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Synergies between Space Research and
Operations - Examples from the International Space Station
Primary
objectives for the International Space Station (ISS) in support of the Vision for Space Exploration included
conducting research to counteract the harmful effects of space on human
health,
test new space technologies, and learn to operate long-duration space
missions. In pursuit of these objectives,
NASA is
interested in closer cooperation between the ISS operational community,
scientists, and engineers. To develop
the exploration vehicles for missions to the moon and Mars, NASA must
test
materials, foods, and medicines to ensure their performance in the
space
environment. These results will enable
important decisions on the materials to be used for future space
vehicles. Another critical factor for the
success on future
missions beyond Earth’s orbit is the capability for repairs of
equipment. On the ISS, the practice of
crewmembers
performing repairs in microgravity will increase our understanding of
the
repair processes in space; when these capabilities are needed during
future space
exploration missions, we will have the knowledge and experience to
perform
them. The ISS is a unique and
irreplaceable training ground for building the operational knowledge
required
to safely conduct future exploration missions, and the growing links
within the
science, engineering and operations communities are reinforcing the
value of that
training. Current interactions between
the communities that support the ISS have already produced many
synergies that
are significantly accelerating NASA’s advancement towards future
exploration
missions in support of the Vision.
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Comparison of the Atomic Oxygen Erosion
Depth and Cone Height of Various Materials at Hyperthermal Energy
Atomic oxygen readily reacts with most spacecraft
polymer materials exposed to low Earth orbital (LEO) environment.
If the atomic oxygen
arrival comes from a fixed angle of impact, the resulting erosion will
foster
the development of a change in surface morphology as material thickness
decreases.
Hydrocarbon and halopolymer materials, as well as graphite, are easily
oxidized
and textured by directed atomic oxygen in LEO at energies of
~4.5eV. What
has been curious is that the ratio of cone height to erosion depth is
quite
different for different materials. The formation of cones under
fixed direction
atomic oxygen attack may contribute to a reduction in material tensile
strength
in excess of that which would occur if the texture-height to
erosion-depth
ratio was very low because of greater opportunities for crack
initiation.
In an effort to try to understand how material composition affects the
cone
height to erosion depth, an experimental investigation was conducted on
18
different materials exposed to a hyperthermal energy directed atomic
oxygen
source (~90eV). The materials were first salt-sprayed to provide
microscopic
local areas that would be protected from atomic oxygen to allow erosion
depth
measurements to be made by scanning microscopy inspection. The
polymers
were then exposed to atomic oxygen produced by an end Hall ion source
which
was operated on pure atomic oxygen. Samples were exposed to an
atomic oxygen
fluence of 1.0E+20 atoms/cm2. The average erosion depth and
average cone
height was determined using field emission scanning electron microscopy
(FESEM).
The experimental ratio of average cone height to erosion depth will be
compared
to polymer composition and other properties.
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Determination
of Ground-Laboratory to In-Space Effective Atomic Oxygen Fluence for DC
93-500
Silicone 2006
MISSE PEACE Polymers Atomic Oxygen
Erosion
Results
Forty-one different polymer samples, collectively
called
the Polymer Erosion and Contamination Experiment (PEACE) Polymers, have
been
exposed to the low Earth orbit (LEO) environment on the exterior of the
International
Space Station (ISS) for nearly four years as part of Materials
International
Space Station Experiment 2 (MISSE 2). The objective of the PEACE
Polymers
experiment was to determine the atomic oxygen erosion yield of a wide
variety
of polymeric materials after long term exposure to the space
environment.
The polymers range from those commonly used for spacecraft
applications,
such as Teflon FEP, to more recently developed polymers, such as high
temperature
polyimide PMR (polymerization of monomer reactants). Additional
polymers
were included to explore erosion yield dependence upon chemical
composition.
The MISSE PEACE Polymers experiment was flown in MISSE Passive
Experiment
Carrier 2 (PEC 2), tray 1, on the exterior of the ISS Quest Airlock and
was
exposed to atomic oxygen along with solar and charged particle
radiation.
MISSE 2 was successfully retrieved during a space walk on July 30, 2005
during
Discovery’s STS-114 Return to Flight mission. Details on the
specific polymers
flown, flight sample fabrication, pre-flight and post-flight
characterization
techniques, and atomic oxygen fluence calculations are discussed along
with
a summary of the atomic oxygen erosion yield results. The MISSE 2
PEACE
Polymers experiment is unique because it has the widest variety of
polymers
flown in LEO for a long duration and provides extremely valuable
erosion
yield data for spacecraft design purposes.
Ground-to-Space Effective Atomic Oxygen
Fluence
Correlation for DC 93-500 Silicone
The objective of this research was to calibrate the ground-to-space
effective
atomic oxygen fluence for DC 93-500 silicone in a thermal energy
electron
cyclotron resonance (ECR) oxygen plasma facility. A technique has been
developed
at NASA Glenn Research Center to determine the equivalent amount of
atomic
oxygen exposure in an ECR ground-test facility to produce the same
degree
of atomic oxygen damage as in space. The approach used was to compare
changes
in the surface hardness of ground test (ECR)-exposed DC 93-500 silicone
with
DC 93-500 exposed to low Earth orbit (LEO) atomic oxygen as part of a
shuttle
flight experiment. The ground-to-space effective atomic oxygen fluence
correlation
was determined based on the fluence in the ECR source that produced the
same
hardness for the fluence in space. A nanomechanical measurement system
operated
in conjunction with an atomic force microscope (AFM) was used to
determine
the surface hardness of the silicones. Hardness vs contact depth
measurements
were obtained for five ECR-exposed DC 93-500 samples (ECR exposed for
18
to 40 h, corresponding to Kapton effective fluences of 4.2×1020 to
9.4×1020
atoms/cm2, respectively) and for space-exposed DC 93-500 from the
Evaluation
of Oxygen Interactions with Materials III (EOIM III) shuttle flight
experiment,
exposed to LEO atomic oxygen (2.3×1020atoms/cm2). Pristine controls for
the
ECR tests and for the EOIM III flight sample were also evaluated. A
ground-to-space
correlation value was determined based on correlation values for four
contact
depths (150, 200, 250, and 300 nm), which represent the near-surface
depth
data. The results indicate that the Kapton effective atomic oxygen
fluence
in the ECR facility needs to be 2.64 times higher than in LEO to
replicate
equivalent exposure damage in the ground test silicone as occurred in
the
space exposed silicone.
Solar Effects on Tensile and Optical Properties of Hubble
Space
Telescope Silver-Teflon Insulation
A section of the retrieved Hubble Space Telescope (HST) solar array
drive
arm (SADA) multilayer insulation (MLI), which experienced 8.25 years of
space
exposure, was analyzed for environmental durability of the top layer of
silver-Teflon
fluorinated ethylene propylene (Ag-FEP). Because the SADA MLI had
solar
and anti-solar facing surfaces and was exposed to the space environment
for
a long duration, it provided a unique opportunity to study solar
effects
on the environmental degradation of Ag-FEP, a commonly used spacecraft
thermal
control material. Data obtained included tensile properties,
solar absorptance,
surface morphology and chemistry. The solar facing surface was
found to
be extremely embrittled and contained numerous through-thickness
cracks.
Tensile testing indicated that the solar facing surface lost 60% of its
mechanical
strength and 90% of its elasticity while the anti-solar facing surface
had
ductility similar to pristine FEP. The solar absorptance of both
the solar
facing surface (0.155 0.032) and the anti-solar facing surface (0.208
0.012) were found to be greater than pristine Ag-FEP (0.074).
Solar facing
and anti-solar facing surfaces were microscopically textured, and
locations
of isolated contamination were present on the anti-solar surface
resulting
in increased localized texturing. Yet, the overall texture was
significantly
more pronounced on the solar facing surface indicating a synergistic
effect
of combined solar exposure and increased heating with atomic oxygen
erosion.
The results indicate a very strong dependence of degradation,
particularly
embrittlement, upon solar exposure with orbital thermal cycling having
a
significant effect.
Detetermination of Ground-Laboratory to In-Space Effective
Atomic
Oxygen Fluence for DC 93-500 Silicone
The objective of this research was to calibrate
the
ground-to-space effective atomic oxygen fluence for DC 93-500 silicone
in
a thermal energy electron cyclotron resonance (ECR) oxygen plasma
facility.
Silicones, a commonly used spacecraft material, do not chemically erode
with
atomic oxygen attack like other organic materials. Silicones
react with
atomic oxygen and form an oxidized hardened silicate surface layer.
Therefore,
the effective atomic oxygen fluence in a ground test facility should
not
be determined based on mass loss measurements, as they are with organic
polymers,
such as Kapton, a polyimide. A technique has been developed at
the Glenn
Research Center to determine the equivalent amount of atomic oxygen
exposure
in an ECR ground test facility to produce the same degree of atomic
oxygen
damage as in space. The approach used was to compare changes in
the surface
hardness of ground test (ECR) exposed DC 93-500 silicone with DC 93-500
exposed
to low Earth orbit (LEO) atomic oxygen as part of a shuttle flight
experiment.
The ground to in-space effective atomic oxygen fluence correlation was
deter¬mined
based on the fluence in the ECR source that produced the same hardness
for
the fluence in-space. Because microhardness measurements need to
be obtained
on the very surface layer of a rubber substrate (with primarily elastic
deformation)
traditional techniques for microhardness that apply large forces and
indenta¬tions
based on plastic deforma¬tion, could not be used. Therefore, a
nanomechanical
measurement system operated in conjunction with an atomic force
microscope
(AFM) was used to determine the surface hardness of the
silicones. The nanomechanical
system can provide ultra light load indentations and can continuously
measure
force and displacement as an indent is made. Hardness versus
contact depth
measurements were obtained for five ECR exposed DC 93-500 samples (ECR
exposed
for 18 hrs to 40 hrs, corresponding to Kapton effective fluences of 4.2
x
1020 to 9.4 x 1020 atoms/cm2, respectively) and for a space exposed DC
93-500
from the Evaluation of Oxygen Interactions with Materials III (EOIM
III)
shuttle flight experiment, exposed to LEO atomic oxygen for 2.3 x 1020
atoms/cm2.
Pristine controls for the ECR tests and for the EOIM III flight sample
were
also evaluated. A ground-to-space correlation value was
determined based
on correlation values for four contact depths (150, 200, 250 & 300
nm),
which represent the near surface depth data. The results indicate
that the
Kapton effective atomic oxygen fluence in the ECR facility needs to be
2.64
times higher than in LEO to replicate equivalent exposure damage in the
ground
test silicone as occurred in the space exposed silicone.
Fast Three-Dimensional Modeling of Atomic
Oxygen Undercutting of Protected Polymers
Snyder, Aaron and Banks, Bruce, ““Fast Three-Dimensional Modeling of
Atomic Oxygen Undercutting of Protected Polymers”, Journal of
Spacecraft and Rockets, Vol. 41, Number 3, pp. 340-344, May-June 2004
A method is presented to model atomic oxygen erosion of protected
polymers in low Earth orbit. Undercutting of protected polymers
by atomic oxygen can occur due to the presence of scratch, crack or
pin-window defects in the protective coatings. As a means of
providing a better understanding of
undercutting processes, a fast method of modeling atomic-oxygen
undercutting of protected polymers has been developed. Current
simulation methods often rely on computationally expensive ray-tracing
procedures to track the surface-to-surface movement of individual
“atoms”. To reduce the burden of time consuming calculations, the
method introduced in this paper replaces computationally demanding
individual particle simulations by substituting a model that utilizes
both a geometric configuration-factor technique, which collectively
governs the diffuse
transport of atoms between surfaces, and an efficient algorithm, which
rapidly computes the cumulative effects stemming from the series of
atomic
oxygen collisions at the surfaces of an undercut cavity. This new
method
facilitates the systematic study of three-dimensional undercutting by
allowing
rapid simulations to be made over a wide range of erosion parameters.
Space Flight Experiments to Measure Polymer
Erosion and Contamination on Spacecraft
Atomic oxygen erosion and silicone contamination are serious issues
that could damage or destroy spacecraft components after orbiting for
an extended period of time, such as on a space station or satellite. An
experiment, the Polymer Erosion And Contamination Experiment (PEACE)
will be conducted to study the effects of atomic oxygen (AO) erosion
and silicone contamination, and it will provide information and
contribute
to a solution for these problems. PEACE will fly 43 different polymer
materials that will be analyzed for AO erosion effects through two
techniques: mass loss measurement and recession depth measurement.
Pinhole cameras will
provide information about the arrival direction of AO, and silicone
contamination pinhole cameras will identify the source of silicone
contamination on
a spacecraft. All experimental hardware will be passively exposed to AO
for up to two weeks in the actual space environment when it flies in
the
bay of a space shuttle. A second set of the PEACE Polymers is being
exposed to the space environment for erosion yield determination as
part of a second experiment, Materials International Space Station
Experiment (MISSE).
MISSE is a collaboration between several federal agencies and aerospace
companies. During a space walk on August 16, 2001, MISSE was attached
to the outside of the International Space Station (ISS) during an
extravehicular activity (EVA), where it began its exposure to AO for
approximately 1
1/2 years. The PEACE polymers, therefore, will be analyzed after both
short-term and long-term AO exposures for a more complete study of AO
effects.
Techniques for Measuring Low Earth Orbital
Atomic Oxygen Erosion of Polymers
Polymers such as polyimide Kapton® and Teflon® FEP (fluorinated
ethylene propylene) are commonly used spacecraft materials due to their
desirable properties such as flexibility, low density, and in the case
of FEP, a
low solar absorptance and high thermal emittance. Polymers on the
exterior
of spacecraft in the low Earth orbit (LEO) environment are exposed to
energetic atomic oxygen. Atomic oxygen reaction with polymers causes
erosion, which is a threat to spacecraft durability. It is therefore
important to understand the atomic oxygen erosion yield (E, the volume
loss per incident oxygen atom) of polymers being considered in
spacecraft design. The most common technique for determining E is
through mass loss measurements. For limited duration exposure
experiments, such as shuttle experiments, where the atomic oxygen
fluence is often so low that mass loss measurements can not produce
acceptable uncertainties, recession measurements based on atomic force
microscopy analyses can be used. Equally necessary to knowing the mass
loss or recession depth for determining the erosion yield of polymers
is
the knowledge of the atomic oxygen fluence that the polymers were
exposed
to in space. This paper discusses the procedures and relevant issues
for
mass loss and recession depth measurements for passive atomic oxygen
erosion
yield characterization of polymers, along with techniques for active
atomic
oxygen fluence and erosion characterization. One active atomic oxygen
erosion
technique discussed is a new technique based on optical measurements.
Details
including the use of both semi-transparent and opaque polymers for
active
erosion measurement are reviewed.
Issues and Consequences of Atomic Oxygen
Undercutting of Protected Polymers in Low Earth Orbit
Hydrocarbon based polymers that are exposed to atomic oxygen in
low Earth orbit are slowly oxidized which results in recession of their
surface. Atomic oxygen protective coatings have been developed which
are both durable to atomic oxygen and effective in protecting
underlying
polymers. However, scratches, pin window defects, polymer surface
roughness
and protective coating layer configuration can result in erosion and
potential failure of protected thin polymer films even though the
coatings
are themselves atomic oxygen durable. This paper will present issues
that cause protective coatings to become ineffective in some cases yet
effective in others dues to the details of their specific application.
Observed in-space examples of failed and successfully protected
materials
using identical protective thin films will be discussed and analyzed.
Proposed approaches to prevent the failures that have been observed
will
also be presented.
Mechanical Properties of Teflon® FEP
Retrieved From the Hubble Space Telescope
Teflon® FEP (fluorinated ethylene propylene) surfaces on the Hubble
Space Telescope (HST) have experienced significant degradation in
mechanical properties during nearly ten years of exposure in the low
Earth orbit (LEO) environment. This paper describes results of
mechanical properties testing of Teflon® FEP materials exposed on HST
for 9.7 years between launch and the third servicing mission (SM3A) and
for 2.8 years between the second servicing mission (SM2) and SM3A.
Results of tensile testing, bend testing and microscopic examination of
crack morphology are described. Effects of post-retrieval heating and
air vs. vacuum storage on the mechanical properties of FEP surfaces are
described as they significantly affect interpretation of results
regarding the durability of FEP on HST. This paper provides comparisons
of the properties of FEP surfaces retrieved during SM3A to previously
reported results for FEP materials retrieved during the first servicing
mission (SM1) and SM2. Environmental exposure conditions for the HST
exposed materials are
also described.
A Sensitive Technique Using Atomic Force
Microscopy to Measure the Low Earth Orbit Atomic Oxygen Erosion of
Polymers
Polymers such as polyimide Kapton and Teflon FEP (fluorinated ethylene
propylene) are commonly used spacecraft materials due to their
desirable properties such as flexibility, low density, and in the case
of FEP low solar absorptance and high thermal emittance. Polymers on
the exterior of spacecraft in the low Earth orbit (LEO) environment are
exposed to energetic atomic oxygen. Atomic oxygen erosion of polymers
occurs in LEO and is a threat
to spacecraft durability. For example, depths of more than 0.0127 cm
thickness
of Kapton and Mylar were eroded away after 5.8 years in LEO on the Long
Duration
Exposure Facility (LDEF). It is therefore important to understand the
atomic
oxygen erosion yield (E, the volume loss per incident oxygen atom) of
polymers
being considered in spacecraft design. Because long-term space exposure
data is rare and very costly, short-term exposures such as on the
shuttle
are often relied upon for atomic oxygen erosion determination. The most
common
technique for determining E is through mass loss measurements. For
limited
duration exposure experiments, such as shuttle experiments, the atomic
oxygen
fluence is often so small that mass loss measurements can not produce
acceptable uncertainties. Therefore, a recession measurement technique
has been developed using selective protection of polymer samples,
combined with post-flight atomic force microscopy (AFM) analysis, to
obtain accurate erosion yields of polymers exposed to low atomic oxygen
fluences. This paper discusses
the procedures used for this recession depth technique along with
relevant
characterization issues. In particular, a polymer is salt-sprayed prior
to flight, then the salt is washed off post-flight and AFM is used to
determine
the erosion depth from the protected plateau. A small sample was
salt-sprayed
for AFM erosion depth analysis and flown as part of the Limited
Duration
Candidate Exposure (LDCE-4,-5) shuttle flight experiment on STS-51.
This
sample was used to study issues such as use of contact versus
non-contact
mode imaging for determining recession depth measurements. Error
analyses
were conducted and the percent probable error in the erosion yield when
obtained by the mass loss and recession depth techniques has been
compared. The recession depth technique is planned to be used to
determine the erosion yield of 42 different polymers in the shuttle
flight experiment PEACE
(Polymer Erosion And Contamination Experiment) planned to fly in 2002
or 2003.
MISSE PEACE Polymers: an International
Space
Station Environmental Exposure Experiment
Forty-one different polymers are being exposed to the low Earth
orbit (LEO) environment on the exterior of the International Space
Station
(ISS) for one year as part of MISSE (Materials International Space
Station
Experiment). MISSE is a materials flight experiment sponsored by the
Air Force Research Lab/Materials Lab and the National Aeronautics and
Space Administration (NASA). A second set of the same polymers is
planned
to be flown as part of PEACE (Polymer Erosion And Contamination
Experiment),
a short duration shuttle flight experiment, and therefore these
forty-one
polymers on ISS are collectively called the MISSE PEACE Polymers. The
objective
of the MISSE PEACE Polymers experiment is to accurately determine the
atomic oxygen (AO) erosion yield of a wide variety of polymeric
materials.
The polymers range from those commonly used for spacecraft
applications,
such as Teflon® FEP, to more recently developed polymers, such as high
temperature polyimide PMR (polymerization of monomer reactants).
Additional
polymers were included to explore erosion yield dependence upon
chemical
composition. Details on the specific polymers being flown, flight
sample
fabrication, and pre-flight characterization techniques will be
discussed.
The MISSE PEACE Polymers experiment was placed on the exterior of ISS
during
a spacewalk on August 16, 2001 and is planned to be retrieved in the
fall
of 2002. The erosion yield data obtained from this experiment will be
compared
with data from the future short duration experiment PEACE and with
predicted
results from models developed by a Canadian group that predicts the AO
erosion yield of organic materials based on their chemical structure.
Having
flight data, and comparing flight data with the predictive model
results,
will be valuable for spacecraft design purposes.
Exposure of Polymer Film Thermal Control
Materials on the Materials International Space Station Experiment
(MISSE)
Seventy-nine samples of polymer film thermal control (PFTC) materials
have been provided by the National Aeronautics and Space Administration
(NASA) Glenn Research Center (GRC) for exposure to the low Earth orbit
environment on the exterior of the International Space Station (ISS) as
part of the Materials International Space Station Experiment (MISSE).
MISSE
is a materials flight experiment sponsored by the Air Force Research
Lab/Materials
Lab and NASA. The paper will describe background, objectives, and
configurations for the GRC PFTC samples for MISSE. These samples
include polyimides,
fluorinated polyimides, and Teflon® fluorinated ethylene propylene
(FEP)
with and without second-surface metallizing layers and/or surface
coatings.
Also included are polyphenylene benzobisoxazole (PBO) and a polyarylene
ether benzimidazole (TOR-LMTM). On August 16, 2001, astronauts
installed
passive experiment carriers (PECs) on the exterior of the ISS in which
were located twenty-eight of the GRC PFTC samples for 1-year space
exposure.
MISSE PECs for 3-year exposure, which will contain fifty-one GRC PFTC
samples,
will be installed on the ISS at a later date. Once returned from the
ISS,
MISSE GRC PFTC samples will be examined for changes in optical and
mechanical
properties and atomic oxygen (AO) erosion. Additional sapphire witness
samples located on the AO exposed trays will be examined for deposition
of contaminations.
Modeling of Transmittance Degradation
Caused
by Optical Surface Contamination by Atomic Oxygen Reaction With
Adsorbed
Silicones
A numerical procedure is presented to calculate transmittance
degradation caused by contaminant films on spacecraft surfaces produced
though the interaction of orbital atomic oxygen (AO) with volatile
silicones and
hydrocarbons from spacecraft component. In the model, contaminant
accretion
is dependent on the adsorption of species, depletion reactions due to
gas-surface collisions, desorption, and surface reactions between AO
and silicone producing SiOx (where x is near 2). A detailed description
of the procedure used to calculate the constituents of the contaminant
layer is presented, including the equations that govern the evolution
of fractional coverage by specie type. As an illustrative example of
film
growth, calculation results using a prototype code that calculates the
evolution
of surface coverage by specie type is presented and discussed. An
example
of the transmittance degradation caused by surface interaction of AO
with
deposited contaminant is presented for the case of exponentially
decaying
contaminant flux. These examples are performed using hypothetical
values
for the process parameters.
Hubble Space Telescope Third Servicing
Mission Retrieved Metallized Teflon FEP Analysis
Following the third servicing mission (SM3A in December ’99) to
the Hubble Space Telescope, analysis was performed on the two returned
panels of multilayer insulation (MLI) as well as two patches. The MLI
panels had been in space since the telescope was launched in April ’90
(9.7 years), while the patches were installed during the second
servicing
mission in February ’97 (2.8 years). This paper provides an overview of
the tests performed on the returned metallized Teflon FEP along with a
summary of results. Testing including determination of mechanical and
optical
properties, crystallinity and fractography. Because of the amount of
material
retrieved and the nominal environmental exposures of the retrieved
materials,
these analyses resulted in a fairly complete understanding of the
degradation
process affecting the materials on the telescope. Test results from
SM3A
materials showed significantly better mechanical strength than second
servicing
mission (SM2) samples.
Evaluation of Optical Degradation of
Aluminized Sapphire Flown On Russian Mir Space Station
Materials degradation in Low Earth Orbit (LEO) is of continued concern,
particularly for long duration space applications such as the
International Space Station (ISS). The Passive Optical Sample Assembly
(POSA) experiment flown on the exterior of the Mir as a risk mitigation
experiment for the ISS was designed to better understand the potential
contamination threats that may be present in the vicinity of the
spacecraft. Deterioration in the optical performance of candidate space
power materials due to the LEO environment and contamination must be
evaluated in order to propose measures to mitigate such deterioration.
The optical properties of optically thick aluminum (>800Ĺ)
on sapphire substrates were evaluated prior to and after the POSA
flight. Total, diffuse, and specular reflectance measurements were
acquired to
quantify the degradation in reflectivity due to contamination. Variable
angle spectroscopic ellipsometry (VASE) covering a spectral range from
0.25 microns to 14 microns was used to identify the type and thickness
of the contamination. This paper summarizes the results of pre- and
post-flight
optical measurements, identifies the mechanisms responsible for optical
properties deterioration, and suggests improvements for the durability
of materials in future missions.
Thermal Contributions to the Degradation of
Teflon® FEP on the Hubble Space Telescope
Metallized Teflon® fluorinated ethylene propylene (FEP) thermal
control material on the Hubble Space Telescope (HST) is degrading in
the
space environment. Teflon® FEP insulation was retrieved during
servicing
missions, which occurred in 1993, 1997 and 1999. During the second
servicing
mission (SM2), the 5 mil aluminized-FEP (Al-FEP) outer layer of
multilayer
insulation (MLI) covering the telescope was found to be cracked in many
locations around the telescope. Teflon® FEP retrieved during SM2 was
more
embrittled than FEP retrieved 2.8 years later from a different
location,
during the third servicing mission (SM3A). Studies have been conducted
to understand the degradation of FEP on HST, and the difference in the
degree
of degradation of FEP from each of the servicing missions. The
retrieved
SM2 material experienced a higher temperature extreme during thermal
cycling
(200°C) than first servicing mission (SM1) and SM3A materials (upper
temperature
of 50°C), therefore an investigation on the effects of heating FEP was
also
conducted. Samples of pristine FEP and SM1, SM2, and SM3A retrieved FEP
were heated to 200°C and evaluated for changes in properties. Heating
at
130°C was also conducted because FEP bi-stem thermal shields are
expected
to cycle to a maximum temperature of 130°C on-orbit. Tensile, density,
x-ray
diffraction (XRD) crystallinity and differential scanning calorimetry
(DSC)
data were evaluated. It was found that heating pristine FEP caused an
increase
in the density and practically no change in tensile properties.
However,
when as-retrieved space samples were heated, the density increased and
tensile
properties decreased. Upon heating, all samples experienced an increase
in crystallinity, with larger increases in the space exposed FEP. These
results indicate that irradiation of FEP in space causes chain
scission,
resulting in embrittlement, and that excessive heating allows increased
mobility of space-environment-induced scissioned chains. Thermal
exposure
was therefore found to have a major impact on the extent of
embrittlement
of FEP on HST.
Prediction and Measurement of the Atomic
Oxygen Erosion Yield of Polymers in Low Earth Orbital Flight
Recently developed models of the erosion of polymeric materials
by AO in low Earth orbit (LEO) have been used for predictive evaluation
of the erosion resistance in LEO for a representative, comprehensive
set of polymers. The established correlations of erosion yield values
with the chemical composition and structure of hydrocarbon polymers,
and with their flammability have been used for predictive evaluation of
the behavior of those materials in LEO. Among these materials, a
variety of aromatic and
aliphatic hydrocarbon polymers, including homopolymers, copolymers, and
terpolymers, have been considered. Predictive estimates have also been
given
for linear-chain fluoro- and fluoro-chloropolymers. With different
degrees
of fluorination, using a recently modified version of the predictive
model,
and the results were in good agreement with the flight data, that exist
to
date. Altogether, predictive evaluations have been performed for more
than
40 polymers, including a few recommended materials for which the lower
and
higher extremes in erosion yield in LEO can be expected, based on their
chemical
composition and structure.
For almost half of the selected materials, there is no date
from neither space nor ground-based experimental testing. For the rest,
the data was collected mostly from the Long Duration Exposure Facility
(LDEF) and several other flight experiments. The predicted erosion
yield values ReLEO were found to be, mostly, in good agreement with the
flight data
for materials, already tested in LEO. A reasonable agreement between
the
two mentioned above predictive correlations, i.e., the one related to
the
chemical composition and structure of the materials, and the one
related
to the material’s flammability was found for ReLEO (pred) for the
majority
of untested materials.
A low Earth orbital space experiment entitled “Polymers
Erosion
and Contamination Experiment,” has been designed and is planned to
allow
measurement of the atomic oxygen erosion yield of a set of 40 different
polymeric materials, whose erosion yields were predicted as described
above.
This will allow direct comparison between predicted and measured
in-space
atomic oxygen erosion yield. The experiment is a Get-Away-Special (GAS
can)
experiment to be conducted in the Shuttle bay that allows atomic oxygen
to impinge on two sets of the 40 types of polymers. One set of polymer
samples
will be analyzed later, using weigh loss to measure atomic oxygen
erosion
yields, and the other set will be evaluated using erosion depth to
measure
atomic oxygen erosion yield. Erosion depth will be measured by means of
salt
or mica flake particles on the polymer surfaces to act on local
protective
coatings, which will result in the production of step height changes
that
are measured by atomic force microscopy. Using this latter technique,
erosion
yield measurements with uncertainties of ~3% can be achieved for
typical
polymers with atomic oxygen fluences of ~5 x 1019 atoms/cm2.
Issues and Effects of Atomic Oxygen
Interactions With Silicone Contamination on Spacecraft in Low Earth
Orbit
The continued presence and use of silicones on spacecraft in low Earth
orbit (LEO) has been found to cause the deposition of contaminant films
on surfaces which are also exposed to atomic oxygen. The composition
and optical properties of the resulting SiOx-based (where x is near 2)
contaminant films may be dependent upon the relative rates of arrival
of atomic oxygen, silicone contaminant, and hydrocarbons. This paper
presents results of in-space silicone contamination tests, ground
laboratory simulation tests, and analytical modeling to identify
controlling processes that affect contaminant characteristics.
Insights into the Damage Mechanism of
Teflon® FEP from the Hubble Space Telescope
Metallized Teflon® FEP (fluorinated ethylene propylene) thermal
control material on the Hubble Space Telescope (HST) has been found to
be degrading in the space environment. Teflon® FEP thermal control
blankets
(space-facing FEP) retrieved during the first servicing mission (SM1)
were found to be embrittled on solar facing surfaces and contained
microscopic
cracks. During the second servicing mission (SM2) astronauts noticed
that
the FEP outer layer of the multi-layer insulation (MLI) covering the
telescope was cracked in many locations around the telescope. Large
cracks were observed on the light shield, forward shell and equipment
bays. A tightly curled piece of cracked FEP from the light shield was
retrieved during SM2 and was severely embrittled, as witnessed by
ground testing. A Failure Review Board (FRB) was organized to determine
the mechanism causing the MLI degradation. Density, x-ray crystallinity
and solid state nuclear magnetic resonance (NMR) analyses of FEP
retrieved during SM1 were inconsistent with results of FEP retrieved
during SM2. Because the retrieved SM2 material curled while in space,
it experienced a higher temperature extreme during thermal cycling,
estimated at 200°C, than the SM1 material, estimated at 50°C. An
investigation on the effects of heating pristine FEP and FEP retrieved
from
the HST was therefore conducted. Samples of pristine, SM1, and SM2 FEP
were heated to 200°C and evaluated for changes in density and
morphology. Elevated temperature exposure was found to have a major
impact on the density of the retrieved materials. Characterization of
polymer morphology of as-received and heated FEP by NMR provided
results that were consistent with the density results. Differential
scanning calorimetry (DSC) was conducted on pristine, SM1 and SM2 FEP.
DSC results provided evidence of chain scission and increased
crystallinity in the space exposed FEP, which supported the density and
NMR
results. Samples exposed to simulated solar flare x-rays, thermal
cycling
and long-term thermal exposure provided information on environmental
contributions
to degradation. These findings have provided insight into the damage
mechanisms
of FEP in the space environment.
Environmental Exposure Conditions for
Teflon
FEP on the Hubble Space Telescope
The outer layer of Teflon® fluorinated ethylene propylene (FEP)
multi-layer insulation (MLI) on the Hubble Space Telescope (HST) was
observed to be significantly cracked at the time of the Second HST
Servicing
Mission (SM2), 6.8 years after HST was launched into low Earth orbit
(LEO).
Comparatively minor embrittlement and cracking were also observed in
FEP
materials retrieved from solar-facing surfaces on HST at the time of
the
First Servicing Mission (3.6 years exposure). After SM2, a Failure
Review
Board was convened to address the problem of degradation of MLI on HST.
In order for this board to determine possible degradation mechanisms,
it
was necessary to consider all environmental constituents to which the
FEP
MLI surfaces were exposed. Based on measurements and various models,
environmental
exposure conditions for FEP surfaces on HST were estimated including:
number
and temperature ranges of thermal cycles; equivalent sun hours; fluence
and absorbed radiation dose of x-rays, trapped protons and electrons,
and
plasma electrons and protons; and atomic oxygen (AO) fluence. This
paper
presents the environmental exposure conditions for FEP on the Hubble
Space
Telescope, briefly describing the possible roles of the environmental
factors
in the observed FEP embrittlement and providing references to the
published
works which describe in detail testing and analysis related to FEP
degradation
on HST.
Analyses of Contaminated Solar Array
Handrail Samples Retrieved from Mir
In January 1998 during the STS-89 mission, an eight section Russian
solar array panel was retrieved after more than ten years exposure to
the orbital space environment on the Russian space station Mir. Two
darkened handrail samples from the Russian solar array have been
evaluated for
contamination; a section of a white paint covered rigid handrail and a
section of woven fabric over-wrapped around a flexible handhold. The
handrail
samples were evaluated using optical microscopy (OM), field emission
scanning
electron microscopy (FESEM) and energy dispersive spectroscopy (EDS).
Optical
properties were also obtained. Microscopy has shown the discolored
areas
to have thick layers of contaminant that has crazed and spalled off in
regions. Energy dispersive spectroscopy revealed that the brown
contaminant
is composed of oxidized silicon. No silicon was present on the
unexposed
fabric over-wrap, and very small amounts were present in the white
paint.
Therefore, the contaminant layer on both samples is attributed to
silicone
contamination from other spacecraft materials that have been oxidized
by
atomic oxygen while in orbit. A significant source of the silicone
contamination
appears to be from the solar array itself.
Consequences of Atomic Oxygen Interaction
with Silicone and Silicone Contamination on Surfaces in Low Earth Orbit
The exposure of silicones to atomic oxygen in low Earth orbit causes
oxidation of the surface, resulting in conversion of silicone to
silica. This chemical conversion increases the elastic modulus to the
surface and initiates the development of a tensile strain. Ultimately,
with sufficient exposure, tensile strain leads to cracking of the
surface enabling the underlying unexposed silicone to be converted to
silica resulting in additional depth and extent of cracking. The use of
silicone coatings for the protection of materials from atomic oxygen
attack is limited because of the eventual exposure of underlying
unprotected polymeric material due to deep tensile stress cracking of
the oxidized silicone. The use of moderate to high volatility silicones
in low Earth orbit has resulted in a silicone contamination arrival at
surfaces which are simultaneously being bombarded with atomic oxygen,
thus leading to conversion of the silicone contaminant to silica. As a
result of these processes, a gradual accumulation of contamination
occurs leading to deposits, which at times have been up to several
microns thick (as in the case of a Mir solar array after 10 years in
space). The contamination species typically consist of silicon, oxygen,
and carbon, which in the synergistic environment of atomic oxygen and
UV radiation leads to increased solar absorptance and reduced solar
transmittance. A comparison of the results of atomic oxygen interaction
with silicones and silicone contamination will be presented
based on the LDEF, EOIM-III, Offeq-3 spacecraft and Mir solar array
in-space
results. The design of a contamination pin-hole camera space
experiment,
which uses atomic oxygen to produce an image of the sources of silicone
contamination, will also be presented.
A Space Experiment to Measure the Atomic
Oxygen Erosion of Polymers and Demonstrate a Technique to Identify
Sources of Silicone Contamination
A low Earth orbital space experiment entitled, "Polymers Erosion and
Contamination Experiment" (PEACE) has been designed as a Get-Away
Special (GAS Can) experiment to be accommodated as a Shuttle in-bay
environmental exposure experiment. The first objective is to measure
the atomic oxygen erosion yields of ~40 different polymeric materials
by mass loss and erosion measurements using atomic force microscopy.
The second objective is to evaluate the capability of identifying
sources of silicone contamination through the use of a pin-hole
contamination camera, which utilizes environmental atomic oxygen to
produce a contaminant source image on an optical substrate.
Durability of Intercalated Graphite in
Epoxy Composites in Low Earth Orbit
The electrical conductivity of graphite epoxy composites can be
substantially increased by intercalating (inserting guest atoms or
molecules
between the graphene planes) the graphite fibers before composite
formation.
The resulting high strength, low density, electrically conducting
composites have been proposed for EMI shielding in spacecraft.
Questions have been
raised, however, about their durability in the space environment,
especially
with respect to outgassing of the intercalates, which are corrosive
species
such as bromine. To answer those concerns, six samples of bromine
intercalated graphite epoxy composites were included in the third
Evaluation of Oxygen Interaction with Materials (EOIM-3) experiment
flown on the Space Shuttle Discovery (STS-46). Changes in electrical
conductivity, optical reflectance, surface texture, and mass loss for
SiO2 protected and unprotected samples were measured after
being exposed to the LEO environment for 42
hours. SiO2 protected samples showed no degradation,
verifying
conventional protection strategies are applicable to bromine
intercalated
composites. The unprotected samples showed that bromine intercalation
does not alter the degradation of graphite-epoxy composites. No bromine
was detected to have been released by the fibers allaying fears that
outgassing
could be disruptive to the sensitive electronics the EMI shield is
meant
to protect.
Durability of Intercalated Graphite in
Epoxy Composites in Low Earth Orbit
The electrical conductivity of graphite epoxy composites can be
substantially increased by intercalating (inserting guest atoms or
molecules between
the graphene planes) the graphite fibers before composite formation.
The
resulting high strength, low density, electrically conducting
composites
have been proposed for EMI shielding in spacecraft. Questions have been
raised, however, about their durability in the space environment,
especially
with respect to outgassing of the intercalates, which are corrosive
species
such as bromine. To answer those concerns, six samples of bromine
intercalated
graphite epoxy composites were included in the third Evaluation of
Oxygen
Interaction with Materials (EOIM-3) experiment flown on the Space
Shuttle
Discovery (STS-46). Changes in electrical conductivity, optical
reflectance,
surface texture, and mass loss for SiO2 protected and unprotected
samples
were measured after being exposed to the LEO environment for 42 hours.
SiO2
protected samples showed no degradation, verifying conventional
protection
strategies are applicable to bromine intercalated composites. The
unprotected
samples showed that bromine intercalation does not alter the
degradation
of graphite-epoxy composites. No bromine was detected to have been
released
by the fibers allaying fears that outgassing could be disruptive to the
sensitive electronics the EMI shield is meant to protect.
Environmental Durability Issues for Solar
Power Systems in Low Earth Orbit
Space solar power systems for use in low Earth orbit (LEO) environment
experience a variety of harsh environmental conditions. Materials used
for solar power generation in LEO need to be durable to environmental
threats such as atomic oxygen, ultraviolet (UV) radiation, thermal
cycling, and micrometeoroid and debris impact. Another threat to LEO
solar power performance is due to contamination from other spacecraft
components. This paper gives an overview of these LEO environmental
issues as they relate to space solar power system materials. Issues
addressed include atomic oxygen erosion of organic materials, atomic
oxygen undercutting of protective coatings, UV darkening of ceramics,
UV embrittlement of Teflon, effects of thermal cycling on organic
composites, and contamination due to silicone and organic materials.
Specific examples of samples from the Long Duration Exposure Facility
(LDEF) and materials returned from the first servicing mission of the
Hubble Space Telescope (HST) are presented. Issues concerning ground
laboratory
facilities which simulate the LEO environment are discussed along with
ground-to-space
correlation issues.
A Comparison of Atomic Oxygen Erosion
Yields of Carbon and Selected Polymers Exposed in Ground Based
Facilities and in Low Earth Orbit
A comparison of the relative erosion yields (volume of material
removed per oxygen atom arriving) for FEP Teflon, polyethylene, and
pyrolytic
graphite with respect to Kapton HN was performed in an atomic oxygen
directed
beam system, in a plasma asher, and in space on the EOIM-III
(Evaluation
of Oxygen Interaction with Materials-III) flight experiment. This
comparison was performed to determine the sensitivity of material
reaction to atomic oxygen flux, atomic oxygen fluence, and vacuum
ultraviolet radiation for enabling accurate estimates of durability in
ground based facilities. The relative erosion yield of pyrolytic
graphite was found not to be sensitive to these factors that for FEP
was sensitive slightly to fluence and possibly ions, and that for
polyethylene was found to be partially VUV and flux sensitive but more
sensitive to an unknown factor. Results indicate that the ability to
use these facilities for material relative durability prediction is
great as long as the sensitivity of particular materials to conditions
such as VUV, and atomic oxygen flux and fluence are taken into account.
When testing materials of a particular group such as Teflon, it may be
best
to use a witness sample made of a similar material that has some
available
space data on it. This would enable one to predict an equivalent
exposure
in the ground based facility.
Evaluation of Space Power Materials Flown
on the Passive Optical Sample Assembly
Evaluating the performance of materials on the exterior of spacecraft
id of continuing interest, particularly in anticipation of those
applications that will require a long duration in low Earth orbit. The
Passive Optical Sample Assembly (POSA) experiment flown on the exterior
of Mir as a risk mitigation experiment for the International Space
Station was designed
to better understand the interaction of materials with the low Earth
orbit environment and to better understand the potential contamination
threats that may be present in the vicinity of spacecraft.
Deterioration in the
optical performance of candidate space power materials due to the low
Earth
orbit environment, the contamination environment, or both, must be
evaluated
in order to propose measures to mitigate such deterioration. The
thirty-two samples of space power materials studied here include solar
array blanket materials such as polyimide Kapton H and SiOx coated
polyimide Kapton H, front surface aluminized sapphire, solar dynamic
concentrator materials
such as silver on spin coated polyimide and aluminum on spin coated
polyimide, CV1144 silicone, and the thermal control paint Z-93-P. The
physical and
optical properties that were evaluated prior to and after the POSA
flight
include mass, total, diffuse, and specular reflectance, solar
absorptance,
and infrared emittance. Additional post flight evaluation included
scanning
electron microscopy to observe surface features caused by the low Earth
orbit environment and the contamination environment, and variable angle
spectroscopic
ellipsometry to identify contaminant type and thickness. This paper
summarizes
the results of pre- and post-flight measurements, identifies the
mechanisms
responsible for optical properties deterioration, and suggests
improvements
for the durability of materials in future missions.
Atomic Oxygen Undercutting of Long
Duration
Exposure facility Aluminized Kapton Multilayer Insulation
Atomic oxygen undercutting is a potential threat to vulnerable
spacecraft materials which have atomic oxygen protective coatings. Such
undercutting is due to the atomic oxygen attack of oxidized materials
at microscopic defects in the protective coatings. These defects occur
during fabrication and handling, or from micrometeoroid and debris
bombardment in space.
An aluminized-polyimide Kapton multi-layer insulation sample that was
located on the leading edge of the Long Duration Exposure Facility has
been used to study low Earth orbit atomic oxygen undercutting. Cracks
in the aluminized coating located around vent holes provided excellent
defect sites for the evaluation of atomic oxygen undercutting. The
experimentally
observed undercut profiles were compared to predictions from Monte
Carlo
models for normal incident space ram atomic oxygen attack. The shape of
the undercut profile was found to vary with crack width, which is
proportional
to the number of oxygen atoms entering the crack. The resulting
profiles
of atomic oxygen undercutting which occurred on the aluminized-Kapton
sample indicated wide undercut cavities in spite of the fixed ram
orientation.
Potential causes of the observed undercutting are presented.
Implications
of the undercutting profiles relevant to Space Station Freedom are also
discussed.
Atomic Oxygen Interaction at Defect Sites
in Protective Coatings on Polymers Flown on LDEF
Although the Long Duration Exposure Facility (LDEF) had exposed
materials with a fixed orientation relative to the ambient
low-Earth-orbital
environment, arrival of atomic oxygen is angularly distributed as a
result
of the atomic oxygen's high temperature Maxwellian velocity
distribution
and the LDEF's orbital inclination. Thus, atomic oxygen entering
defects
in protective coatings on polymeric surfaces can cause wider undercut
cavities than the size of the defect in the protective coating. Because
only a small fraction of atomic oxygen reacts upon first impact with
most
polymeric materials, secondary reactions with lower energy thermally
accommodated
atomic oxygen can occur. The secondary reactions of scattered and/or
thermally
accommodated atomic oxygen also contribute to widening the undercut
cavity
beneath the protective coating defect. As the undercut cavity enlarges,
exposing more polymer, the probability of atomic oxygen reacting with
underlying
polymeric material increases because of multiple opportunities for
reaction.
Thus, the effective atomic oxygen erosion yield for atoms entering
defects
above that of the unprotected material. Based on the results of
analytical
modeling and computational modeling, aluminized Kapton multilayer
insulation
exposed to atomic oxygen on row 9 lost the entire externally exposed
player
of polyimide Kapton, yet based on the results of this investigation,
the
bottom surface aluminum film must have remained in place, but crazed.
Atomic
oxygen undercutting at defect sites in protective coatings on graphite
epoxy composites indicates that between 40 to 100 percent of the atomic
oxygen thermally accommodates upon impact, and that the reaction
probability
of thermally accommodated atomic oxygen may range from 1.1x10-6
to 2.1x10-3, depending upon the degree of thermal
accommodation
upon each impact.
Atomic Oxygen Interactions with Protected
Organic Materials on the Long Duration Exposure Facility
The Long Duration Exposure Facility (LDEF) has provided an excellent
opportunity to understand the nature of directed atomic oxygen
interactions with protected polymers and composites. Although there
were relatively few samples of materials with protective coatings on
their external surfaces on LDEF which were exposed to a high atomic
oxygen fluence, analysis of such samples has enabled an examination of
the shape of atomic oxygen undercut cavities at defect sites in the
protective coatings. Samples of front-surface aluminized (Kapton®)
polyimide were inspected by scanning electron microscopy to identify
and measure crack defects in the aluminum protective coatings. After
chemical removal of the aluminum coating, measurements were also made
of the width of the oxidized undercut cavities below the crack defects.
The LDEF flight undercut cavity geometries were then compared to the
Monte Carlo computational model undercut cavity predictions. The
comparison of the LDEF results and computational modeling indicates
agreement in specific undercut cavity geometries for atomic oxygen
reaction probabilities dependant upon the 0.68 to 3.0 power if the
energy. However, no single energy dependency was adequate to replicate
flight results over a variety of aluminum crack widths.
Prediction of In-Space Durability of
Protected Polymers Based on Ground Laboratory Thermal Energy Atomic
Oxygen
The probability of atomic oxygen reacting with polymeric materials is
orders of magnitude lower at thermal energies (<0.1 eV) than at
orbital impact energies (4.5 eV). As a result, absolute atomic oxygen
fluxes at thermal energies must be orders of magnitude higher than
orbital energy
fluxes, to produce the same effective fluxes (or same oxidation rates)
for
polymers. These differences can cause highly pessimistic durability
predictions
for protected polymers, and polymers which develop protective metal
oxide
surfaces as a result of oxidation if one does not make suitable
calibrations.
A comparison was conducted of undercut cavities below defect sites in
protected
polyimide Kapton samples flown on the Long Duration Exposure Facility
(LDEF) with similar samples exposed in thermal energy oxygen plasma.
The
results of this comparison were used to quantify predicted material
loss
in space based on material loss in ground laboratory thermal energy
plasma
testing. A microindent hardness comparison of surface oxidation of a
silicone
flown on the Environmental Oxygen Interaction with Materials III
(EOIM-III)
experiment with samples exposed in thermal energy plasmas was similarly
used to calibrate the rate of oxidation of silicone in space relative
to
samples in thermal energy plasmas exposed to polyimide Kapton effective
fluences.
Monte Carlo Computational Techniques for
Prediction of Atomic Oxygen Erosion of Materials
Materials on the surface of spacecraft in low Earth orbit (LEO)
are exposed to the remnants of the Earth's upper atmosphere. Energetic
solar photons cause photodissociation of O2 to produce
highly
reactive atomic oxygen. As spacecraft orbit through the atomic oxygen,
impact energies of 4.5± 1 eV result with an arrival flux sufficient to
cause polymeric materials to be oxidized at rates high enough
durability
concerns. To increase materials durability adequate to meet spacecraft
mission lifetime requirements, atomic oxygen protective coatings have
been applied over polymers. Such coatings typically consist of metal
oxide
thin films. The durability of such protected polymers used for solar
array
blankets and thermal control is limited as a result of microscopic
defects
in the protective films.
Environmental Exposure Conditions for
Teflon® FEP on the Hubble Space Telescope
The outer layer of Teflon® fluorinated ethylene propylene
(FEP) multi-layer insulation (MLI) on the Hubble Space Telescope (HST)
was observed to be significantly cracked at the time of the Second HST
Servicing Mission (SM2), 6.8 years after HST was launched into low
Earth orbit (LEO). Comparatively minor embrittlement and cracking were
also observed in FEP materials retrieved from solar-facing surfaces on
HST at the time of the First Servicing Mission (3.6 years exposure).
After SM2, a Failure Review Board was convened to address the problem
of degradation of MLI on HST.
In order for this board to determine possible degradation mechanisms,
it
was necessary to consider all environmental constituents to which the
FEP
MLI surfaces were exposed. Based on measurements and various models,
environmental exposure conditions for FEP surfaces on HST were
estimated including; number and temperature ranges of thermal cycles;
equivalent sun hours; fluence
and absorbed radiation dose of x-rays, trapped protons, and plasma
electrons and protons; and atomic oxygen (AO) fluence. This paper
presents the environmental exposure conditions for FEP on the Hubble
Space Telescope, briefly describing the possible roles of the
environmental factors in the observed FEP embrittlement and providing
references to the published works which describe in detail testing and
analysis related to FEP degradation on HST.
Simulated Solar Flare X-Ray and Thermal
Cycling Durability Evaluation of Hubble Space Telescope Thermal Control
Candidate Replacement Materials
During the Hubble Space Telescope (HST) second servicing mission (SM2),
astronauts noticed that the multi-layer insulation (MLI) covering the
telescope was damaged. Large pieces of the outer layer of MLI
(aluminized Teflon® fluorinated ethylene propylene (Al-FEP))
were cracked in several locations around the telescope. A piece of
curled up Al-FEP was retrieved by the astronauts and was found to be
severely embrittled, as witnessed by ground testing. The national
Aeronautics and Space Administration (NASA) Goddard Space Flight Center
(GSFC) organized a HST MLI Failure
Review Board (FRB) to determine the damage mechanism of the Al-FEP in
the HST environment, and to recommend a replacement thermal control
outer
layer to be installed on HST during the subsequent servicing missions.
Candidate
thermal control replacement materials were chosen by the FRB and tested
for
environmental durability under various exposures and durations by GSFC
and
NASA Glenn Research Center (GRC). This paper describes durability
testing
at GRC of candidate materials which were exposed to charged particle
radiation,
simulated solar flare x-ray radiation, and thermal cycling under load.
Samples
were evaluated for changes in solar absorptance and tear resistance.
Descriptions
of environmental exposures and durability evaluations of these
materials
are presented.
Analysis of Retrieved Hubble Space
Telescope Thermal Control Materials
The mechanical and optical properties of the thermal control materials
on the Hubble Space Telescope (HST) have degraded over the nearly 7
years the telescope has been in orbit. Astronaut observations and
photographs from the second servicing mission (SM2) revealed large
cracks in the metallized Teflon® FEP, the outer layer of the
mulit-layer insulation
(MLI), in many locations around the telescope. Also, the emissivity of
the bonded metallized Teflon® FEP radiator surfaces of the
telescope has increased over time. Samples of the top layer of the MLI
and radiator material were retrieved during SM2, and a thorough
investigation
into the degradation following in order to determine the primary cause
of damage. Mapping of the cracks on HST and the ground testing showed
that
thermal cycling with deep-layer damage from electron and proton
radiation
are necessary to cause the observes embrittlement. Further, strong
evidence
was found indicating that chain scission (reduced molecule weight) is
the
dominant form of damage to the metallized Teflon® FEP.
Mechanical Properties Degradation of
Teflon® FEP Returned form the Hubble Space Telescope
After 6.8 years in orbit, degradation has been observed in the
mechanical properties of second-surface metallized Teflon®
FEP (fluorinated ethylene propylene) used on the Hubble Space telescope
(HST) on the outer surface of the multi-layer insulation (MLI) blankets
and on radiator
surfaces. Cracking of FEP surfaces on HST was first observed upon close
examination of samples with high solar exposure retrieved during the
first servicing mission (SM1) conducted 3.6 years after HST was put
into
orbit. Astronaut observations and photographs from the second servicing
mission (SM2), conducted after 6.8 years on orbit, revealed severe
cracks
in the FEP surfaces of the MLI on many locations around the telescope.
This paper describes results of mechanical properties testing of FEP
surfaces
exposed for 3.6 and 6.8 years to the space environment on HST. These
tests
include bend testing, tensile testing, and surface micro-hardness
testing.
Investigation of Teflon FEP Embrittlement
on Spacecraft in Low Earth Orbit
Teflon FEP (fluorinated ethylene-propylene) is commonly used on
exterior spacecraft surfaces in the low Earth orbit (LEO) environment
for thermal control. Silverized or aluminized FEP is used for the outer
layer of thermal control blankets because of its low solar absorptance
and high thermal emittance. FEP is also preferred over other spacecraft
polymers because of its relatively high resistance to atomic oxygen
erosion.
Because of its low atomic oxygen erosion yield, FEP has not been
protected
in the space environment. Recent, long term space exposures such as on
the Long Duration Exposure Facility (LDEF, 5.8 years in space), and the
Hubble Space Telescope (HST, after 3.6 years in space) have provided
evidence
of LEO environmental degradation because of long durations and the
different
conditions (such as differences in altitude) of the exposures. Samples
of
FEP from LDEF and from HST (retrieved during its first servicing
mission)
have been evaluated for solar induced embrittlement and for synergistic
effects of solar degradation and atomic oxygen. Micro-indenter results
indicate
that the surface hardness increased as the ratio of atomic oxygen
fluence
to solar fluence decreased for the LDEF samples, but the solar
exposures
were higher. Cracks induced during bend testing were significantly
deeper
for the HST samples with the higher solar exposure than for LDEF
samples with similar oxygen fluence to solar fluence ratios. If solar
fluences are compared, the LDEF samples appear as damaged as the HST
samples, except
that HST had deeper induced cracks. The results illustrate difficulties
in comparing LEO exposed materials from different missions. Because the
HST FEP appears more damaged than the LDEF FEP based on the depth of
embrittlement,
other causes for FEP embrittlement in addition to the atomic oxygen and
ultraviolet (UV) radiation, such as thermal effects and the possible
role
of soft x-ray radiation, need to be considered. FEP that was exposed to
soft x-rays in a ground test facility, showed embrittlement similar to
that
witnessed in LEO, which indicates that the observed differences between
LDEF
and HST FEP might be attributed to the different soft x-ray fluences
during
these two missions.
Degradation of FEP Thermal Control
Materials Returned from the Hubble Space Telescope
After an initial 3.6 years of space flight, the Hubble Space Telescope
(HST) was serviced through a joint effort with the National Aeronautics
and Space Administration (NASA) and the European Space Agency (ESA).
Multi-layer insulation (MLI) was retrieved from the electronics boxes
of the two magnetic sensing systems (MSS), also called the
magnetometers, and from the returned solar array (SA-I) drive arm
assembly. The top layer of each MLI assembly is fluorinated ethylene
propylene (FEP, a type of Teflon). Dramatic changes in material
properties were observed when comparing areas of high solar fluence to
areas of low solar fluence. Cross sectional analysis shows atomic
oxygen (AO) erosion values of up to 25.4m m (1 mil). Greater
occurrences of through-thickness cracking and surface microscopy were
observed in areas of high solar exposure. Atomic force microscopy (AFM)
showed increases in surface microhardeness measurements with increasing
solar exposure. Decreases in FEP tensile strength and elongation were
measured when compared to non-flight material. Erosion yield and
tensile results are compared with FEP data from the Long Duration
Exposure Facility (LDEF). AO erosion yield data, solar fluence values,
contamination, micrometeoroid or debris (MMD) impact sites, and optical
properties are presented.
Effects of Heating on Teflon® FEP Thermal
Control Material from the Hubble Space Telescope
Metallized Teflon® FEP (fluorinated ethylene propylene)
thermal control material on the Hubble Space Telescope (HST) is
degrading in the space environment. Teflon® FEP thermal
control blankets (space-facing FEP) retrieved during the first service
mission (SM1) were found to be embrittled on solar facing surfaces and
contained microscopic cracks. During the second servicing mission (SM2)
astronauts noticed that the FEP outer layer of the multi-layer
insulation (MLI) covering the telescope was cracked in many locations
around the telescope. Large cracks were observed on the light shield,
forward shell, and equipment bays. A tightly curled piece of cracked
FEP from the light shield was retrieved during SM2 and was severely
embrittled, as witnessed by ground testing. A Failure Review Board
(FRB) was organized to determine the mechanism causing the MLI
degradation. Density, x-ray crystallinity, and solid state nuclear
magnetic resonance (NMR) analyses of FEP retrieved during SM1 were
inconsistent with results of FEP retrieved during SM2. Because the
retrieved SM2 material curled while in space, it experienced a higher
temperature extreme during thermal cycling, estimated at 200° C, than
the SM1 material, estimated at 50° C. An investigation on the effects
of heating pristine and FEP exposed on HST was therefore conducted.
Samples of pristine, SM1, and SM2 FEP were heated to 200° C and
evaluated for changes in density and morphology. Elevated temperature
exposure was
found to have a major impact on the density of the retrieved materials.
Characterization of polymer morphology of as-received and heated FEP
samples by NMR provided results that were consistent with the density
results. These findings have provided insight to the damage mechanisms
of FEP in the space environment.
Hubble Space Telescope Metalized Teflon®
FEP Thermal Control Materials: On-Orbit Degradation and Post-Retrieval
Analysis
During the Hubble Space Telescope (HST) second servicing mission (SM2),
degradation of unsupported Teflon® FEP (fluorinated ethylene
propylene), used as the outer layer of the multi-layer insulation (MLI)
blankets, was evident as large cracks on the telescope light shield. A
sample
of the degraded outer layer was retrieved during the mission and
returned
to Earth for ground testing and evaluation. The results of the Teflon®
FEP sample evaluation and additional testing of pristine Teflon®
FEP led the investigative team to theorize that the HST damage was
caused by thermal cycling with deep-layer damage from electron and
proton radiation which allowed the propagation of cracks along stress
concentrations, and
that the damage increased with the combined total dose of electrons,
protons,
ultraviolet and x-ray radiation along with thermal cycling. This paper
discusses
the testing and evaluation of the retrieved Teflon® FEP.
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