the materials under investigation were exposed to is 8.5 x
1013 electrons/cm2, which corresponds to approximately 10 years at 600-800 km (LEO) orbit, delivered during 5.5 hours. A 24-h vacuum dehydration bakeout was performed on all samples at
60 C prior to radiation exposure to remove any absorbed water. Details of the electron irradiation
procedure are reported elsewhere [28].
The space-simulated VUV exposure was conducted using three Resonance KrLM-LQD12
lamps installed in the Jumbo chamber at SCICL. The lamp spectrum is composed of 1.0 parts
and 2.38 parts of photons with peak wavelengths at 116.5 nm and 123.6 nm, respectively [29].
The photon flux at each wavelength was estimated as 5.5 x 1020 photons/ (seconds · cm2 · A) at
116.5 nm and 1.06 x 1020 photons/ (seconds · cm2 · A) at 123.6 nm. Based on Lyman-Alpha flux
in-vacuum conversion [30], samples receive an average of 50.9 equivalent sun hours (ESH) per
single sample carousel rotation. The expected solar net flux for 6 months of the MISSE-16 mission
was calculated to be ~1094 ESH, assuming that ISS completes approximately sixteen 90-minute
revolutions per day. Samples studied in this work were exposed to ~ 6 months of LEO space
weather in ~ 23 hours.
Exposure testing of sample materials to an 8 km/s oxygen atom beam was conducted
according to ASTM-E2089-15A using the FAST source at the Physical Sciences Inc. The effective
peak atomic oxygen fluence during the exposure was calculated using the known witness sample
density (Kapton© H, 1.427g/cm3) and LEO erosion yield of the same material (3 x 10-24 cm3/Oatom). Effective peak atomic oxygen fluence during the run was 3.1 × 1020 O-atom/cm2 which
7
corresponds to 6 weeks of LEO exposure. Prior to exposure, all samples were stored in the vacuum chamber for 24 hours to remove water. They were then evacuated and weighed over time
to monitor water absorption and allow for the dry mass calculation. Details of the AO-irradiation
procedure may be found in [31].
Orbital Irradiation Experiments
The LEO exposure time of each MSCs is presented in Table 2.
TABLE 2. MSCs exposure times.
MSC
Exposure time
Wake
193 days  6.3 months
Zenith
194 days  6.3 months
Ram
183 days  6.0 months
Upon reviewing the radiation data for MISSE-16, the Aegis Aerospace operations team
observed that the data may have been affected by noise in the active radiation sensor. Consequently, the team has decided to wait for the passive dosimeter data to become available. These
data will be acquired after the MISSE science carriers are retrieved and processed by the dosimeter manufacturer. This will enable our team to obtain more accurate radiation measurements
and mitigate the effects of any potential noise in the active radiation sensor data.
Characterization protocol for flight duplicate materials
As part of the characterization process, several different measurements were conducted
to assess the properties of the test samples, with a focus on their optical characteristics. Optical
measurements were performed using techniques such as direct hemispherical reflectance (DHR),
reflectance (R), transmittance (T), Fourier-transform infrared spectroscopy (FTIR), and bidirectional reflectance distribution function (BRDF) measurements. In addition, surface morphology
characterization was undertaken using scanning electron microscopy (SEM) and atomic force
microscopy (AFM). To access the changes in radiation-induced electrical conductivity of the materials, surface potential decay (SPD) [32] and ASTM-D-257 techniques were employed. Finally,
visible and infrared (IR) images of all radiation-damaged materials were taken in vacuo using the
same Basler daA1600 camera and illumination scheme as was utilized in MSCs.
AFM measurements were performed using Bruker Dimension ICON instrument allowing measurement of surface roughness up to 5 𝜇m on areas as large as 200 𝜇m x 200 𝜇m. SEM
8
measurements were performed using Hitachi SU-8230 which features novel cold field emission
(CFE) gun for improved imaging and analytical performance.
The DHR measurements were performed in according to the optical data acquisition procedure reported elsewhere [33]. UV/Vis transmission measurements were performed with the
Cary 5000 UV-Vis-NIR spectrometer. The Hemispherical Conical Reflectance Factor (HCRF), the
laboratory approximation of true Bidirectional Reflectance Distribution Function (BRDF) measurements, of pristine samples with fixed viewing angles (0◦ , 30◦ , 60◦ , -30◦ , and -60◦ ) and variable
illumination angles (0◦ - 70◦ degrees range, exact values of illumination angles are specified separately for each measurement), was evaluated in the principal plane of illumination. The samples
were illuminated with a collimated 50 Watt tungsten halogen that approximated a blackbody
source over the range of 350-2500 nm. An Analytical Spectral Devices Spectrometer was used
to collect radiance measurements over this spectral range [34]. FTIR measurements were performed with Nicolet 6700 FT-IR spectrometers from Thermo Electron Corporation, with a wavenumber resolution of 2 nm. For reflectance measurements, an Analytical Spectral Device (ASD)
field spectrometer, ASD FieldSpec4, with a range from 300-2500 nm and a full-width half maximum of 3 nm at 700 nm, and 10 nm at 1400/2100 nm, respectively, was employed. In addition,
mass loss of AO-exposed materials was measured with 10 microgram accuracy using the microgram balance and then AO-erosion rate was estimated.
Geographically the members of the research team are separated by hundreds of miles. In
particular, the electron and UV space-simulated exposure was performed at the SCICL laboratory
whereas AO exposure was outsourced to Physical Sciences Inc [35] (Figure 1). Some of the
characterizations of the space weather effects on material properties of spacecraft materials were
in situ at the SCICL and, mainly, ex situ at the GTRI. To ensure the accuracy of the characterization experiments, it is important to guarantee that all collaborators receive space-weather-exposed samples that have undergone minimal recovery from their atmospheric exposure, and that
all characterization procedures are carried out with minimal air exposure. To comply with these
requirements, we have developed a procedure for handling spacecraft materials that are sensitive
to atmospheric exposure [36]. This procedure allows sufficient time for the delivery of air-sensitive
materials and for conducting all intended characterization experiments on radiation-damaged materials.
9
Figure 1. Transfer logistics of spacecraft materials exposed to the space-simulated environments. The blue lines correspond to the transfer of electron-irradiated and VUV-exposed samples, the red lines correspond to the transfer of AO-exposed materials.
Characterization protocol for flight materials
Visible and IR images of every sample were sequentially acquired using a Basler daA1600
camera on MSCs mounted on zenith, ram, and wake ISS faces. The experimental cadence includes daily images during the first week of the materials’ exposure, weekly during the next two
months, and monthly till the end of the mission. It should be noted that there was a sensor anomaly that prevented AEGIS Aerospace from fully opening the swing side of wake and ram MSCs.
Thus, data collection was not performed for the four samples, namely, #12 (Mylar® M021), #13
(CORIN XLS), #14 (Thermalbright® N), and #15 (Alumina), placed on the swing side of ram and
wake MSCs. Our team will have the chance to examine all of the materials after they return to
Earth, so for these particular samples at least two data points will be obtained, namely, at their
pristine state (taken on Earth prior to the mission) and at the most damaged state (to be taken
again on Earth after MISSE- FF comes back on April’23).
Optical images were downlinked to Earth within days of the performed measurements.
After initial analysis of those images, a machine learning (ML) approach based on a radial basis
function (RBF) network [37] was utilized to extract reflectance spectra from RGB/IR images. Once
images are downloaded from the on-orbit MISSE-FF cameras, the RGB data are selected from
non-oversaturated regions under both visible and IR lighting conditions. Next, these pixel counts
were white-balanced given the scene context, which is possible because the same settings have
been applied to the flight and ground cameras. Finally, the RGB MISSE-FF data were converted
10
into the L*a*b* color space and fed into the RBF network. The ML algorithm produced an estimated reflectance spectrum for each image. A summary of the approach utilizing ML and computer vision algorithms to retrieve reflectance data from RGB images is shown in figure 2.
Figure 2. Summary of approach utilizing ML and computer vision algorithms to retrieve reflectance data from RGB images.
RESULTS
Due to the abundance of samples in our study, it is not feasible to present the results for
all materials. Therefore, we will focus our discussion on one particular material, namely, Mylar®
M021. Table 3 summarizes the surface roughness of pristine, electron-irradiated, AO- and VUVexposed Mylar® M021 accessed by AFM. Average surface roughness (𝑅𝑎) values are averages
of several 5𝜇m x 5𝜇m scans taken at different parts of the sample. Representative 5𝜇m x 5𝜇m
AFM scans of pristine and exposed to different components of space-simulated environment Mylar® M021 material are shown in Figure 3. Since no significant effect of electron and VUV exposure
on the surface roughness of the material was observed, whereas AO exposure significantly affected the material’s morphology, SEM images of only AO-exposed Mylar® M021 samples were
taken, as shown in Figure 4.
TABLE 3. Surface roughness of Mylar® M021 exposed to different components of spacesimulated environment.
Pristine
Electron-irradiated
AO-exposed
VUV-exposed
7.8 nm
3.6 nm
187.2 nm
4.6 nm
11
Figure 3. Representative 5𝜇m x 5𝜇m AFM scans of pristine, electron-irradiated, AO-exposed,
and VUV-exposed Mylar® M021.
Mass loss, erosion rate, and relative erosion rate with respect to the Kapton® HN of the 9
mil Mylar® M021 sample after AO exposure was 7.89 mg, 7.29 x 10-21 g/cm2/O-atom, and 1.70,
respectively.
Figure 4. Representative 200 m x 200 m SEM images of pristine and AO-exposed Mylar
M021.
Figure 5 presents the DHR spectra of control (pristine), electron-irradiated, and VUVexposed Mylar M021 material in 400 - 2000 nm spectral range taken at different levels of spacesimulated exposure. Since AO irradiation was performed at a different facility, no DHR data were
taken for the AO-exposed material. Comparison of transmission characteristics (within 200 – 2000
nm wavelength range) of pristine and exposed to the different components of the space-simulated
weather Mylar M021 is presented in figure 6. The respective FTIR transmission curves are
shown in figure 7. Representative HCFR data which were taken at one selected pair of viewing
and illumination angles (I = 0 and v = 30) is presented in figure 8. The absolute reflectance
measurements were performed for all studied flight duplicate samples; however, results for the
AO- and VUV-exposure are still being processed. Figure 9 presents the results of such measurements for pristine and electron-irradiated Mylar M021 material with 9 mil and 10 mil thicknesses.
Finally, the electrical conductivity of Mylar M021 was initially planned to be evaluated by
two complementary methods, namely, SPD and ASTM D-257. Whereas the measurements and
subsequent analyses of the latter have been completed, and the results are presented in Table
12
4, the SPD measurements were only partially finished due to hardware issues affecting the material set. The findings obtained from SPD measurements will be presented in a future publication.
Figure 5. DHR measurements of (a) electron-irradiated and (b) VUV-exposed Mylar M021.
Figure 6. Transmission curves of Mylar M021 material exposed to different components of
the space-simulated environment.
13
Figure 7. FTIR Transmission curves of Mylar M021 material exposed to (a) high energy electrons, (b) AO, and (c) VUV components of the space-simulated environment.
14
Figure 8. HCFR measurements taken at 0
Figure 9. Absolute reflectance measurements
illumination angle (i) and 30 viewing angle
of Mylar M021 in its pristine state and after
(v) of Mylar M021 material exposed to dif-
electron irradiation. Adapted from [38].
ferent damaging types of irradiation.
TABLE 4. Bulk resistivity (·cm) of Mylar® M021 exposed to different components of
space-simulated environment.
Material
ASTM D-257
100V
500V
1000V
17
17
Pristine
3.6 x 1017
5.6 × 10
Electron-irradiated
1.4 x 1016
8.4 × 1016 1.3 × 1017
AO-exposed
8.3 x 1016
1.4 × 10
VUV-exposed
2.0 × 10
17
17
17
5.7 × 10
5.1 × 10
17
1.3 × 10
17
5.1 × 10
RGB and IR images of Mylar® M021 samples taken during the MISSE-16 orbital experiment. Due to Mylar® M021 (sample #12) being mounted on the swing side of the MCSs alongside
three other samples, its imagery on wake and ram MCSs was impacted by a sensor malfunction,
leading to the delivery of only the images captured by the zenith MSC during the mission, as
shown in Figure 10. The respective reflectance values extracted from the zenith RGB pictures
taken on orbit during the MISSE-16 mission using ML algorithm are presented in Figure 11.
15
(a)
(b)
Figure 10. (a) RGB and (b) IR images of Mylar® M021 delivered from the zenith face of ISS
characterized by the predominant VUV exposure. Data were taken during 08/31/22 – 02/23/23.
Figure 11. Predicted reflectance values of Mylar® M021 extracted from the RGB pictures taken
on orbit during the MISSE-16 mission using ML algorithm.
DISCUSSION
Among three different components of LEO space environment, namely, high energy electron irradiation, AO- and VUV-exposure, the AO irradiation has the most detrimental effect on the
surface morphology of the Mylar® M021. Exposure of polymer materials to AO at 8 km/s is sufficient to break the polymer bonds and induce oxidative decomposition, resulting in substantial
erosion of polymer surfaces which manifests itself as mass loss, thinning, and texture roughening.
AO exposure resulted in an increase of the surface roughness from 7.8 nm (pristine material) to
187.2 nm. Due to the nature of electron and VUV particles we did not expect to reveal significant
16
surface modifications of the electron-bombarded and VUV-exposed samples, which was confirmed by the visual inspection of irradiated materials as well as AFM data. SEM measurements
confirmed the erosion of the AO-exposed Mylar® M021. It is noteworthy that both electron irradiation and VUV exposure resulted in a decrease in surface roughness of the material in comparison to the pristine sample. This reduction in surface roughness may indicate an increase in radiation-induced crosslinking.
DHR measurements of the electron-irradiated and VUV-exposed samples revealed the
degradation of reflectance after electron exposure in 400 – 700 nm region whereas the prolonged
VUV exposure (23 hours) resulted in increased reflectance values in the entire studied wavelength range (figure 5). Absolute reflectance measurements of electron-irradiated Mylar® M021
(figure 9) demonstrated similar trend. Degradation of reflectance after electron bombardment may
be attributed to the radiation-induced condensation of aromatic rings into compact carbonaceous
clusters [39]. The observed reflectance behavior after VUV exposure may be attributed to the
changes in surface microscale smoothness caused by VUV-induced chemical and/or structural
changes in polymer surfaces. UV photons are energetic enough to dissociate molecular bonds,
resulting in photo-oxidation reactions that scissor or crosslink the polymer chains causing surface
morphology modifications. Additional analysis of this PET material as well as similar materials in
this class is necessary to accurately characterize the nature of the chemical changes caused by
VUV exposure to PET materials.
The transmission properties of Mylar® M021 material in the 350 nm – 2000 nm range are
expected to undergo significant degradation following AO exposure, resulting in a complete loss
of transparency and an opaque appearance. Oppositely, after high-energy electrons and VUV
irradiation, the transmittance of the studied material did not change considerably. Interestingly,
that electron and AO exposure did not change the band edge position of the material with respect
to its pristine state (~ 303 nm), whereas VUV exposure caused a shift of the band edge to the ~
348 nm, suggesting VUV-exposure induced shrinkage of the bandgap. This bandgap shrinkage
may be caused by the presence of excited states and free radicals in the bandgap of the VUVexposed PET, or, alternatively, by the formation of carbonyl groups and other oxygen-containing
species within the polymer matrix which alters the electronic structure of the material.
FTIR measurements showed no new peaks generated after the exposure to the LEO
space weather components, suggesting that no new chemical moieties are formed. FTIR spectra
of radiation-damaged materials demonstrated slightly increased overall transmission after high
energy bombardment and AO exposure. HCRF measurements of irradiated material demonstrated a reduction of signal intensity after VUV exposure.
17
After electron irradiation, the bulk resistivity of Mylar® M021 was found to have only a
minor reduction, which was not significantly affected by exposure to AO or VUV radiation. These
results suggest that the electrical properties of this PET variety are stable under the given levels
of exposure to the LEO environment components simulated in this study.
Visual inspection of images taken at MSC mounted on the zenith ISS face (Figure 10a),
characterized by dominant VUV exposure, revealed noticeable changes in the appearance of
Mylar® M021 in the visible spectral range. The material's color changed from a bright white at the
beginning of the experiment (Day 1) to a creamy yellow at the end of the experiment (Month 6),
indicating a shift in the material's optical properties. These changes may be attributed to photochemical reactions such as the accumulation of degradation products and discoloration due to
exposure to ultraviolet radiation.
No signs of mechanical degradation caused, for example, by micrometeoroid impacts
were observed during the experiment (figure 10b) indicating that the Mylar® M021 material remained structurally intact under the conditions of the study. However, other forms of mechanical
degradation, such as abrasion or fatigue, may have occurred and will be studied upon return of
the MISSE-16 experiment to Earth on April-May 2023.
Analysis of RGB pictures taken during the MISSE-16 mission using ML algorithm allowed
for the extraction of predicted reflectance values of Mylar® M021 under predominant VUV irradiation. The results presented in Figure 11 show that the reflectance of Mylar® M021 changed in a
complex manner during LEO exposure. Specifically, it increased by factor of two during the first
week of LEO exposure, then it steadily decreased during the first two months of exposure, followed by a gradual restoration to a steady value. Such results suggest a dependence of material
reflectance on multiple factors, including background electron irradiation. Also, it should be noted
that the accuracy of these predicted values may be affected by the quality of the images and the
assumptions made by the ML algorithm. Therefore, further research and validation are necessary
to fully understand the behavior of Mylar® M021 under VUV irradiation.
CONCLUSIONS
A comprehensive physical and chemical characterization of both novel and heritage
spacecraft materials during simulated space weather experiments is essential for establishing
correlation factors between true space exposure and accelerated space weather experiments
conducted on ground facilities. This will enable accurate prediction of on-orbit material performance based on laboratory-based testing. Such characterization can provide insights into the
underlying mechanisms of material degradation in space and inform the development of more
18
reliable materials for future space missions. In this study, we presented our approach to investigate the changes in optical properties of various spacecraft polymers under both true space and
space-simulated LEO conditions, using Mylar® M021 as a representative material. Furthermore,
we attempted to establish a connection between the changes in optical properties of the spacecraft polymer, Mylar® M021, and the alterations in surface morphology and charge transport properties induced by various components of the space-simulated LEO environment.
ACKNOWLEDGMENTS
This work was partially supported by the Georgia Tech Research Institute Independent Research
and Development (IRAD). The authors would like to express their gratitude to all colleagues who
contributed to this project, including Mr. Rayn Hoffmann, Drs. Jainisha Shah, Miles Bengtson,
Alexey Sokolovskiy, and Mrs. Sydney Collman from the SCICL/AFRL (Albuquerque, New Mexico,
USA) for preparation of electron-irradiated and VUV-exposed materials, as well as for the characterization of flight duplicate materials with DHR and ASTM D-257 techniques; Drs. Heather
Cowardin (NASA) and Jacqueline Reyes (University of Texas at El Paso) for performing absolute
reflectance measurements on flight duplicate materials and thorough analysis of their results; Dr.
Daniel Engelhart (University of New Mexico) for the valuable insights, expertise, and dedication
which were instrumental in the success of this research endeavor; and Mr. Timothy R. Scott from
DuPont de Nemours, Inc, for providing materials for this research.
CONFLICT OF INTEREST
On behalf of all authors, the corresponding author states that there is no conflict of interest.
DISCLAIMERS
The views expressed are those of the author and do not necessarily reflect the official policy or
position of the Department of the Air Force, the Department of Defense, or the U.S. government.
The appearance of external hyperlinks does not constitute endorsement by the United States
Department of Defense (DoD) of the linked websites, or the information, products, or services
contained therein. The DoD does not exercise any editorial, security, or other control over the
information you may find at these locations.
19
REFERENCES
1.
https://starwalk.space/en/news/spacex-starlink-satellites-night-sky-visibility-guide#is-it-ok-to-
launch-such-a-large-number-of-starlink-satellites
2. Y. Lu, Q. Shao, H. Yue, F. Yang. A review of the space environment effects on spacecraft in
different orbits. IEEE access 7, pp. 93473-88 (2019)
3. A. M. Abd El-Hameed. Radiation effects on composite materials used in space systems: a
review. NRIAG Journal of Astronomy and Geophysics. 11(1), pp. 313-24 (2022)
4. K. K. de Groh, B. A. Banks, S. K. Miller, J. A. Dever. Degradation of Spacecraft Materials.
InHandbook of Environmental Degradation of Materials, pp. 601-645. William Andrew Publishing
(2018)
5. G. Allegri, S. Corradi, M. Marchetti, V. K. Milinchuk. On the degradation of polymeric thin films
in LEO space environment. In Materials in a Space Environment 540, pp. 255-264 (2003)
6. F. Awaja, J. B. Moon, M. Gilbert, S. Zhang, C. G. Kim, P. J. Pigram. Surface molecular degradation of selected high performance polymer composites under low earth orbit environmental
conditions. Polymer degradation and stability 96(7), pp. 1301-9 (2011)
7. K. K. de Groh, B. A. Banks. Atomic oxygen erosion data from the MISSE 2-8 missions (No.
NASA/TM—2019-219982) (2019)
8. J. C. Yang, K. K. de Groh. Materials Issues in the Space Environment. MRS Bulletin 35, pp.
12-19 (2010)
9. J. K. McTernan, S. G. Bilén, and L. H. Krause. Facility for real-time test and verification of LEO
space plasma phenomena. In Applied Space Environments Conference, Huntsville, AL. (2017)
10. E. Grossman, I. Gouzman, G. Lempert, Y. Noter, Y. Lifshitz. Assessment of atomic-oxygen
flux in low-Earthorbit ground simulation facilities. Journal of Spacecraft and Rockets 41(3), pp.
356-9 (2004)
11. J. Kleiman, Z. Iskanderova, Y. Gudimenko, S. Horodetsky. Atomic oxygen beam sources: A
critical overview. In Materials in a Space Environment 540, pp. 313-324 (2003)
12. J. A. Dever, 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, pp. 465‑501. Norwich, NY: William Andrew Publishing (2005)
20
13. K. K. de Groh, B. A. Banks, J. A. Dever, K. J. Jaworske, S. K. Miller, E. A. Sechkar, S. R.
Panko. NASA Glenn Research Center’s Materials International Space Station Experiments
(MISSE 1-7). Proceedings of the International Sym. on “SM/MPAC&SEED Experiment,” JAXASP-08-015E, pp. 91-119 (2009).
14. K. K. de Groh, B. A. Banks. MISSE-Flight Facility Polymers and Composites Experiment 1-4
(PCE 1-4). NASA TM-20205008863 (2021)
15. T. Gold. Lunar-Surface Closeup Stereoscopic Photography. Apollo 14 - Preliminary Science
Report pp. 272 - 239 (1969)
16. S. Ramayanti, P. A. Budiantoro, A. Fauzi, E. Fitrianingsih, E. N. Nasser. Comparative study
of deployable satellite solar panel structure between Carbon Fiber Reinforced Polymer and Al7075 Honeycomb. In 2022 IEEE International Conference on Aerospace Electronics and Remote
Sensing Technology (ICARES), pp.1-6.
17. S. Y. Park, H. S. Choi, W. J. Choi, H. Kwon. Effect of vacuum thermal cyclic exposures on
unidirectional carbon fiber/epoxy composites for low earth orbit space applications. Composites
Part B: Engineering 43(2), pp. 726-38 (2021)
18. S. Sethi, B. C. Ray. Environmental effects on fibre reinforced polymeric composites: Evolving
reasons and remarks on interfacial strength and stability. Advances in colloid and interface science 217, pp. 43-67 (2015)
19. H. Cowardin, P. Seitzer, K. Abercromby, E. Barker, T. Schildknecht. Characterization of orbital
debris photometric properties derived from laboratory-based measurements. In 2010 AMOS
Technical Conference No. JSC-CN-21447.
20. T. J. White, D. J. Broer. Programmable and adaptive mechanics with liquid crystal polymer
networks and elastomers. Nature materials 14(11), pp. 1087-98 (2015)
21. E. A. Campo. Complete Part Design Handbook (Vol. 162). Munich, Germany:: Hanser (2006)
22. S. Gupta. Application of Liquid Crystals in Space Activities. Frontiers of Polymer Research, pp. 189-193 (1991)
23. J. Cepeda-Rizo, J. Gayle, J. Ravich. The multilayer insulation (MLI) blanket. In: Thermal and
Structural Electronic Packaging Analysis for Space and Extreme Environments, pp. 33–40. CRC
Press (2021)
21
24. M. A. Hasan, S. Rashmi, A. Esther, P. Y. Bhavanisankar, B. N. Sherikar, N. Sridhara, A. Dey.
Evaluations of silica aerogel-based flexible blanket as passive thermal control element for spacecraft applications. Journal of Materials Engineering and Performance 27(3), pp. 1265–1273
(2018)
25. Petroffe, G., Beouch, L., Cantin, S., Chevrot, C., Aubert, P.-H., Dudon, J.-P., Vidal, F.: Thermal regulation of satellites using adaptive polymeric materials. Solar Energy Materials and Solar
Cells 200, 110035 (2019)
26. H. Nikpourian, A. R. Bahramian. Thermo-physical properties of multilayer super insulation:
The role of aerogel blanket. Thermal Science and Engineering Progress 20, 100751 (2020)
27. R. Cooper, R. Hoffman. Jumbo space environment simulation and spacecraft charging chamber characterization. Technical report Air Force Research Laboratory, Space Vehicles Directorate
Kirtland AFB (2015)
28. D. P. Engelhart, E. Plis, S. Humagain, S. Greenbaum, D. Ferguson, R. Cooper, R. Hoffmann.
Chemical and electrical dynamics of polyimide film damaged by electron radiation. IEEE Transactions on Plasma Science 45 (9), pp. 2573–2577 (2017)
29. Resonance KrLM-LQD12 lamp brochure. Available at: https://resonance.on.ca/index_htm_files/LampComparisonBrochure.pdf
30. J. L. Lean, A. Skumanich. Variability of the Lyman alpha flux with solar activity. Journal of
Geophysical Research: Space physics 88(A7): 5751-5759 (1983)
31. E. Plis, M. Bengtson, D. P. Engelhart, G. Badura, T. Scott, H. Cowardin, J. Reyes, R. Hoffmann, A. Sokolovskiy, D. Ferguson, J. Shah, S. Horne. Characterization of novel spacecraft
materials under high energy electron and atomic oxygen exposure. In AIAA SCITECH 2022 Forum, p. 0797 (2022)
32. A. R. Frederickson, J. R. Dennison. Measurement of conductivity and charge storage in insulators related to spacecraft charging. IEEE transactions on nuclear science 50(6), 2284–2291
(2003)
33. M. Bengtson, J. Maxwell, R. Hoffmann, R. Cooper, S. Schieffer, D. Ferguson, W. R. Johnston,
H. Cowardin, E. Plis, D. Engelhart. Optical characterization of commonly used thermal control
22
paints in a simulated GEO environment. The Advanced Maui Optical and Space Surveillance
Technologies Conference p. 33 (2018)
34. G. Badura, E. Plis, C. R. Valenta. Extending Laboratory BRDF Measurements towards Radiometric Modeling of Resident Space Object Spectral Signature Mixing. n Advanced Maui Optical
and Space Surveillance Technologies Conference (AMOS) (2021)
35. R. H. Krech, G. E. Caledonia. Novel oxygen atom source for material degradation studies.
Final report, 1 March 1986-1 September 1988. Physical Sciences, Inc., Andover, MA (United
States); 1988 Oct 1
36. S. Collman. E. A. Plis, J. R. Shah, M. T. Bengtson, R. C. Hoffmann, D. C. Ferguson. Operational procedure for handling of spacecraft materials sensitive to atmospheric exposure. Journal
of Spacecraft and Rockets 60(2), pp. 685-8 (2023)
37. R. M. Nguyen, D. K. Prasad, M. S. Brown.Training-based spectral reconstruction from a single
RGB image. In European Conference on Computer Vision 2014 Sep 6 (pp. 186-201). Springer,
Cham
38. H. Cowardin, J. Reyes, E. A. Plis, R. C. Hoffmann, G. P. Badura, J. R. Shah, S. E. J. Collman,
M. T. Bengtson, D. P. Engelhart. Spectral Characterization of Modern Spacecraft Materials. Published as a proceeding of Advanced Maui Optical and Space Surveillance Technologies Conference (AMOS) (2022)
39. T. Steckenreiter, E. Balanzat, H. Fuess, C. Trautmann. Chemical modifications of PET induced by swift heavy ions. Nuclear Instruments and Methods in Physics Research Section B:
Beam Interactions with Materials and Atoms 131(1-4), pp.159-66 (1997)
23
CEAS Space Journal
https://doi.org/10.1007/s12567-023-00495-y
ORIGINAL PAPER
Characterization of polyethylene terephthalate (PET) materials
under high‑energy electron exposure
Jainisha R. Shah1 · Miles T. Bengtson2 · Sydney Collman1 · Ryan C. Hoffmann3 · Dale C. Ferguson3 ·
Daniel P. Engelhart4 · Elena A. Plis1,5
Received: 17 October 2022 / Revised: 31 March 2023 / Accepted: 3 April 2023
© The Author(s), under exclusive licence to Council of European Aerospace Societies 2023
Abstract
External spacecraft materials play an important role in satellite protection from the harsh space environment.Research has
shown that the physical, chemical, and optical properties of matter change continuously as a result of exposure to solar
radiation and aggressive chemical species produced in Earth’s upper atmosphere. Thorough knowledge of the material
properties evolution throughout a planned mission lifetime helps to improve the reliability of spacecraft. Moreover, the
establishment of correlation factors between true space exposure and accelerated space weather experiments at ground
facilities enables accurate prediction of on-orbit material performance based on laboratory-based testing. The presented
work evaluates the radiation effects of different doses of high-energy electron exposure on surface morphology, optical, and
charge transport properties of two materials from the PET family, ­Melinex®454 and ­Mylar®M021.
Keywords High-energy electron exposure · Geosynchronous Earth orbit (GEO) · Low Earth orbit (LEO) · Reflectance
spectroscopy · Surface morphology · Charge transport
1 Introduction
Polyethylene terephthalate (PET) materials are proposed
for many applications in a spacecraft industry, including construction of inflatable structures for satellite deorbiting, [1] deployable membrane antenna structures [2],
fabrication of flexible and stretchable electronics for harsh
Jainisha R. Shah and Elena A. Plis have contributed equally to this
work.
radiation environment [3, 4], and foldable organic solar cell
arrays manufacturing [5, 6]. Most importantly, PET films
are utilized in multilayer insulation (MLI) blankets that are
employed on the exterior surfaces of spacecraft for passive
thermal control purposes. Spacecraft operate in a harsh and
demanding environment, which requires materials that can
withstand extreme conditions such as temperature fluctuations, high levels of radiation, and vacuum. The ability to
maintain an acceptable temperature range during all phases
1
Assurance Technology Corporation, 84 South Street,
Carlisle, MA 01741, USA
2
Sydney Collman
collman@assurtech.com
Postdoctoral Fellow, National Research Council Research
Associateship Program, Kirtland AFB, Albuquerque,
NM 87117, USA
3
Ryan C. Hoffmann
AFRL.RVBorgmailbox@us.af.mil
Air Force Research Laboratory, Space Vehicles Directorate,
Kritland AFB, Albuquerque, NM 87117, USA
4
Department of Chemistry and Chemical Biology, University
of New Mexico, 1128 University Blvd NE, Albuquerque,
NM 87102, USA
5
Georgia Tech Research Institute (GTRI), 925 Dalney St NW,
Atlanta, Georgia 30318, USA
* Jainisha R. Shah
shah@assurtech.com
Miles T. Bengtson
AFRL.RVBorgmailbox@us.af.mil
Dale C. Ferguson
AFRL.RVBorgmailbox@us.af.mil
Daniel P. Engelhart
dpengelhart@unm.edu
Elena A. Plis
Elena.Plis@gtri.gatech.edu
13
Vol.:(0123456789)
J. R. Shah et al.
of operation is an important task to ensure the success of
the mission [7, 8].
MLI composite materials work by limiting the amount
of radiative heat transfer through multiple layers of thin
reflectors (shields) and spacer materials. The shields are
generally PET or polyimide (PI) metal-coated films with
high mechanical strength and low thermal conductivity.
An ideal thermal blanket completely reflects the incident
radiation. Although these perfect reflectance characteristics
have not been obtainable in practice yet [9] and research is
being conducted on alternatives to traditional MLI blankets,
such as those based on silica aerogels [10–13], which offer
improved reflectivity and durability, PET-based shields are
still an essential part of spacecraft thermal control system.
By nature of their usage, MLI blankets are exposed to
harsh space weather environments comprising high vacuum,
solar ultraviolet (UV) radiation, thermal cycling, and impacts
from micrometeoroids and artificial orbital debris. At low
Earth orbit (LEO), single-oxygen atoms (atomic oxygen,
AO) are the prevailing source of material degradation
whereas at geosynchronous Earth orbit (GEO), highly
energetic electrons are the dominant species interacting with
the spacecraft surface [14, 15].
Space-weather events can cause damage to the MLI
layers on spacecraft. This damage can weaken the adhesive
that holds the layers of MLI together, leading to their
delamination [16, 17] thus contributing to the population
of high-area-to mass (HAMR) space debris objects in
orbit [18]. By comparing the reflectance spectra of the
HAMR objects with the laboratory spectra of known
spacecraft materials, researchers were able to identify the
specific reflective material present in the MLI layers. This
finding supports the hypothesis that the HAMR debris was
formed as a result of delamination of the MLI layers [19–22].
The effects of dominant GEO environment species, high
energy electrons, on mechanical and charge accumulation
properties of PI [23] and PET [24–28] films causing the
degradation and fragmentation of MLI have been studied.
However, the number of studies devoted to the optical
changes of PET under the influence of space weather are
limited [29] even though understanding these optical changes
is invaluable for identifying and tracking debris clouds and
determining the origin of space debris [30–32]. This task
is accomplished by comparing the reflectance spectra of
observed orbital bodies (such as HAMR objects) with a
library of known (laboratory) reflectance spectra of different
materials. However, if PET layers are changing under the
influence of space weather, this must be taken into account
when analyzing and interpreting reflectance spectra. Any
changes in the optical properties of PET could potentially
lead to incorrect identification of debris materials. Moreover,
the radiation-induced alteration of optical properties may
serve as a proxy measurement for other, less tractable
13
measurements such as electrical conductivity, embrittlement,
surface morphology, and chemical reactivity.
In the presented work we studied changes of surface
morphology, optical, and charge transport properties of
two materials from the PET family, M
­ elinex ®454 and
®
­M ylar M021, under high-energy (100 keV) electron
irradiation. Evaluation of space weather effects on the
change in material properties of MLI components is helpful
for understanding of on-orbit characteristics of external
spacecraft materials. Various techniques were employed to
access the radiation-induced materials dynamics, such as
atomic force microscopy (AFM), directional hemisperical
reflectance (DHR), and surface potential decay (SPD)
measurements. Additionally, we present the astronomical
color index of chosen materials as a function of electron
fluence.
The astronomical color index of a given material is
a convenient metric that is experimentally tractable to
remote observers in situations where measurement of
a full reflection spectrum is not feasible [33]. A body
of literature is available on the utilization of the filter
photometry approach for optical observation data analysis
and interpretation. For example, Payne et al [34] emphasized
the importance of color indices to distinguish the satellites
with different configurations. Further, Cowardin et al [35]
used the same approach to determine the man-made space
body fragment brightness distribution. Beamer et al [36]
used the color-color technique for analyzing the wavelength
band intensity of reflected light from objects in space. Pearce
et al [37] used the clustering of the color indices of Russian
SL-12 rocket bodies for the identification of one unique
object (SL-12 RB 2012-012D). Finally, attempts to classify
spacecraft materials into families according to color index
derived from their reflectivity spectral curves were reported
by Reyes et al [31]. In our recent work [30] we compared all
possible color indexes combinations of g’ (406–544 nm),
r’ (558–682 nm), i’ (705–835 nm), and the z’ (838–1094
nm) passbands of Sloan Digital Sky Survey (SDSS) [38]
astronomical filter set to illustrate the importance of the
proper choice of filter combinations since the use of some
filter combinations result in more dramatic radiation-induced
changes to the color index of the material. Knowledge of
the evolution of a material’ color index may be harnessed
to provide information about a material’ chemical state and
physical properties using remote observations [35–37, 39].
2 Experimental Details
2.1 Materials
This study was conducted using two PET materials, 5 mil
(127 𝜇m) thick ­Melinex®454 film and 9 mil (229 𝜇m) thick
Characterization of polyethylene terephthalate (PET) materials under high‑energy electron…
­Mylar®M021 film from DuPont de Nemours, Inc. Visually,
the pristine ­Melinex®454 sample appears as a high clarity
film with both sides pre-treated to promote adhesion to most
Industrial coatings. Pristine ­Mylar®M021 is a low shrinkage,
low moisture absorption, and high thermal durability film
with a white opaque appearance. The studied materials had
no coatings applied.
2.2 Irradiation Procedure
Samples were irradiated with high energy (100keV)
mono-energetic electron radiation from a Kimball Physics
EG8105-UD flood electron gun in the Spacecraft Charging
and Instrument Calibration Laboratory (SCICL) at Kirtland
Air Force Base in New Mexico, USA [40]. The energy of
the electron beam was selected based on the continuous
slowing down approximation (CSDA) ranges of highenergy electrons [41]. The details regarding the anticipated
estimate of penetration depth of 100 keV electrons into
each material is not investigated and surpass the scope of
this manuscript. It is important to recognize that the space
environment is characterized by a broad distribution of
electron energies, which can cause energy deposition and
charging of spacecraft surfaces.
Samples were mounted over a carousel that rotated
through the hot spot of the electron beam to ensure
uniform irradiation. Sample size was 2.5 cm2 and reflective
metal surface of the same area was utilized as a backing
substrate.Copper tape and aluminum foil were employed
as backing materials for both irradiation experiments,
respectively. Prior to electron bombardment, a dehydration
bake-out of the loaded carousel was performed for 12 h at
600 C using a vacuum oven. More details of the electron
irradiation procedure have been reported elsewhere [42].
Irradiation was performed with two different electron beam
fluences, 8.5 x 1013 electrons/cm2 and 9.2 x 1014 electrons/
cm2 , corresponding to 6 h and 24 h of irradiation time,
respectively, which could be representative of a range of
different space environments. During materials irradiation,
background pressure was 3 × 10−7 Torr. The temperature of
the sample holder was not measured.
2.3 Characterization Methods
2.3.1 Optical properties
The directional hemispherical reflectance (DHR) of PET
samples was measured in situ before and during the electron
irradiation process in accordance with the optical data
acquisition procedure reported elsewhere [43]. In particular,
the data collection procedure began by measuring white and
black standards (Spectralon and Acktar Black, respectively)
using a Spectralon integrating sphere mounted on a robotic
arm. The Spectralon sphere was then moved to measure each
of the samples mounted on the rotating platform. During
the 10-minute measurement process, the electron beam
was extinguished to avoid doing damage to the Spectralon
standard.
In addition, the optical properties of pristine and
radiation-damaged samples were assessed in the UV/
Vis (200–800 nm) spectral region using a Cary 2000
spectrophotometer with spectral resolution of 2 nm. Finally,
color ratio plots were generated using astronomical Sloan
Digital Sky Survey (SDSS) [38] filters to show changes in
spectral brightness as a function of electron fluence [31].
Color ratio plots were generated from the measured
reflectance curves calculating the ratio of brightness between
two filter passbands (color index) using the equation 1
( )
I
A − B = −2.5 A
(1)
IB
where A and B represent the two filter passbands of
interest, and Ix is the brightness of the band x obtained by
integrating the reflectance curve over the wavelength range
of a given band. Differences between the brightness in the
g’ band (408–545 nm) and the z’ band (865–960 nm) were
determined.
2.3.2 Surface characterization
Surface morphology and roughness of studied materials
were examined using Bruker Dimension ICON atomic
force microscopy (AFM) allowing measurement of surface
roughness up to 5 𝜇m on areas as large as 200 𝜇m x 200 𝜇m.
2.3.3 Charge transport properties
The volume resistivity of pristine and irradiated PET samples
was evaluated using the surface potential discharge (SPD)
measurements [44] performed in a vacuum environment
with a low-energy (5 keV) Kimball Physics EGPS-2017B
electron gun and a TREK probe model 370 high-speed
electrostatic voltmeter. The back surface of the sample was
attached to the grounded backplane via copper tape with
conductive adhesive. To conduct the SPD experiment, the
front surface of the irradiated material was bombarded by
a beam of electrons during a short period of time, 1-2 s,
immediately after which the non-contact voltmeter was
positioned 1-2 mm from the surface and began to record
the surface potential. After the front of the charge body had
reached the grounded backplane, the dissipation of charge
was primarily determined by the loss of electrons from the
material. SPD measurements were performed in darkness to
eliminate the possibility of optically excited states obscuring
our analysis. Using the SPD method, the dark resistivity of
13
J. R. Shah et al.
the material may be derived from a plot of surface potential
versus time, or decay curve, using equation 2:
𝜌=
𝜏
𝜖0 𝜖r
(2)
where 𝜏 is charge decay time in seconds determined from
the linear fit of the post-transit region of the decay curve.
It represents the time it takes electrons deposited from the
beam to traverse the material and be lost to the grounded
backing plate. The 𝜖0 and 𝜖r are the permittivity of free space
and relative permittivity of the material, respectively. The
conductivity of the material is then calculated as inversely
proportional to the resistivity of the material. Whereas
the constant voltage method conforming to ASTM D-257
standard is commonly utilized by the material manufacturers
to attest the charge transport properties, the SPD method is
more applicable to test materials with irradiation-induced
heterogeneity under space-simulated conditions [45].
2.4 Sample handling
Unlike radiation-induced material degradation measured by
the in situ DHR measurements which were performed under
high vacuum conditions (< 10−6 Torr), the post-irradiation
measurements such as AFM or UV/Vis optical tests were
performed at different facilities, necessitating considerable
air exposure to the materials between their irradiation and
characterization. Since air has been shown to obfuscate
results for electron irradiated organic polymers (e.g., [46]),
handling and characterization protocols of the irradiated
air-sensitive materials must be scrutinized and carefully
controlled. To reduce the healing rate of PET materials
irradiated with high energy electrons upon exposure to
atmosphere, the two-stage sample packaging procedure
Fig. 1 Representative 5𝜇m x
5𝜇m AFM scans of (a) pristine
and electron-irradiated with (b)
8.5 x 1013 electrons/cm 2 and (c)
9.2 x 1014 electrons/cm 2 PET
materials
13
comprising vacuum sealing performed with a ­Henkelman®
industrial vacuum sealer was employed [47].
3 Results
Figure 1 shows representative AFM scans for each studied
sample under different electron irradiation dose. Table 1
summarizes the measured surface roughness values of pristine and electron-irradiated samples. Average surface roughness ( Ra ) values were the average of several 5 𝜇 m x 5 𝜇 m
scans taken at different parts of the respective samples.
Surface roughness of both PET materials decreased
after LEO irradiation with 8.5 x 1013 electrons/cm2 fluence
which is attributed to the coalescence of some of the original
defects. Higher irradiation fluence, 9.2 x 1014 electrons/cm2,
caused the formation of multiple point defects, probably
arcing sites, which contribute to the increased average
roughness of the materials. The surface between the arcing
sites is smoother compared to the pristine material and this
is easier to observe for the ­Mylar®M021.
Absolute hemispherical reflectance of ­Mylar®M021 and
­Melinex®454 samples measured at zero electron fluence of
impinging 100 keV electrons (pristine material), and two
Table 1 Surface roughness of pristine and electron-irradiated polymer samples
Fluence (e/cm 2)
Mylar®M021, Ra (nm)
Melinex®454,
Ra (nm)
Pristine
7.8
3.6
8.2
5.0
4.2
7.6
8.5 x 1013
9.2 x 1014
Characterization of polyethylene terephthalate (PET) materials under high‑energy electron…
Fig. 2 Absolute hemispherical reflectance curves for
(a) ­Mylar®M021 and (b)
­Melinex®454 samples irradiated
with different fluences of 100
keV electrons measured during
irradiation
Table 2 Volume resistivity of M
­ ylar®M021 and M
­ elinex®454 samples
𝜌 (Ω ⋅ cm)
­Mylar®M021
Pristine
4.3 x 1018
8.5 x 1013
4.8 x 1021
9.2 x 10
Pristine
3.8 x 1019
9.6 x 1018
13
8.5 x 10
7.5 x 1018
9.2 x 1014
1.6 x 1019
14
®
Melinex 454
reflection spectrum is not feasible [33]. Among twelve possible combinations of the g’, r’,i’, and z’ filters from the SDSS
astronomical filter set, the g’-z’ color index showed the
greatest variation as a function of electron fluence and this
filter combination was applied to the investigated films [30].
Fig. 3 UV–Vis transmittance spectra of ­
Mylar®M021 and
­Melinex®454 samples irradiated with different fluences of 100 keV
electrons
4 Discussion
different electron fluences is shown in Fig. 2. The aluminum
backing of the materials served as an optical mirror during
the DHR measurements, making the “reflection spectra”
superpositions of reflected and transmitted light.
The UV–Vis transmittance spectra of M
­ ylar ®M021
®
and ­Melinex 454 samples that were exposed to different
fluences of 100 keV electrons are shown in Fig. 3. Transmittance of both irradiated samples is decreased with an
increase of irradiation dose, more for ­Melinex®454 than for
the ­Mylar®M021.
The volume resistivity values of M
­ ylar ®M021 and
®
­Melinex 454 samples irradiated with different fluences of
100 keV electrons measured by SPD method are summarized
in Table 2.
The astronomical color index of a given material is
a convenient metric that is experimentally tractable to
remote observers in situations where measurement of a full
Whereas the shape of the absolute hemispherical reflectance
spectrum of ­Mylar®M021 film is more complicated than that
of the ­Melinex®454, several common features are identified,
such as the absorption band at 830 nm which is attributed to
the partial light absorption in the aluminum of the backing
plate. Intensity of the 1650 nm feature is not significantly
affected by electron irradiation, indicating that whatever
chemical moiety is responsible for it is not influenced by
the impinging high energy electrons.
A notable decrease of absolute reflectance of both
materials in a short wavelength range, 400–675 nm for
­M ylar ®M021 and 400–830 nm for ­M elinex ®454, was
observed with increased exposure duration to the high
energy electrons. Further increase in electron irradiation dose caused the 20% decrease in reflectance values
in 830–1800 nm range of M
­ elinex®454 film. Oppositely,
®
the ­Mylar M021 film showed a nearly constant value of
13
J. R. Shah et al.
absolute hemispherical reflectance in 1000–1800 nm region,
with a maximum of 10% reflectance increase in 830–1000
nm range. The latter may be due to the radiation-promoted
smoothness due to the radiation-induced crosslinking of
the ­Mylar®M021 compared to the pristine material, since
the reduced roughness causes a decrease of the spreading
of reflected light and, consequently, promotes more specular light. Although there is a lack of available data on the
relationship between the reflected light and the radiationchanged topography of PET materials, researchers have
investigated the impact of surface roughness on reflected
light intensity in other materials such as pigmented plastics
and paper [48, 49].
The decreased transmittance of the electron-irradiated
PETs suggests that exposure to high-energy electrons leads
to a decrease in the optical transparency of these films. This
decrease in transmittance can be attributed to the formation
of defects and color centers in the films due to the ionization
and excitation of the polymer chains by the high-energy
electrons.
The radiation stability of PET has been observed by other
researchers [50] and was attributed to the modification of
aromatic rings from di-substituted to mono-substituted
benzene groups [22]. The condensation of aromatic rings
into compact carbonaceous clusters may also result in
increased absorption of PET materials in a short wavelength
range. Indeed, the increased absorption of the irradiated PET
films, beyond 500 nm for M
­ ylar®M021 and in 300–600 nm
®
range for ­Melinex 454, is shown in Fig. 2.
Interestingly, the volume resistivity of the electronirradiated PET samples either increased ­(Mylar®M021)
or did not change significantly ­(Melinex ® 454). The
similar phenomenon has been observed by Chaudhary
et al [51] and Oproiu et al [52] (decreased conductivity
of PET after electron irradiation) and is attributed to the
increased crosslinking of the PET chains due to the electron
irradiation, which may obstruct the charge carrier hopping
from one chain to another chain resulting in decrease of
electrical conductivity. This result is surprising, considering
that one successful method for increasing the conductivity of
other organic polymers based on a polyimide backbone (such
as ­Kapton®) is to introduce conductive carbon clusters into
the polymer. Studies of the role of crosslinking in surface
roughening of electron-irradiated PET films are limited;
however, the smoother surface of electron-irradiated PET
samples may also be a manifestation of radiation-induced
increased cross-linking and demonstrates an example of
another aromatic thermoplastic, polystyrene [53]. One
very promising technique to study the fundamental causes
of these changes is in-situ vibrational spectroscopy, a
technique that is currently in development in the SCICL lab
in collaboration with the University of New Mexico.
13
The g’-z’ color index plot was utilized for the characterization of spectral brightness with increased electron fluence
of both studied PET materials. As shown in Fig. 4, both
materials demonstrated a monotonic increase of g’-z’ versus
electron fluence characteristics, resulted in a change of the
g’-z’ index by factor of 6.3 from the initial (pristine) to the
irradiated with maximum electron fluence values for both
samples, in particular, from of 0.08 to 0.50 for M
­ ylar®M021
and from 0.15 to 0.95 for ­Melinex®454. By understanding
these optical changes i.e. the spectral regions over which the
materials are stable and those where it varies with electron
exposure, remote observers can get information about the
material’s state. This knowledge about the history of the
material would be invaluable for identifying and tracking
debris clouds and determining the origin of the space debris
5 Conclusion
By understanding the behavior of satellite components
throughout their mission lifetime, Earth-based observers
may glean more detailed information from unresolved
imagery to help prevent space-based catastrophes and to
better understand the on-orbit life cycle of commonly used
spacecraft materials. This knowledge of on-orbit material
degradation will inform spacecraft designers and enable
the construction of more robust spacecraft designs as well
as improve the abilities of spacecraft operators to conduct
accurate and timely anomaly resolution.
In this paper, we investigate a dependence between
changes in optical behavior and other material properties
of two PET polymers irradiated with high-energy electrons.
Fig. 4 g’-z’ color index of pristine and irradiated with different fluences of 100 keV electrons M
­ ylar®M021 and ­Melinex®454 samples
Characterization of polyethylene terephthalate (PET) materials under high‑energy electron…
The alteration of volume conductivity observed in both
types of PET materials following electron irradiation,
which could be indicative of diverse space environments,
implies that differential spacecraft charging models, such
as NASA/Air Force Spacecraft Charging Analyzer Program
(NASCAP), should be revised to incorporate time- and
environment-dependent material properties to accurately
account for material aging. Further, as these polymers are
among commonly used spacecraft surface materials, the
accompanying optical changes may be of a magnitude that
is discernible via resolved or unresolved remote imaging
Acknowledgements Authors would like to thank DuPont de Nemours,
Inc, for providing materials for this research.
Author Contributions JRS and EAP conceptualized the idea of the
study. JRS, SC, and MTB performed all the experiments reported in this
manuscript excluding the surface morphology and FTIR measurements
which were performed by EAP. Optical, charge transport, and surface
morphology data were analysed by JRS, DPE, and EAP. Manuscript
was written by JRS and EAP. RCH, DPE, and DCF reviewed, provided
feedback, and approved the manuscript.
Funding This work was partially supported by Air Force Office of
Scientific Research, Remote Sensing and Imaging Physics Portfolio
(Dr. Michael Yakes) Grant 20RVCOR024 and Georgia Tech Research
Institute Independent Research and Development (IRAD) program.
Data availability Not applicable
Code Availability Not applicable
Declarations
Disclaimer The views expressed are those of the author and do not
necessarily reflect the official policy or position of the Department of
the Air Force, the Department of Defense, or the U.S. government.
Conflict of interest Authors declare no conflict of interest/competing
interests.
Ethical approval Not applicable
Consent to participate All authors consented to the participation in
research described in this manuscript.
Consent for publication All authors consented to the publication of
this manuscript.
References
1. Koryanov, V., Toporkov, A., Kazakovtsev, V., Nedogarok, A.: The
technology of using inflatable structures made of special materials
to adapt the de-orbiting system for removing spacecraft. In: IOP
Conference Series: Materials Science and Engineering, vol. 812,
p. 012006 (2020). IOP Publishing
2. Liu, Z.-Q., Qiu, H., Li, X., Yang, S.-L.: Review of large spacecraft
deployable membrane antenna structures. Chin J Mech Eng 30(6),
1447–1459 (2017)
3. Almuslem, A.S., Shaikh, S.F., Hussain, M.M.: Flexible and
stretchable electronics for harsh-environmental applications.
Adv Mater Tech 4(9), 1900145 (2019)
4. Cordill, M.J., Taylor, A.A., Berger, J., Schmidegg, K., Dehm,
G.: Robust mechanical performance of chromium-coated polyethylene terephthalate over a broad range of conditions. Philoso
Mag 92(25–27), 3346–3362 (2012)
5. Zhao, X., Hao, H., Dong, J., Cheng, F., Hao, J., Xing, J., Liu,
H.: Fast extraction of electron across the interface of nanowire
ch3nh3pbi3/zno on flexible pet substrate. Materials Letters 197,
139–142 (2017)
6. Liu, C., Xiao, C., Xie, C., Li, W.: Flexible organic solar cells:
Materials, large-area fabrication techniques and potential applications. Nano Energy 89, 106399 (2021)
7. Meseguer, J., Perez-Grande, I.: Spacecraft thermal control
woodhead publishing limited (2012)
8. Gilmore, D.G., Gilmore, D.G., Donabedian, M.: Spacecraft
Thermal Control Handbook vol. 1. Aerospace Press El Segundo,
CA, ??? (2002)
9. Cepeda-Rizo, J., Gayle, J., Ravich, J.: The multilayer insulation
(MLI) blanket. In: Thermal and Structural Electronic Packaging
Analysis for Space and Extreme Environments, pp. 33–40. CRC
Press, ??? (2021)
10. Hasan, M.A., Rashmi, S., Esther, A., Bhavanisankar, P.Y.,
Sherikar, B.N., Sridhara, N., Dey, A.: Evaluations of silica
aerogel-based flexible blanket as passive thermal control element for spacecraft applications. J Mater Eng Perform 27(3),
1265–1273 (2018)
11. Petroffe, G., Beouch, L., Cantin, S., Chevrot, C., Aubert, P.-H.,
Dudon, J.-P., Vidal, F.: Thermal regulation of satellites using
adaptive polymeric materials. Solar Energy Mat Solar Cells
200, 110035 (2019)
12. Nikpourian, H., Bahramian, A.R.: Thermo-physical properties of multilayer super insulation: The role of aerogel blanket.
Thermal Sci Eng Progress 20, 100751 (2020)
13. Fesmire, J.E.: Aerogel-based insulation materials for cryogenic
applications. In: IOP Conference Series: Materials Science and
Engineering, vol. 502, p. 012188. IOP Publishing (2019)
14. Pilipenko, V., Yagova, N., Romanova, N., Allen, J.: Statistical
relationships between satellite anomalies at geostationary orbit
and high-energy particles. Adv Space Res 37(6), 1192–1205
(2006)
15. Han, J.-H., Kim, C.-G.: Low earth orbit space environment simulation and its effects on graphite/epoxy composites. Composite
structures 72(2), 218–226 (2006)
16. Flegel, S.K., Gelhaus, J., Möckel, M., Wiedemann, C., Krag,
H., Klinkrad, H., Vörsmann, P.: Multi-layer insulation model for
master-2009. Acta Astronautica 69(11–12), 911–922 (2011)
17. Krag, H., Klinkrad, H., Jehn, R., Flegel, S., Schildknecht, T.:
Conclusions from esa space debris telescope observations on
space debris environment modelling. In: Proceedings of the
European Conference on Space Debris, Darmstadt, Germany
(2009)
18. Früh, C., Kelecy, T.M., Jah, M.K.: Coupled orbit-attitude
dynamics of high area-to-mass ratio (hamr) objects: influence
of solar radiation pressure, earth’s shadow and the visibility in
light curves. Celestial Mechanics Dynamical Astronomy 117(4),
385–404 (2013)
19. Cowardin, H., Abercromby, K., Barker, E., Seitzer, P., Mulrooney, M., Schildknecht, T.: An assessment of geo orbital
debris photometric properties derived from laboratory-based
measurements. In: Advanced Maui Optical and Space Surveillance Technologies Conference, 1, 25 (2009)
20. Barker, E., Abercromby, K.J., Abell, P.: Reflectance spectra
comparison of orbital debris, intact spacecraft, and intact rocket
13
J. R. Shah et al.
21.
22.
23.
24.
25.
26.
27.
28.
29.
30.
31.
32.
33.
34.
35.
36.
bodies in the geo regime. In: 5th European Conference on Space
Debris (2009)
Schildknecht, T., Vananti, A., Krag, H., Erd, C.: Physical characterization of high amr debris by optical reflectance spectrometry.
In: Proceedings of the International Astronautical Congress, p. 6
(2010)
Steckenreiter, T., Balanzat, E., Fuess, H., Trautmann, C.: Chemical modifications of pet induced by swift heavy ions. Nuclear
Instruments Methods Phy Res Sect B: Beam Interactions with
Mater Atoms 131(1–4), 159–166 (1997)
Plis, E.A., Engelhart, D.P., Cooper, R., Johnston, W.R., Ferguson,
D., Hoffmann, R.: Review of radiation-induced effects in polyimide. App Sci 9(10), 1999 (2019)
Czeremuszkin, G., Latreche, M., Wertheimer, M.: Charging/discharge events in coated spacecraft polymers during electron beam
irradiation in a scanning electron microscope. Nuclear Instrument
Methods Phy Res Sect B: Beam Interact with Mater Atoms 185(1–
4), 88–99 (2001)
Hoffmann, R., Cooper, R., Engelhart, D., Plis, E., Ferguson, D.:
Localized electron trap modification as a result of space weather
exposure in highly disordered insulating materials. Technical
report, AIR FORCE RESEARCH LAB KIRTLAND AFB NM
KIRTLAND AFB (2017)
Griseri, V.: Behavior of dielectrics in a charging space environment and related anomalies in flight. IEEE Transa Dielect Electrical Insulation 16(3), 689–695 (2009)
Edwards, D.L., Hubbs, W.S., Wertz, G.E., Hoppe, D.T., Nehls,
M.K., Hollerman, W.A., Gray, P.A., Semmel, C.L.: Electron radiation effects on candidate solar sail material. High Perform Polymers 16(2), 277–288 (2004)
Kupchishin, A.I., Taipova, B.G., Voronova, N.A., Kupchishin,
A.A., Shakhanov, K.: Dose dependences of the mechanical
strength of polymer materials. In: Key Engineering Materials,
vol. 781, pp. 25–29. Trans Tech Publ (2018)
Engelhart, D.P., Murray, V.J., Plis, E.A., Fulford, K., Ferguson,
D.C., Hoffmann, R.: Effect of electron flux on the degradation
of organic polymers in a simulated geo environment. In: AIAA
Scitech 2021 Forum, p. 0066 (2021)
Engelhart, D.P., Reyes, J.A., Murray, V.G., Ferguson, D.C., Hoffmann, R.C.: Identifying radiation induced spectroscopic changes
in polyimide materials. The Journal of the Astronautical Sciences,
1–11 (2022)
Reyes, J.A., Cowardin, H.M.: Spectral characterization of spacecraft materials used in hypervelocity impact testing. In: Algorithms, Technologies, and Applications for Multispectral and
Hyperspectral Imaging XXVII, vol. 11727, p. 117271 (2021).
International Society for Optics and Photonics
Reddy, V., Battle, A., Campbell, T., Chodas, P., Conrad, A., Engelhart, D., Frith, J., Furfaro, R., Farnocchia, D., Hoffmann, R., et al.:
Spectral characterization of 2020 so (2021)
Payne, T.E., Castro, P.J., Moody, J.W., Beecher, E.A., Fisher,
M.D., Acosta, R.I.: A discrimination analysis of sloan and johnson photometric systems for non-resolved object characterization.
In: Advanced Maui Optical and Space Surveillance Technologies
Conference (AMOS) Conference Proceedings (2016)
Payne, T.E., Gregory, S.A., Dentamaro, A., Ernst, M., Hollon,
J., Kruchten, A., Chaudhary, A.B., Dao, P.D.: Development and
evaluation of new methods for estimating albedo-areas for threeaxis stabilized geosynchronous satellites. J Astro Sci 66, 170–191
(2019)
Cowardin, H., Seitzer, P., Abercromby, K., Barker, E., Buckalew,
B., Cardona, T., Krisko, P., Lederer, S.: Observations of titan iiic
transtage fragmentation debris. In: Advanced Maui Optical and
Space Surveillance Technologies (AMOS) Conference (2013)
Beamer, D., Abeywickrema, U., Banerjee, P.: Polarization vector
signatures for target identification. In: Polarization Science and
13
37.
38.
39.
40.
41.
42.
43.
44.
45.
46.
47.
48.
49.
50.
51.
52.
Remote Sensing VIII, vol. 10407, p. 104070. International Society
for Optics and Photonics (2017)
Pearce, E.C., Ford, H.A., Schildknecht, T., Reddy, V., Block, A.D.,
Rockowitz, K.: Rapid characterization of geosynchronous space
debris with 5-color near-ir photometry. In: Proceedings AMOS
Conference (2017)
Kent, S.M.: Sloan digital sky survey. Astrophysics and Space Science 217(1), 27–30 (1994)
Reyes, J.A.: Reflectance spectral characterization and taxonomy
applications of spacecraft materials to aid space situational awareness (2022)
Cooper, R., Hoffman, R.: Jumbo space environment simulation
and spacecraft charging chamber characterization. Technical
report, Air Force Research Laboratory, Space Vehicles Directorate Kirtland AFB (2015)
Pal, P., Varshney, V., Gupta, D.: Approximate expressions for
continuous slowing down approximation ranges of high energy
electrons and positrons. Nuclear Instruments Method Phy Res
Sect B: Beam Interactions with Materials Atoms 21(1–4), 14–19
(1987)
Engelhart, D.P., Plis, E., Humagain, S., Greenbaum, S., Ferguson,
D., Cooper, R., Hoffmann, R.: Chemical and electrical dynamics
of polyimide film damaged by electron radiation. IEEE Transact
Plasma Sci 45(9), 2573–2577 (2017)
Bengtson, M., Maxwell, J., Hoffmann, R., Cooper, R., Schieffer,
S., Ferguson, D., Johnston, W.R., Cowardin, H., Plis, E., Engelhart, D.: Optical characterization of commonly used thermal control paints in a simulated geo environment. In: The Advanced
Maui Optical and Space Surveillance Technologies Conference,
p. 33 (2018)
Frederickson, A.R., Dennison, J.R.: Measurement of conductivity
and charge storage in insulators related to spacecraft charging.
IEEE transact nuclear sci 50(6), 2284–2291 (2003)
Plis, E.A., Engelhart, D.P., Likar, J., Hoffmann, R.C., Cooper,
R., Ferguson, D.: Electrical behavior of carbon-loaded kapton
for spacecraft applications. J Spacecraft Rockets 55(3), 775–777
(2018)
Plis, E.A., Engelhart, D.P., Cooper, R., Ferguson, D.C., Hoffmann, R.: Effect of environment on charge transport properties
of polyimide films damaged by high-energy electron radiation. J
Vacuum Science Tech B, Nanotech Microelect: Materials, Proces,
Measurement, Phenomena 36(5), 052906 (2018)
Collman, S., Plis, E.A., Shah, J.R., Bengtson, M.T., Hoffmann,
R.C., Ferguson, D.C.: Operational procedure for handling of
spacecraft materials sensitive to atmospheric exposure. Journal
of Spacecraft and Rockets, 1–4 (2022)
Arino, I., Kleist, U., Rigdahl, M.: Effect of gloss and texture on
the color of injection-molded pigmented plastics. Polymer Eng
Sci 45(5), 733–744 (2005)
Xu, R., Fleming, P.D., Pekarovicova, A., Bliznyuk, V.: The effect
of ink jet paper roughness on print gloss. J Imag Science Technol
49(6), 660–666 (2005)
Mishra, R., Tripathy, S., Sinha, D., Dwivedi, K., Ghosh, S., Khathing, D., Müller, M., Fink, D., Chung, W.: Optical and electrical properties of some electron and proton irradiated polymers.
Nuclear Instruments and Methods in Physics Research Section
B: Beam Interactions with Materials and Atoms 168(1), 59–64
(2000)
Chaudhary, N., Koiry, S., Singh, A., Tillu, A., Jha, P., Samanta, S.,
Debnath, A., Aswal, D., Mondal, R., Acharya, S.: Electron beam
induced modifications in flexible biaxially oriented polyethylene
terephthalate sheets: Improved mechanical and electrical properties. Materials Chem Phy 189, 237–244 (2017)
Oproiu, C., Martin, D., Toma, M., Marghitu, S., Jianu, A.: Transitory and permanent effects of electron beam irradiation on insulating materials. Nuclear Instruments and Methods in Physics
Characterization of polyethylene terephthalate (PET) materials under high‑energy electron…
Research Section B: Beam Interactions with Materials and Atoms
166, 669–675 (2000)
53. Ting, Y.-H., Liu, C.-C., Park, S.-M., Jiang, H., Nealey, P.F.,
Wendt, A.E.: Surface roughening of polystyrene and poly (methyl
methacrylate) in Ar/O 2 plasma etching. Polymers 2(4), 649–663
(2010)
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Spacecraft Material Characterization Using Reflectance Spectra
Extracted from RGB/IR Color Images
Jainisha R. Shah∗
Assurance Technology Corporation, Carlisle, MA, 01741, USA
Downloaded by GA Inst of Technology on May 9, 2023 | http://arc.aiaa.org | DOI: 10.2514/6.2023-1957
Miles T. Bengtson†
National Research Council Research Associateship Program, 3550 Aberdeen Ave., Kirtland AFB, New Mexico 87117,
USA
Ryan C. Hoffmann, Dale C. Ferguson‡
Air Force Research Laboratory, Space Vehicles Directorate, USA, Kirtland AFB, Albuquerque, NM, 87117, USA
Sydney E.J. Collman§
Assurance Technology Corporation, Carlisle, MA, 01741, USA.
Elena A. Plis¶
Georgia Tech Research Institute (GTRI), Atlanta, GA, 30318, USA; Assurance Technology Corporation, Carlisle, MA,
01741, USA
Daniel P. Engelhart‖
University of New Mexico, Albuquerque, NM, 87131, USA
Gregory P. Badura∗∗
Georgia Tech Research Institute (GTRI), 925 Dalney St NW, Atlanta, GA 30318
Heather M. Cowardin††
NASA Johnson Space Center, Orbital Debris Program Office, 2101 NASA Pkwy., Houston, TX 77058, USA
Timothy R. Scott‡‡
DuPont de Nemours, Inc, Durham, NC 27703, USA
The space environment is detrimental to the exterior of the spacecraft and imposes stringent
requirements on spacecraft materials. Space materials are often expected to perform for
15-20 years in the harsh space environment without significant degradation of their optical,
electrical, mechanical and thermal properties. Further, as commercial activities increase in the
space domain, development of predictive capability for material evolution is critical for agile
adaption of novel materials. Therefore, it is very important to understand the effects of the space
environments on the spacecraft materials. The Materials on the International Space Station
∗ Scientist, Assurance Technology Corporation, Carlisle, MA, 01741, USA.
† Postdoctoral Fellow, National Research Council Research Associateship Program, 3550 Aberdeen Ave., Kirtland AFB, New Mexico 87117,
USA.
‡ Air Force Research Laboratory, Space Vehicles Directorate, USA, Kirtland AFB, Albuquerque, NM, 87117, USA
§ Assurance Technology Corporation, Carlisle, MA, 01741, USA.
¶ Senior Research Engineer, Georgia Tech Research Institute (GTRI), 925 Dalney St NW, Atlanta, GA 30318; Assurance Technology Corporation,
Carlisle, MA, 01741, USA
‖ Research Professor, Department of Chemistry and Chemical Biology, University of New Mexico, 300 Terrace St NE, Albuquerque, NM, 87131.
∗∗ Research Scientist, Georgia Tech Research Institute (GTRI), 925 Dalney St NW, Atlanta, GA 30318.
†† Laboratory and In Situ Lead Orbital Debris Business Unit Manager, NASA Johnson Space Center, Orbital Debris Program Office, 2101 NASA
Pkwy., Houston, TX 77058, USA
‡‡ Business Development Leader, Aerospace & Defense, DuPont de Nemours, Inc, Durham, NC 27703.
1
Copyright © 2023 by the American Institute of Aeronautics and Astronautics, Inc. All rights reserved.
Downloaded by GA Inst of Technology on May 9, 2023 | http://arc.aiaa.org | DOI: 10.2514/6.2023-1957
Experiment Flight Facility (MISSE-FF) has flown numerous material samples to investigate
the effect of LEO space weather exposure on the performance and durability of materials and
devices. As part of the MISSE-16 mission, which launched on July 15, 2022, we are flying fifteen
novel and well-characterized materials in the LEO environment for a duration of six months.
Changes in spectral reflectivity will be measured in real time throughout the mission with an
RGB/IR camera. These time-resolved data will function as a “space-truth” reference for our
team’s ongoing laboratory-based space weather-material interaction experiments. Correlation
of the MISSE-16 data with extensive ground testing of duplicate samples under simulated
space weather conditions will enable development of fundamental chemical models for material
degradation. This paper discusses preliminary results from the ground test campaign to
collect the RGB/IR images for pristine and damaged materials and the development of machine
learning algorithms to extract reflectance spectra from the color images.
I. Introduction
xterior spacecraft materials play an important role in protecting satellites from the harsh space environment and
are disproportionately understudied in the literature. Research has shown that the physical, chemical, and optical
properties of matter change continuously as a result of exposure to solar radiation and aggressive chemical species
produced in Earth’s upper atmosphere [1]. Thorough knowledge of the material properties evolution throughout a
planned mission lifetime helps to improve the reliability of spacecraft. Moreover, the establishment of correlation
factors between true space exposure and accelerated space weather experiments at ground facilities enables accurate
prediction of on-orbit material performance based on laboratory-based testing.
The Materials International Space Station Experiment Flight Facility (MISSE-FF) on the International Space Station
(ISS) is the perfect test bed to generate benchmark data with which to validate the efficacy of ground-based space weather
simulation experiments. Our ultimate goal is to fly several novel and well-understood materials on the MISSE-FF and
monitor the changes in spectral reflectivity that occur as a result of exposure to different components of the low Earth
orbit (LEO) environment. Energy deposition from the environment leads to chemical changes in the material that in
turn alter the optical properties. Since unique damage occurs at each face of ISS, the mission will acquire spectroscopic
data of identical material sets placed on ram (direction of travel, AO exposure), zenith (direction away from Earth,
VUV exposure), and wake faces (opposite to the travel direction, electron exposure). MISSE flight duration will be
approximately six-months; daily images of each sample will be taken for the first 30 days; weekly images of each sample
for the next two months; monthly images are then taken for the remaining duration of the mission. This experiment will
allow us to observe the progression of space weather damage in real time which can then be compared to ground-based
space weather simulations. Figure 1 shows an illustration of the MISSE-FF concept of operations.
E
II. Experimental Details
MISSE flight sample collection comprises several different class of polymers including polyimides (PIs), polyethylene
terephthalates (PET) materials, liquid crystal polymers (LCP), PI/Polyhedral Oligomeric Silsesquioxanes (POSS)
materials and carbon and glass reinforced polymers (CFRP and GFRP). Table 1 below lists the materials and the material
class that they belong to.
After deploying the MISSE-16 mission on the International Space Station (ISS) in August 2022, we expect to
begin receiving red, green, blue and infrared (RGB/IR) images of the materials throughout the 6 months mission.
These time-resolved data will function as a “space-truth” reference for our team’s ongoing laboratory-based space
weather-material interaction experiments. The RGB/IR color data obtained from the experiment will be the first of its
kind and allows for the observation of the space weather damage on the materials in real time.
Further, within the framework of the same project, replica of the hardware used on MISSE to collect the reflectance
data has been installed in the Jumbo space irradiation chamber at the Spacecraft Charging and Instrument Calibration
Laboratory (SCICL) at the Air Force Research Laboratory [2]. Figure 2 shows the ground test unit of the MISSE camera
and light setup installed in the SCICL facility.
The reflectance data of control (pristine) and material aged under the LEO-simulated environment (i.e., electron and
VUV – exposed) will be collected using the same lighting and camera system as installed on board the MISSE-FF.
Figure 3 shows the MISSE camera hardware with the “white” LED illumination spectrum and the camera response for
each RGB/IR band.
2
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Fig. 1 Planned MISSE experiment aboard the ISS in 2022. Materials will be illuminated with infrared (IR)
and white light sources, and imaged with an RGB/IR camera on a periodic basis. Materials will be mounted in
positions of varying ultraviolet radiation, ionizing particle (electron) radiation, and atomic oxygen exposure.
(a) Camera and lights in front of the sample carousel
Fig. 2
(b) Camera apparatus with only the white lights on
MISSE camera board installation in SCICL Jumbo Chamber
III. Preliminary Analysis
The large amount of data collected for these experiments will be analyzed using machine learning techniques
developed by our colleagues at the Georgia Tech Research Institute (GTRI). To train the machine learning algorithm
initially, color and IR images of Spectralon color standards with well-characterized reflectance spectra were collected
using the MISSE camera hardware. Figure 4 shows several of these images and illustrates the difference between
visible and IR images. For each of the materials, pristine visible and IR images as well as the directional hemispherical
reflectance (DHR) spectra were collected.
The materials will then be subjected to electron irradiation and VUV irradiation to simulate damage seen by the
samples on each face of the ISS. Visible and IR camera images and DHR measurements will be performed at varying
time intervals to monitor the changes in the optical properties of the materials and will be used to continue training the
machine learning algorithm. Figure 5 is an example of visible and IR images for two of the pristine materials Kapton
HN®and GFRP.
3
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Table 1
No.
1
2
3
4
5
6
7
8
9
10
11
12
13
14
15
List of materials studied for the MISSE-FF
Material
Polyimide (PMDA)
Polyimide (PMDA)
Polyimide (PMDA)
Polyimide (PMDA)
Polyimide (PMDA)
Polyimide (PMDA)
Polyimide (PMDA)
Carbon Fiber Reinforced Polymer
Glass Fiber Reinforced Polymer
Liquid Crystal Polymer
Polyethylene terephthalate
Polyethylene terephthalate
PI/Polyhedral Oligomeric Silsesquioxane
PI/Polyhedral Oligomeric Silsesquioxane
Aluminum Oxide
Abbrev.
PI
PI
PI
PI
PI
PI
PI
CFRP
GFRP
LCP
PET
PET
POSS
POSS
ALUM
Trade Name
Kapton CR®
Kapton CS®
Kapton WS®
Kapton XC®
Kapton TF®
Kapton 200DR9®
Kapton HN®
Economyplate™Carbon Fiber
G-10/FR4 Glass Epoxy
Zenite®
Melinex®454
Mylar®MO21
CORIN®XLS
Thermalbright®N
Alumina
a *Trade names and trademarks are used in this report for identification only. Their usage does not constitute
an official endorsement, either expressed or implied, by the National Aeronautics and Space Administration.
Fig. 3
Illumination/Detection scheme for the MISSE experiment
IV. Machine Learning
A key factor in fully characterizing the materials flying on MISSE is the ability to extract reflectance spectra
from RGB/IR images. Materials damaged in the terrestrial vacuum chambers can be characterized with a suite of
4
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Fig. 4
Visible and IR LED camera calibration with colored Spectralon standards
laboratory instruments. On-orbit experiments, however, are typically limited to only capturing pictures or very limited
characterizations of material properties. Therefore, we are developing a method to extract reflectance spectra from
RGB/IR images. Prior work has addressed how reflectance spectra can be used as a proxy to determine other important
material properties such as conductivity and embrittlement (i.e., [3, 4]). Thus, developing a method to estimate
reflectance spectra from camera images will allow future experiments to obtain more detailed information about material
evolution.
Spectral extraction from camera images is an under constrained problem since it involves mapping from a lowdimensional space (RGB/IR pixel counts) to a high-dimensional space (reflectance as a function of wavelength). In
other words, there are an infinite number of spectra which can produce a given RGB/IR image. Machine learning
techniques are often used to approach such problems. We are developing a machine learning approach which uses
a radial basis function (RBF) network [5]. The RBF network is a proven method in the color-vision community for
performing a non-linear interpolation of sparsely spaced data. This feature is critical for our dataset because the data is
non-uniformly distributed within our chosen colorspace, as is shown in Figure 6.
The RBF model in this study is a mapping ( 𝑓 ) between input color data and a spectral reflectance data. The mapping
can be denoted as 𝑓 : R4 → R 𝑘 , where 4 denotes the 4 color channels (L*a*b and IR) and 𝑘 denotes the number of
spectral channels, with 𝑘 > 4. The mapping is given by the equation [5]:
𝑀
∑︁
||x − 𝜇i ||
𝑓 (x) = 𝜔0 +
𝜔𝑖 𝜙
𝜎𝑖
𝑖=1
(1)
where x ∈ R4 is the input color data, 𝑓 (x) ∈ R 𝑘 is the mapped spectral reflectance vector, 𝜙(.) denotes the radial basis
function, ||.|| denotes the Euclidean distance function, 𝜔𝑖 denotes the weight of the 𝑖 𝑡 ℎ RBF, 𝜇i ∈ R4 denotes the center
of the 𝑖 𝑡 ℎ RBF in color space, and 𝜎𝑖 is a scalar denoting the radius of the 𝑖 𝑡 ℎ RBF. A custom Keras layer representing
this RBF node was written in order to learn the trainable parameters (𝜇i , 𝜙i , and 𝜔𝑖 ) via standard backpropagation
algorithm approaches [6].
The RBF model will be trained using camera images and DHR spectra for numerous pristine and damaged material
samples collected in the ground test campaign. When images are downloaded from the on-orbit MISSE camera, the
RGB data will be selected from a non-over-saturated region under both visible and IR lighting conditions. Next, these
pixel counts will be white balanced given the scene context, which is possible because the same exposure settings have
been applied to the flight and ground test cameras and a white Spectralon standard will be visible in each of the ground
test images. Finally, the RGB data from MISSE will be converted into L*a*b* color space and fed into the trained
Radial Basis Function (RBF) network. The machine learning algorithm will then give an estimated reflectance spectrum
for each image. The entire training and subsequent reflectance retrieval pipeline is outlined in Figure 7.
5
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Fig. 5 Visible and IR illuminated camera images for pristine Kapton HN®(a and b) and pristine GFRP (c and
d)
An example of the error in spectrum retrieval for the hypertuned RBF network model is shown in Figure 8. The
errors are shown for the validation portion of our ground test set. The validation portion of this dataset comprised
approximately 70 samples of color (L*a*b* and IR) and spectral information that the RBF network model did not see
during the training procedure. This model was rather trained on a separate 140 samples of color and spectral information.
In other words, there was a 2/1 split in the training and validation datasets, respectively. The dark lines represent the
mean reconstruction error and the transparent bars represent the spread of the error.
From these results, we can see that the model is able to retrieve the spectrum to within 10% error for the vast
majority of samples across wavelengths of 350 nm to 900 nm. The lowest overall errors occur near the green (500 nm)
and red (650 nm) wavelength regions, suggesting that we can place the highest confidence in spectrum retrieval for these
regions. The highest overall errors seem to occur in the blue (400 nm) region. Surprisingly, it appears that there are
difficulties in reconstructing the spectra of some of the white samples (Alumina and Kapton WS) in the blue wavelength
region. This could be due to a number of factors such as lower spectral responsivity of the camera system in the blue
region or a limited number of samples spanning the convex hull of color information in the blue region of L*a*b*IR
space. These issues will be considered when retrieving spectra from on-orbit data.
V. Conclusion
Correlation of the MISSE data with extensive ground testing of flight duplicate samples under simulated space
weather conditions will enable development of fundamental chemical models for material degradation. Spacecraft
designers will benefit from the insight into material behavior changes throughout a mission lifetime. Further, a detailed
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Fig. 6
A scatter plot showing the distribution of our ground test sample data in the L*a*b* color space.
Fig. 7
Diagram illustration the training procedure for the RBF network.
predictive knowledge of space weather-induced material color change will enable more robust and accurate space domain
awareness by allowing remote observers to glean knowledge about a spacecraft by examining the spectral signature of
unresolved images. Use of spacecraft color in this way will help with enable remote diagnosis and anomaly resolution.
7
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Fig. 8
RBF network spectral reconstruction error on the validation portion of the ground dataset.
VI. Acknowledgements
Authors would like to thank DuPont de Nemours, Inc, for providing polyimide materials for this research.This work
was partially supported by the Air Force Office of Scientific Research, Remote Sensing and Imaging Portfolio (Dr.
Michael Yakes) Grant 20RVCOR024 and Georgia Tech Research Institute Independent Research and Development
(IRAD) program.
VII. Disclaimers
The views expressed are those of the author and do not necessarily reflect the official policy or position of the
Department of the Air Force, the Department of Defense, or the U.S. government.
Trade names and trademarks are used in this report for identification only. Their usage does not constitute an official
endorsement, either expressed or implied, by the National Aeronautics and Space Administration.
8
References
[1] Engelhart, D., Cooper, R., Cowardin, H., Maxwell, J., Plis, E., Ferguson, D., Barton, D., Schiefer, S., and Hoffmann, R., “Space
weathering experiments on spacecraft materials,” The Journal of the Astronautical Sciences, Vol. 66, No. 2, 2019, pp. 210–223.
[2] Cooper, R., and Hoffman, R. C., “Jumbo Space Environment Simulation and Spacecraft Charging Chamber Characterization,”
2015.
[3] Plis, E. A., Engelhart, D. P., Cooper, R., Ferguson, D. C., and Hoffmann, R. C., “Physical and Spectrometric Analysis of
Electron-Damaged LDPE,” IEEE Transactions on Plasma Science, Vol. 47, No. 8, 2019, pp. 3752–3758.
[4] Rahnamoun, A., Engelhart, D. P., Humagain, S., Koerner, H., Plis, E., Kennedy, W. J., Cooper, R., Greenbaum, S. G., Hoffmann,
R., and van Duin, A. C., “Chemical dynamics characteristics of Kapton polyimide damaged by electron beam irradiation,”
Polymer, Vol. 176, 2019, pp. 135–145.
Downloaded by GA Inst of Technology on May 9, 2023 | http://arc.aiaa.org | DOI: 10.2514/6.2023-1957
[5] Nguyen, R. M., Prasad, D. K., and Brown, M. S., “Training-based spectral reconstruction from a single RGB image,” European
Conference on Computer Vision, Springer, 2014, pp. 186–201.
[6] Chollet, F., et al., “Keras,” https://keras.io, 2015.
9
in the Space
Environment
Judith C. Yang and Kim K. de Groh,
Guest Editors
Abstract
To explore higher, farther, and faster, scientists and engineers have developed
advanced materials for manned spacecraft and satellites for a range of sophisticated
applications in transportation, global positioning, exploration, and communication.
Materials used in space are exposed to vacuum, intense ultraviolet radiation from the
sun, and ionizing radiation that results in material damage as well as charging
(electrostatic discharge effects), micrometeoroids and debris impacts, and thermal
cycling (typically from −175 to 160°C). In terms of materials degradation in space, the
low Earth orbit (LEO), where LEO is defined as 200–1000 km above the Earth’s
surface, is a particularly challenging synergistic environment, since atomic oxygen
(AO) is present along with all other environmental elements. Hence, this special issue
focuses primarily on the materials issues experienced in LEO by space environmental
exposure, such as on the exterior of the International Space Station and the Hubble
Space Telescope, and the challenges and opportunities of ground-based laboratory
sources to mimic LEO. The combination and comparison of both in-flight and groundbased experiments are needed for the development of predictive understanding of the
materials degradation and AO passivation mechanisms in LEO. Such insights are
essential for the development of advanced materials and coatings to ensure the longterm durability and performance of vehicles employed in space.
Introduction and Background
Space exploration began with gazing at
the stars; as early as 3000 BC, Babylonian
astronomers made methodical observations of the skies. The telescope was
invented in ~1600 AD, and Galileo Galilei
introduced the telescope to the study of
astronomy in 1609; the developments and
utilization of the telescope have had a
tremendously profound impact on our
understanding of the universe. However,
it was not until the last century that we
accomplished the spectacular achievement
of sending engineered spacecrafts and
humans into space. The first successful
orbital launch was of the Soviet unmanned
Sputnik (“Satellite I”) mission on October
4, 1957, which weighed about 83 kg
and orbited Earth at a height of ~250 km.
By 1960, the first communications and
weather satellites were launched. By 1969,
Neil Armstrong became the first human to
step on the moon. As space programs
developed during the past 50 years, the
focus has shifted from single-use space
vehicles to re-usable and long-term vehicles, such as the Space Shuttle and the
International Space Station (ISS)—both of
which are substantial engineering achievements—that must withstand long-term
exposure to the harsh space environment.
The orbits in space are classified as geostationary Earth orbit (GEO), which is a
geosynchronous orbit 36,000 km directly
above the equator, and medium Earth
orbit (MEO), which is between GEO and
the low Earth orbit (LEO), where LEO is
defined as 200–1000 km above the Earth’s
surface. Each orbit is used for different
applications. The GEO is primarily for
communication satellites; MEO is used for
navigation satellites such as the Global
Positioning System; and LEO is for Earth
12
https://doi.org/10.1017/S0883769400100004 Published online by Cambridge University Press
700
600
Altitude (km)
Materials Issues
observation for scientific or military
purposes as well as manned spacecraft,
such as the ISS. In each orbit, external
spacecraft materials are subjected to
vacuum, intense ultraviolet (UV) radiation from the sun, ionizing particle
radiation (protons/electrons), micrometeoroids and debris impacts, and thermal
cycling (typically from −175 to 160°C), all
of which can result in material damage.
In addition, on-orbit charging (electrostatic
discharge) can affect spacecraft components. Each orbit has its own peculiarities:
in GEO, electrostatic discharge is of more
concern than in other orbits; in MEO, particle radiation levels can be especially high
(Van Allen belts); and in LEO, atomic
oxygen (AO) is an additional concern not
present in higher altitude orbits. AO is
generated by the photo-dissociation of
molecular oxygen by intense UV sunlight,
where the vacuum conditions are
sufficiently low in LEO (Figure 1), such
that the probability for the AO to recombine is low. Since spacecraft travel is at
~8 km s−1 in LEO, the space vehicles ram
into the AO with a relative kinetic energy
of ~4.5 eV. Therefore, in terms of materials
in space, LEO is an intriguing and complex environment, since all environmental
elements occurring in space are present
(Figure 2), if not always to their maximum extent. Hence, this special issue
focuses primarily on the materials issues
experienced in LEO.
As shown in Figure 1, AO is the most
predominant species in LEO. It is also
present in other planetary orbital environments, such as in low Mars orbit. At
spacecraft velocities, AO is energetic
enough to cause bond breakage and subsequent oxidation in many spacecraft
materials. The oxidation products of most
polymers are gas species; therefore, extensive material erosion occurs.1–5 AO can
produce serious structural, thermal, or
optical degradation of spacecraft components.6,7 Figure 3 shows AO erosion of
Teflon-fluorinated ethylene propylene
around a small protective particle after
H
500
N2
400
O2
300
He
O
107
108
Ar
200
100
104
105
106
109
1010
1011
1012
Number Density (cm–3)
Figure 1. Atmospheric composition in
the low Earth orbit.30
MRS BULLETIN • VOLUME 35 • JANUARY 2010 • www.mrs.org/bulletin
Materials Issues in the Space Environment
direct ram exposure in space.8 Several of
the articles in this issue emphasize polymer, carbon-based, and coating materials
because these are essential components to
space vehicles—such as structural materials for the hull and manipulator arm,
paints to control the solar absorption and
thermal emissivity, sealants, and solar
array blankets. Polymers such as Kapton
and Teflon-fluorinated ethylene propylene are also commonly used as thermal
insulation blankets, often covering large
surface areas on spacecraft, such as on the
Hubble Space Telescope.
Degradation of Materials in Space
Comparison between Space Flight
Experiments and Ground-Based
Testing
To determine mechanisms of AO
erosion in the space environment, both inspace and ground-based laboratory experiments are necessary. Though in-space
experiments are expensive and difficult to
control, actual space flight data of the AO
erosion yield (volume loss in cm3 per incident oxygen atom) is needed to assess
the durability of a material for mission
applicability, since, currently, ground AO
facilities do not perfectly simulate all
aspects of the space environment.5
The first long-duration in-space materials exposure experiment was the Long
Duration Exposure Facility (LDEF),
deployed on April 7, 1984 and retrieved on
January 12, 1990. LDEF was a school bussized near cylindrical structure that flew
with 57 science and technology experiments.9,10 Since the LDEF remained in LEO
for 5.8 years, NASA discovered that the
LEO environment damages many materials dramatically, especially polymers and
some metals.5,11 Because of the intriguing
information from LDEF and observations
from other spacecraft and ground tests,
additional in-space experiments have been
flown. The Materials International Space
Station Experiment (MISSE) is a series of
materials flight experiments consisting of
trays, called passive experiment containers,
which are exposed to the space environment on the exterior of the ISS (Figure 4).
The objective of MISSE is to test the longterm stability and durability of materials
and devices in the space environment. The
MISSE 6 mission (MISSE 6A and 6B) was
launched on March 13, 2008, on the space
shuttle Endeavour; it contained more than
1200 material samples along with quartz
crystal monitor experiments and other
active experiments for real-time kinetic
information on the degradation of
a variety of materials. Examples of the
many different types of materials flown on
MISSE 6 include thin-film polymers,
metals, metal oxides, semiconductor materials, shape-memory composites, ceramic
materials, optical reflector materials, radiation shielding materials, thermal control
coatings, and protective materials and
coatings. Numerous active experiments
were also flown, such as environment
monitors and detectors (for AO, radiation,
and low-energy ion detectors), calorimeters and emissivity sensors, and
microelectromechanical systems (MEMS)
devices. These materials, sensors, and
devices are the efforts of many organizations, including various NASA centers,
U.S. Department of Defense agencies, the
space industry, and several universities in
the United States.
Spacecraft designers must take into
account not only the physical and mechanical properties of materials they are proposing for use in space with regard to their
intended function but also what the
expected properties will be at the end of
the mission due to degradation by the
environment to which they will be
subjected. Because spacecraft are often
used for exploration and discovery to usually unknown environments, the first indication of an environmental problem is
usually through performance failure. Even
with some advanced knowledge of the
environment in LEO, unforeseen degradation issues can still occur, such as those
experienced on the solar array blanket box
of the ISS, as shown in Figure 5, or the
Primary Degradaon Mechanisms in LEO
Fast Atomic Oxygen
Hyperthermal
4.5 eV
Vacuum Ultraviolet
Radiaon
Thermal Cycling
–175 to +160°C
Secondary Degradaon Mechanisms in LEO
Debris/Meteoroid
Impacts
Solar Radiaon
X-rays
Electron + Proton
Damage
Figure 2. The primary and secondary
degradation mechanisms in the low
Earth orbit.
Figure 3. Atomic oxygen erosion of
Teflon-fluorinated ethylene propylene in
space.8
Figure 4. The International Space Station (photographed in March 2008) with a close-up
photo of MISSE 6A and 6B on the Columbus Laboratory.31
MRS BULLETIN • VOLUME 35 • JANUARY 2010 • www.mrs.org/bulletin
https://doi.org/10.1017/S0883769400100004 Published online by Cambridge University Press
13
Materials Issues in the Space Environment
Figure 5. Atomic oxygen undercutting degradation of the solar array wing blanket box cover
on the International Space Station after one year of low Earth orbit space exposure.8,31,32
Figure 6. Severe degradation to the aluminized-Teflon outer layer of multilayer insulation on
the Hubble Space Telescope after 19 years of space exposure.
thermal control blankets on the Hubble
Space Telescope,6,12–17 as shown in Figure 6.
Material interactions with the environment and with combined environments
can be complex and difficult to predict
without validation testing.
Due to the need for validation testing
in ground facilities, spaceflight experiments are also needed to determine the
correlation factors between space exposure and ground facilities. This allows for
more accurate in-space performance
predictions to be made based on groundlaboratory testing, since multiple differences exist between ground facilities and
space exposure.12 For AO testing, these
differences include variations in species,
energies, thermal exposures, and radiation exposures, all of which may result in
different synergistic reactions and erosion
rates.18 Past ground-based radiation environmental durability tests have indicated
that exposing materials in accelerated
tests to environmental model-predicted
spacecraft mission elements does not simulate the extent of damage that occurs in
the space environment.18 One approach to
overcoming the difficulties in simulating
the space environment using groundbased testing is to calibrate the facility
using data from actual space-exposed
materials (e.g., ISS, Hubble Space Telescope)
to determine exposure levels required to
replicate degraded properties observed in
space, which is discussed by Miller and
Banks in this issue.
Though the article by Miller and Banks is
focused on the LDEF, Hubble Space
Telescope, and the MISSE missions as
examples of materials degradation in the
LEO, numerous international spaceflight
missions have been conducted by space
agencies besides NASA, including the
European Space Agency and the Japan
Aerospace Exploration Agency, and are
summarized by Edwards et al. in this issue.
Fundamental Mechanisms of
Oxidation by Hyperthermal Atomic
Oxygen
A large amount of empirical information
has been obtained about the AO degradation of a large variety of materials in the
harsh and dynamic space environment
14
https://doi.org/10.1017/S0883769400100004 Published online by Cambridge University Press
from LDEF and MISSE missions,3,7,9–11,18–22
but an additional fundamental understanding is needed. Gaining a better fundamental
understanding of the unusual oxidation
mechanisms that are occurring in LEO
requires an ideal laboratory source of AO
that best simulates the LEO environment;
this ideal source would produce a high flux
(>1015 atoms/cm2 s) of a direct beam of
mono-energetic (kinetic energy of 5
eV/atom) oxygen atoms in the ground electronic state (3P) with no impurities or UV
light present. This hyperthermal AO source
is a demanding engineering problem;5 fortuitously, the laser detonation source comes
closest to this ideal. A high-energy pulse of
CO2 laser energy is focused into a conical
expansion nozzle that has been partially
filled with O2 by a pulsed-beam valve that
ignites plasma at the nozzle throat. The
shock wave induced by the plasma dissociates the molecular oxygen to hyperthermal
AO. Only a handful of sites around the
world own such an AO source, primarily
for space programs (see the article by
Tagawa and Minton in this issue). The reactions of materials with hyperthermal AO
are different than those with thermal AO22,23
or molecular oxygen.22,24–26 Tagawa and
Minton describe fundamental mechanisms
of hyperthermal AO reactions discovered
by the utilization of this novel source under
well-controlled vacuum conditions.
Micrometeoroid and Debris Impacts
(MMOD)
Though AO is the primary species in
attacking materials in LEO, there are
other species that contribute to the degradation of materials in space (Figure 2).
For example, space debris, which consists
of hypervelocity objects produced by
satellites in orbit, and naturally occurring
ultrahigh velocity meteoroids both present a significant hazard to operational
spacecraft. During the last 30 years, the
number of spacecraft orbiting the Earth has
increased dramatically, leading to an enormous amount of space debris in
LEO. Collisions with space debris or
micrometeoroids can result in numerous
types of spacecraft degradations such as
dysfunctions in solar cells, degradation
in materials properties, contamination
of optical devices, or even destroying
satellites. These collisions also can form
new additional debris, thereby significantly increasing the erosion hazard
for future missions. As an example of
current MMOD studies, NASA Glenn
Research Center is studying the effect
of MMOD impacts on the rubber seals
to be used between NASA’s Crew
Exploration Vehicle and ISS and other
spacecraft such as the lunar module
MRS BULLETIN • VOLUME 35 • JANUARY 2010 • www.mrs.org/bulletin
Materials Issues in the Space Environment
Altair. These seals prevent the loss of
cabin air when spacecrafts dock together
and hatches are opened so supplies
and crew can pass between them. Seals
similar to flight seals have been made and
impacted with hypervelocity particles and
then leak tested.27,28 In this way, the particles and kinetic energies that cause seal
failures are determined and the risk of
failure deduced through modeling the
space environment. Figure 7 shows a
rubber seal being hit by an aluminum
projectile and the resulting crater.28 The
article by Grossman et al. in this issue is an
overview of the effects of degradation by
micrometeoroids and space debris.
Thermal Cycling, Solar and (V)UV
Radiation, and Synergistic Effects
Common to all orbits is the degradation
of materials by particle radiation
(protons/electrons), UV light, and thermal cycling. The photovoltaic energy
sources of satellites—solar arrays—are
ideal to discuss these effects, since they
represent one of the most exposed locations on any spacecraft. Photovoltaic
arrays need to have the highest possible
specific power output with regard to mass
and surface area, while showing as little
degradation as possible under intense UV
radiation and ionizing radiation for lifetimes of up to 15 years. In addition, the
arrays need to be durable to thermomechanical stresses induced by temperature fluctuations up to 200°C. In this issue,
Zimmermann presents the current design
solutions to withstand these environmental loads for extended periods in space for
this subassembly common to almost all
space missions and advanced material
solutions for increasing their performance. The challenges in representative
ground testing and extrapolation to endof-life conditions for proton, electron, and
UV radiation of solar modules are
addressed. Thermal cycling is critical,
especially with regard to mechanical
defects that can grow and degrade the
module performance on orbit. These
defects have to be identified and recorded
by specialized detection methods, and
their impact in space has to be treated
statistically.
the polymer. However, even this method
produces surfaces that mechanically
degrade upon exposure at LEO, albeit
more slowly. The article by Kleiman
describes the development of novel
coating materials for protection in the
LEO and GEO environments.
Summary and Outlook
Development of Coating Materials
To protect spacecraft from the harsh
space environment, a wide variety of
durable materials and coatings are utilized, including oxide films, metals,
and semiconductors. The most common
approach to protecting susceptible spacecraft materials from AO erosion is to coat
the material with a thin AO-durable film,
such as SiOx. Unfortunately, microscopic
scratches, dust particles, or other imperfections in the substrate surface can result
in defects in the protective coating. These
coating defects can provide pathways for
AO attack, and undercutting erosion of
the substrate can occur.8 Undercutting erosion can be a serious threat to component
survivability, as shown in Figure 5, where
AO undercutting erosion has severely
degraded the P6 Truss port solar array
wing two-surface aluminized-Kapton
blanket box cover on the ISS (a more
detailed discussion is covered in the
Miller article).29
Recently, protective methods have been
developed that reduce the rate of erosion
of polymers by AO. Both silyation
and ion implantation methods have
been devised to enhance the interfacial
integrity between polymers and coatings,
because polymers can be coated with
paints for thermal control and oxidebased films for protection. The deposited
films of silicon material, produced by both
methods, are subsequently oxidized to
produce stable, graded, and protective
layers that extend into the subsurface of
Space has been explored by manmade
vehicles for more than 50 years. The fact
that most of the missions conducted during this period have met their intended
objective, even with some observed materials degradation, injects a degree of confidence in the suitability of their materials
for the space environment. This “space
heritage,” however, is no replacement for a
systematic study of the material behavior
in space, since it provides no information
on the limit of a particular material or
manufacturing technology. An accurate
knowledge of these boundary conditions,
however, is essential for the success of
increasingly challenging missions, which
push the size, performance, and lifetime of
spacecrafts in well-known orbits and even
more so for missions entering into new
frontiers. In the near future, NASA envisions the building of a moon base, investigations of asteroids, robotic missions, and
perhaps human travel to Mars. Whichever
new place we travel to will have
“weather” that we will need to understand
in detail and develop a predictive knowledge of its environment. As we enter into
the next frontier, corresponding new materials challenges for protection and energy
sources will continually be present.
Therefore, extensive investigations on
the interaction of the various environmental species with spacecraft materials have
been, and continue to be, conducted on
Earth in several countries with active space
programs as well as in space (e.g., Materials
International Space Station Experiment
missions on the International Space
Station). Most important for materials
research is the examination of each effect
systematically as well as to have all effects
present to determine synergistic effects.
Recent investigations have revealed a variety of unusual reaction mechanisms and
possible synergistic effects. The following
articles represent the various critical materials degradation issues in space, particularly in the low Earth orbit environment,
from scientists internationally recognized
in this field.
Acknowledgments
Figure 7. Ultrahigh speed images of a 0.7-mm diameter aluminum projectile at 8.17 km/s hitting
a 5.1-mm wide seal; 2.1 μs between images; the damaged seal did not leak significantly.28
MRS BULLETIN • VOLUME 35 • JANUARY 2010 • www.mrs.org/bulletin
https://doi.org/10.1017/S0883769400100004 Published online by Cambridge University Press
JudithYang thanks the Air Force Office of
Scientific Research and NASA for past support in the area of materials degradation
in the low Earth orbit. The editors thank
15
Materials Issues in the Space Environment
Claus Zimmermann and Gary Pippin for
their critical reading of the introductory
article. We are grateful to all of the authors
for their contributions to this special issue.
References
1. B.A. Banks, K.K. de Groh, E. Baney-Barton,
E.A. Sechkar, P.K. Hunt, A. Willoughby,
M. Bemer, S. Hope, T. Koo, C. Kaminsky, E.
Youngstorm, NASA/ TM-1999 20918 (1999).
2. B.A. Banks, S.K. Rutledge, J.A. Brady,
J.E. Merrow, Proc. Conf. NASA /SDIO Space
Environment Effects on Materials Workshop (June
1993).
3. A.S. Levine, Proc. Conf. LDEF Second Post—
Retrieval Symposium (1993).
4. J.I. Kleiman, Z.A. Iskanderova, Y.I.
Gudimenko, W.D. Morison, R.C. Tennyson,
Can. Aeronaut. Space J. 45, 148 (1999).
5. M.R. Reddy, J. Mater. Sci. 30, 281 (1995).
6. J.A. Townsend, P.A. Hansen, J.A. Dever, K.K.
de Groh, B.A. Banks, L. Wang, C. He, High
Perform. Polym. 81 (1999).
7. J.I. Kleiman, Z. Iskanderova, Eds., Protection
of Materials and Structures from Space Environment ICPMSE-6 (Kluwer Academic Publishers, The Netherlands, 2003).
8. B.A. Banks, K.K. de Groh, S.K. Miller, Proc.
Conf. Materials Research Society, Boston, MA (2004).
9. B.A. Stein, P.R. Young, Proc. Conf. LDEF
Materials Workshop (1992).
10. A.F. Whitaker, Proc. Conf. LDEF Materials
Results for Spacecraft Applications (1994).
11. R.C. Tennyson, G.R. Cool, D.G. Zimcik,
Proc. Conf. LDEF—69 Months in Space (1993).
12. K.K. de Groh, B.A. Banks, J.A. Dever, J.C.
Hodermarsky, High Perform. Polym. 16, 319 (2004).
13. K.K. de Groh, J.R. Gaier, R.L. Hall,
M.P. Espe, D.R. Cato, J.K. Sutter, D.A.
Scheiman, High Perform. Polym. 12, 83 (2000).
14. K.K. de Groh, J.A. Dever, J.K. Sutter, J.R.
Gaier, J.D. Gummow, D.A. Scheiman, C. He,
High Perform. Polym. 13, S401 (2001).
15. J.A. Dever, K.K. de Groh, B.A. Banks, J.A.
Townsend, High Perform. Polym. 11, 123 (1999).
16. J.A. Dever, K.K. de Groh, B.A. Banks, J.A.
Townsend, J.L. Barth, S. Thomson, T. Gregory,
W. Savage, High Perform. Polym. 12, 125 (2000).
17. J.A. Dever, K.K. de Groh, R.K. Messer, M.W.
McClendon, M. Viens, L.L. Wang, J.D.
Gummow, High Perform. Polym. 13, S373 (2001).
18. A.H. Stambler, K.E. Inoshita, L.M. Roberts,
C.E. Barbagallo, K.K. de Groh, B.A. Banks,
Proc. Conf. 9th International Conference, “Protection
of Materials and Structures from Space
Environment,” Toronto, Canada, 19–23 May 2008.
19. L.J. Leger, S.L. Koontz, J.T. Visentine, D.
Hunton, Proc. Conf. LDEF Materials Results for
Spacecraft Application (1994).
20. B.A. Banks, L. Gebauer, C.M. Hill, 1st LDEF
Post-Retrieval Symposium (Kissimmee, FL,
1991).
21. D.G. Gilmore, Spacecraft Thermal Control
Handbook, 2nd edition (The Aerospace Press, El
Segundo, CA, 2002).
22. J.I. Kleiman, R.C. Tennyson, Eds., Protection
of Materials and Structures from the Low Earth
Orbit Space Environment (Kluwer Academic
Publishers, The Netherlands, 1999).
23. J.I. Kleiman, Y. Gudimenko, Z.A.
Iskanderova, R.C. Tennyson, W.D. Morrison,
Judith C. Yang
Kim K. de Groh
Judith C. Yang, Guest
Editor for this issue of
MRS Bulletin, can be
reached by e-mail at
judyyang@pitt.edu.
Yang is a professor
in the Mechanical
Engineering and
Materials Science
Department at The
University of Pittsburgh,
with secondary appoint-
ments in the Chemical
Engineering and Physics
Departments. She earned
her BS degree in physics
at the University of
California, Berkeley, and
her PhD degree at
Cornell University. Yang
was a National Science
Foundation international
post-doctoral research
fellow at Max-Planck-
Institut für
Metallforschung, in
Stuttgart, Germany, and
then a post-doctoral and
visiting lecturer at the
University of Illinois,
Urbana-Champaign. She
joined The University of
Pittsburgh in 1999. Her
research interests include
oxidation, catalysis,
and advanced electron
microscopy—especially
in situ. Yang is the
recipient of the
Chancellor’s
Distinguished Research
Award and National
Aeronautics and Space
Administration (NASA)
and Beyond Petroluem
(BP America) faculty
fellowships.
Kim K. de Groh, Guest
Editor for this issue of
MRS Bulletin, can be
reached at the NASA
16
https://doi.org/10.1017/S0883769400100004 Published online by Cambridge University Press
M.S. McIntyre, R. Davidson, Surf. Interface Anal.
23, 335 (1995).
24. M. Kisa, T. Minton, J.C. Yang, J. Spacecr.
Rockets 43, 431 (2006).
25. L. Li, T. Minton, J.C. Yang, J. Phys. Chem.
C 111, 6763 (2007).
26. L. Li, L. Wang, T.K. Minton, J.C. Yang, Proc.
Conf. Materials Research Society, Symposium NN
(Boston, MA, 2004).
27. H.C. de Groh III, C.A. Gallo, H.K. Nahra,
Proc. Conf.1st AIAA Atmospheric and Space
Environments Conference (AIAA-2009-3524) (San
Antonio, TX, 2009).
28. H.C. de Groh III, B. Steinetz, Proc. Conf. AIAA
Joint Propulsion Conference (AIAA-2009-5249)
(Denver, CO, 2009).
29. B.A. Banks, S.K. Rutledge, B.M. Auer, F.
DiFilippo, Materials Degradation in Low Earth
Orbit (LEO) (TMS Society, 1990), pp. 15–33.
30. R.G. Roble, in The Upper Mesosphere and
Lower Thermosphere: A Review of Experiment and
Theory, Geophysical Monograph 87, pp. 1–21, 1995.
31. K.K. de Groh, B.A. Banks, J.A. Dever,
D.J. Jaworske, S.K. Miller, E.A. Sechkar, 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-SP-08-015E, March 2009, pp. 91–119.
32. K.K. de Groh, B.A. Banks, J.A. Dever, D.J.
Jaworske, S.K. Miller, E.A. Sechkar, S.R. Panko,
“NASA Glenn Research Center’s Materials
International Space Station Experiments
(MISSE 1–7),” NASA TM-2008-215482,
■
December 2008.
Glenn Research Center,
21000 Brookpark Rd.,
MS 309-2, Cleveland,
OH 44135, USA;
tel. 216-433-2297;
fax 216-433-2221;
and e-mail kim.k.degroh
@nasa.gov.
de Groh is a senior
materials research
engineer in the Space
Environment and
Experiments Branch at
the National Aeronautics
and Space
Administration (NASA)
Glenn Research Center,
where she has conducted
research and mentored
students for the past 20
years. She received her
BS (1985) and MS (1987)
degrees in materials
science from Michigan
State University. Kim
specializes in areas
relating to the durability
of spacecraft materials in
the low Earth orbit space
environment. She is the
principal investigator
for 12 International
Space Station external
experiments, and has
studied the degradation
of Teflon insulation
retrieved from the
Hubble Space Telescope
during each of its
servicing missions. Kim
has more than 95
technical publications
and is co-author of the
chapter “Degradation
of Spacecraft Materials”
in the Handbook of
Environmental
Degradation of Materials
(2005). Her honors
include NASA’s Space
Flight Awareness
Honoree Award and
NASA’s Exceptional
Achievement Medal.
She also received the
Society of Automotive
MRS BULLETIN • VOLUME 35 • JANUARY 2010 • www.mrs.org/bulletin
Materials Issues in the Space Environment
Bruce Banks
David L. Edwards
Irina Gouzman
Eitan Grossman
Yugo Kimoto
Engineers’ (SAE
International) J. Cordell
Breed Award for Women
Leaders; the Rotary
National Award for
Space Achievement
(RNASA) Stellar Award,
and, in 2009, Kim was
inducted into the Ohio
Women’s Hall of Fame.
has been the recipient of
six Space Act Awards,
three Federal Laboratory
Consortium Awards, and
the Arthur S. Fleming
Award for selection as
“One of the 10
Outstanding Young
Employees of the Federal
Government.”
Bruce Banks can be
reached at the NASA
Glenn Research Center,
Cleveland, OH 44135,
USA; tel. 216-433-2308;
and e-mail Bruce.A.
Banks@nasa.gov.
Banks is the retired
chief of the ElectroPhysics Branch at the
National Aeronautics
and Space Administration
(NASA) Glenn Research
Center and currently is
a senior physicist at
Alphaport Corporation,
serving NASA. During
his 43 years with NASA,
he has lead research
activities in electric
propulsion technology,
thin-film coatings, surface
texturing processes,
atomic oxygen
interactions, and space
environmental effects.
Banks’ efforts have
resulted in 58 spaceflight
experiments or functional
applications of technology
developed by his research
team. He also has
authored 191 technical
publications and 38
patents. He is the most
patented researcher in the
history of NASA Glenn
Research Center. Banks
David L. Edwards can
be reached at the NASA
Marshall Space Flight
Center, M/S EV44,
Huntsville, AL 35812,
USA; tel. 256-544-4081;
and e-mail David.
L.Edwards@NASA.gov.
Edwards has served as
the branch chief of the
Natural Environments
Branch in the
Engineering Directorate
at the National
Aeronautics and Space
Administration (NASA)
Marshall Space Flight
Center since December
2005. He received a
bachelor’s degree from
the University of North
Alabama in physics in
1985, a master’s degree
from Auburn University
in physics in 1989, and a
PhD degree in materials
engineering from
Auburn University in
1999. Edwards started
his career as an engineer
in the Environmental
Effects Branch in the
Materials and Processes
Laboratory and has
served in a variety of
positions, including
team lead of the Space
Environments Effects
Team, branch chief of
the Environmental
Effects Branch, MSFC
co-resident manager at
the Jet Propulsion
Laboratory for the
Jupiter Icy Moons
Orbiter Program, and
a special assignment as
Center Point of Contact
on the Agency One
NASA Team.
Eitan Grossman can be
reached at Soreq NRC,
Space Environment
Section, Yavne 81800,
Israel; tel. 972-8-9434121;
fax 972-8-9434403;
and e-mail
eitan@soreq.gov.il.
Grossman is the head
of the Materials Group of
the Space Environment
Section at Soreq NRC,
Israel. He received his
PhD degree in materials
engineering from BenGurion University in
Beer-Sheva, Israel. His
research interests include
interaction of space
environment components, such as atomic
oxygen, hypervelocity
debris impact, ultraviolet, and ionizing
radiation, with materials.
Grossman specialized
in polymers and surface
analysis techniques with
emphasis on SPM. His
group at Soreq works
on developing new
materials that will
withstand the harsh
environment of the
low Earth orbit.
J.I. Kleiman can be
reached at the Integrity
Testing Laboratory Inc.,
80 Esna Park Dr., #7-9,
Markham, Ont. L3R 2R7,
Ontario, Canada;
tel. 905-415-2207;
and e-mail jkleiman@
itlinc.com.
Kleiman is the
president and CEO
of the Integrity Testing
Laboratory Inc. which
he formed in 1989. He
earned his BASc degree
from the Institute for
Civil Engineers in
Vilnius, Lithuania, his
MSc degree in solid-state
physics and material
science from the
Technion–Israel Institute
of Technology, and his
PhD degree from the
University of Toronto
Institute for Aerospace
Studies (UTIAS),
Canada. Among
Kleiman’s scientific and
commercial interests, the
topics on development of
protective materials,
coatings, and surface
modification
technologies for space
and aerospace
applications occupy a
very important place.
In 1992, Kleiman
established, and has
since organized, the BiAnnual International
Space Forum on
“Protection of Materials
and Structures from the
Low Earth Orbit Space
Environment” (ICPMSE
series). He has more than
120 scientific papers and
Irina Gouzman can be
reached at Soreq NRC,
Space Environment
Section, Yavne 81800,
Israel; tel. 972-8-9434412;
fax 972-8-9434403; and
e-mail irina@soreq.
gov.il.
Gouzman is a senior
research scientist in the
Materials Group, Space
Environment Section, at
Soreq NRC, Israel. After
receiving her PhD degree
from the Technion–Israel
Institute of Technology’s
Chemistry Department
in 1999, she joined Soreq
NRC, where she works
on the effects of space
environment on
materials durability.
Her research interests
are in the areas of
applied surface science,
nano-diamond films
deposition, and materials
characterization by a
variety of electron
spectroscopy techniques.
In 2006, Gouzman
received an award of
professional excellence
from the Israel
Government
Commissionership.
MRS BULLETIN • VOLUME 35 • JANUARY 2010 • www.mrs.org/bulletin
https://doi.org/10.1017/S0883769400100004 Published online by Cambridge University Press
Yugo Kimoto can be
reached by e-mail at
kimoto.yugo@jaxa.jp.
Kimoto is a researcher
and associate fellow in
the Electronic Devices
and Materials Group of
the Aerospace Research
and Development
Directorate, which is
part of the Japan
Aerospace Exploration
Agency (JAXA).
17
Materials Issues in the Space Environment
J.I. Kleiman
Sharon K.R. Miller
presentations in
reputable international
magazines and
conferences and served
as the chairman of the
American Society for
Metals (ASM
International), Ontario
Chapter.
Consortium Technology
Transfer Award in 2009,
a Research and
Development 100 award
in 2002, and the NASA
Exceptional Service
Medal in 2001.
Sharon K.R. Miller can
be reached at NASA
Glenn Research Center,
21000 Brookpark Rd., MS
309-2, Cleveland, OH
44135, USA;
tel. 216-433-2219;
fax 216-433-2221;
and e-mail sharon.k.
miller@nasa.gov.
Miller is a senior
research engineer at the
National Aeronautics and
Space Administration
(NASA) Glenn Research
Center, where her work
for the last 28 years has
primarily focused on
atomic oxygen
interactions with
materials, thin-film
coatings, and surface
texturing. She holds a
master’s degrees in
materials science and
engineering (1988) and
business administration
(1996) from Case Western
Reserve University, and
a bachelor’s degree in
chemical engineering
(1984) from Cleveland
State University. Miller
has more than 90
publications in these
areas and five patents.
She has won numerous
awards, including a
Federal Laboratory
Timothy K. Minton
can be reached at the
Department of
Chemistry and
Biochemistry, Montana
State University,
Bozeman, MT 59717,
USA; tel. 406-994-5394;
and e-mail tminton@
montana.edu.
Minton is a professor
of chemistry at Montana
State University,
Bozeman. His research
focuses on crossed-beams
and beam-surface
scattering studies of
energy transfer and
reaction dynamics at
hyperthermal energies,
with applications to
space vehicle interactions
with Earth’s outer
atmosphere. Minton also
is a senior editor of the
Journal of Physical
Chemistry.
Masahito Tagawa can be
reached at the Graduate
School of Engineering,
Kobe University,
Rokko-dai 1-1, Nada,
Kobe 657-8501, Japan;
tel. 81-78-803-6126; and
e-mail tagawa@mech.
kobe-u.ac.jp.
Tagawa is an associate
professor at Kobe
University, where he
studies the space
Timothy K. Minton
Masahito Tagawa
Adrian P. Tighe
Marc Van Eesbeek
Ronen Verker
Claus G. Zimmermann
environmental effect on
materials, especially on
atomic oxygen–induced
material erosion in low
Earth orbit. He also is a
guest researcher with the
Japanese Aerospace
Exploration Agency
(JAXA).
the ground-based
simulation of space
environmental effects on
materials (particularly
UV radiation and atomic
oxygen), and laserinduced contamination
effects on space optics.
Tighe also has managed
the development of ESA’s
recent materials
spaceflight experiments.
His previous academic
research centered on the
experimental simulation
of the orbital debris
environment.
Marc Van Eesbeek can be
reached by e-mail at
Marc.van.Eesbeek@
esa.int.
Van Eesbeek is head
of the Materials Space
Evaluation and Radiation
Effects Section at the
European Space
Research and Technology
Centre.
Section, Yavne, 81800,
Israel; tel. 972-8-943-4397;
fax 972-8-943-4403; and
e-mail rverker@soreq.
gov.il.
Verker is a PhD degree
student at Tel-Aviv
University and a
research scientist at
Soreq NRC. His PhD
degree work focuses on
the effects of the low
Earth orbit space
environment on
durability of POSSpolyimide hybrid
nanocomposites. His
research interests
include the effects of
hypervelocity impacts
on space durability of
polymers, polymer
fractography,
development of in-orbit
material degradation
detectors, and combined
effects of the low Earth
environment on
polymers.
Ronen Verker can be
reached at Soreq NRC,
Space Environment
Claus G. Zimmermann
can be reached at
EADS Astrium, 81663
Adrian P. Tighe can be
reached at the European
Space Agency, ESAESTEC/TEC-QEM,
Keplerlaan 1, 2200AG
Noordwijk, The
Netherlands;
tel. 31-71-5654190; and
e-mail Adrian.Tighe@
esa.int.
Tighe is a materials
engineer at the
European Space Agency
(ESA) in the Materials
Space Evaluation and
Radiation Effects
Section. He received his
PhD degree in space
engineering science
from the University of
Southampton, UK, in
2000. His current
specializations include
18
https://doi.org/10.1017/S0883769400100004 Published online by Cambridge University Press
MRS BULLETIN • VOLUME 35 • JANUARY 2010 • www.mrs.org/bulletin
Materials Issues in the Space Environment
Munich, Germany;
tel. 49-89-607-23873;
fax 49-89-607-24202;
and e-mail claus.
zimmermann@astrium.
eads.net.
Zimmermann is an
expert for solar array
technology in the satellite
division of European
Aeronautic Defense and
Space Company (EADS)
Astrium. He received his
diploma in 1997 from the
University of Bonn,
Germany, and his PhD
degree in 2000 from the
University of Augsburg,
Germany—both in
physics. In 1995 and
1996, he was visiting
scientist at the University
of Newcastle, Australia,
involved in surface
analysis by specialized
sputtering techniques.
From 1997 to 2000,
Zimmermann was a
visiting scholar at the
Materials Research
Laboratory of the
University of Illinois,
where he worked on
the fundamentals of
metal nanoparticle
substrate interaction.
He joined the Materials
and Processes Section
of EADS Astrium in
MRS BULLETIN • VOLUME 35 • JANUARY 2010 • www.mrs.org/bulletin
https://doi.org/10.1017/S0883769400100004 Published online by Cambridge University Press
2001, focusing on the
development and
qualification of materials
and manufacturing
processes for the space
environment. Since 2004,
as a member of the
Electronic Subsystem
Group, Zimmermann
has worked on materials
issues related to
photovoltaic power
generation in space.
In this role, he has
pioneered the use of
electroluminescence
as a characterization
tool for space solar cells.
Zimmermann also
has published more
than 10 papers and has
been granted two
patents. In 1999, he
received the MRS
Graduate Student Gold
■
Award.
19
https://doi.org/10.1017/S0883769400100004 Published online by Cambridge University Press