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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.
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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.
Preliminary Analysis of Polymer Film
Thermal Control and Gossamer Materials Experiments on Materials
International Space Station Experiment (MISSE 1 and MISSE 2)
A total of 31 samples were included in the
National Aeronautics and Space Administration (NASA) Glenn Research
Center (GRC) Polymer Film Thermal Control (PFTC) and Gossamer Materials
experiments, which were exposed to the low Earth orbit environment for
nearly four years on the exterior of the International Space Station
(ISS) as part of the Materials International Space Station Experiment
(MISSE 1 and MISSE 2). MISSE is a materials flight experiment
sponsored by the Air Force Research Lab/Materials Lab and NASA.
This paper describes objectives, materials, and characterizations for
the MISSE 1 and MISSE 2 GRC PFTC and Gossamer Materials samples.
Samples included films of polyimides, fluorinated polyimides, and
TeflonÒ fluorinated ethylene propylene (FEP) with and without
second-surface metalizing layers and/or surface
coatings. Also included were films of polyphenylene
benzobisoxazole (PBO)
and a polyarylene ether benzimidazole (TOR-LMTM). Polymer film
samples were
examined post-flight for changes in mechanical and optical properties.
The
environment in which the samples were located was characterized through
analysis
of sapphire contamination witness samples and samples dedicated to
atomic
oxygen (AO) erosion measurements. Results of the preliminary
analyses of
the PFTC and Gossamer Materials experiments are discussed.
Effects of the Space Environment on
Polymer Film Materials Exposed on the Materials International Space
Station Experiment (MISSE 1 and MISSE 2)
A total of 28 polymer film samples were included in the National
Aeronautics and Space Administration (NASA) Glenn Research Center (GRC)
Polymer Film Thermal
Control (PFTC) and Gossamer Materials Experiments, which were exposed
to
the low Earth orbit environment for nearly four years on the exterior
of
the International Space Station (ISS) as part of the Materials
International Space Station Experiment (MISSE 1 and MISSE 2).
MISSE is a materials flight experiment sponsored by the Air Force
Research Lab/Materials Lab and NASA. This paper will describe
objectives, materials, and characterizations for the MISSE 1 and MISSE
2 GRC PFTC and Gossamer Materials samples. Samples included films
of polyimides, fluorinated polyimides, and TeflonÒ fluorinated ethylene
propylene (FEP) with and without second-surface metalizing layers
and/or surface coatings. Also included were films of
polyphenylene benzobisoxazole (PBO) and a polyarylene ether
benzimidazole (TOR-LMTM). Polymer film samples were examined
post-flight for changes in mechanical and optical properties and
for atomic oxygen (AO) erosion. Results of the preliminary
analyses of the PFTC and Gossamer Materials Experiments are discussed.
Issues and Advancements in
Space
Durable Multi-Functional Thermal Control Coatings
Passive spacecraft thermal control coatings are required to possess
properties of low solar absorptance, high thermal emittance, and
stability
to survive the space environment for the mission duration. The
white paint
coatings Z-93, YB-71 and S13G/LO, originally developed in the 1960s,
have
been successfully used for satellite thermal control and have served as
standards for spacecraft white thermal control paints. Since
their original
development, these coatings have gone through re-formulations as
original
raw materials became unavailable; however, their replacement products
continue
to serve as standards for spaceflight thermal control. Unique
conditions
of space exploration and space science missions have required that
additional
functionalities be incorporated into spacecraft thermal control
coatings.
Coating development efforts have addressed needs for long-life
stability,
surface conductivity, and the ability to clean coating surfaces.
Advancements
in development of lightweight composite structures for spacecraft have
led
to the need for thermal control coatings that are adherent and
compatible
with these composite substrates, whereas the original formulations of
white
paints were developed for application to aluminum substrates. The
pursuit
of nuclear reactor powered spacecraft for future missions requires
coating/substrate
systems which are not only compatible with harsh space radiation
environmental
exposures, but must also perform at higher temperatures than have been
previously
required. Future missions to the lunar and Martian surfaces will
additionally
require thermal control coatings for which dust accumulation can be
mitigated.
Although advancements continue in the area of thermal control materials
technologies,
thermal control coatings are not currently commercially available to
meet
all of these advanced requirements. This paper presents some of
the unique
challenges for thermal control material systems for future space
missions
and some current approaches to meeting these challenges.
Effects of Vacuum Ultraviolet
Radiation of Various Wavelength Ranges on Teflon FEP Film
This paper describes testing to investigate the
effects of vacuum ultraviolet (VUV) radiation on Teflon® fluorinated
ethylene
propylene (FEP) film, examining differences in mechanical properties
degradation for samples of 50.8 m thickness exposed to VUV of various
lower cut-off wavelengths. Samples were illuminated in a
high vacuum facility by deuterium lamps, which provide radiation in the
115-400 nm wavelength range, but
with the highest intensity being below 200 nm. Windows of
fused silica,
crystalline quartz, and magnesium fluoride provided lower cut-off
wavelengths
of 155, 140, and 115 nm, respectively. Lamp intensity was
measured in
the 115-200 nm wavelength range throughout the sample exposures. The
determined
intensities were used to estimate intensity and incident energy of
various
wavelength ranges, between 115 and 400 nm. Samples were analyzed
for tensile
strength and elongation at failure. The effects of radiation
exposures
of different wavelength ranges were compared and discussed in terms of
the
expected depth to which radiation of various wavelengths is deposited
into
FEP.
Use of Hubble Space Telescope
Degradation Data for Ground-based Durability Projection of EPTFE on ISS
Ground-based environmental durability tests have indicated that
exposing materials in accelerated tests to environmental model
predicted spacecraft mission exposures of known degradation 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 exposure levels required to replicate degraded properties
observed in space. This paper describes a ground-to-space
correlation method that uses a multiple step 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 fluorinated ethylene propylene (FEP) thermal
insulation covering the Hubble Space Telescope (HST). This method
was designed to damage the back surface of equivalent thickness ePTFE
to the same amount of scission damage as occurred in 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.
The analysis indicates that after a 10 year mission, the ISS ePTFE will
have an extremely embrittled front surface, with surface cracks induced
under any given strain, and a very ductile back surface. This
study also found that a thermal induced strain of 0.1 will develop in
the ePTFE, and under this strain condition, microscopic cracks will
start developing very early in the mission at the exposed surface and
develop to a depth of ≈ 300 μm after 10 years.
Vacuum Ultraviolet Radiation and Atomic
Oxygen Durability Evaluation of HST Bi-Stem Boom Thermal Shield
Materials
Bellows-type thermal shields were proposed for use on the Hubble Space
Telescope (HST) solar array bi-stem booms to reduce the thermal
gradient-induced jitter during orbital thermal cycling. Candidate
thermal shield materials included aluminized FEP Teflon with and
without protective coatings for durability to atomic oxygen (AO) and
combined AO and ultraviolet (UV) radiation. NASA Lewis (now Glenn)
Research Center performed vacuum ultraviolet (VUV) radiation and AO
durability testing of candidate materials as part of an overall program
coordinated by NASA Goddard Space Flight Center (GSFC) to evaluate the
on-orbit durability of these thermal shield materials.
Coating adhesion problems were observed for samples having AO-
and the combined AO/UV-protective coatings which were attributed to
exposure to rapid thermal cycling used to simulate thermal cycling on
orbit. Such adhesion problems led to production of coating flakes from
the material which could pose a significant risk to HST optics if the
coated materials were used for the bi-stem boom thermal shields. No
serious degradation was observed for the uncoated aluminized Teflon as
evaluated by optical microscopy, although atomic force microscopy (AFM)
revealed that an embrittled surface layer would build up on the
uncoated Teflon surface due to ultraviolet
radiation exposure. This embrittled layer was not completely removed by
AO erosion. Despite the formation of this embrittled layer, no cracks
or
particle flakes were produced for the uncoated material upon exposure
to
VUV and AO.
Uncoated aluminized FEP Teflon was determined to be the most
appropriate thermal shield material and was used on the replacement
solar arrays installed during the December 1993 First HST Servicing
Mission.
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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 metalizing 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.
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.
Effect of Air and Vacuum Storage on the
Tensile Properties of X-ray Exposed Aluminized-FEP
Metallized Teflon® FEP (fluorinated ethylene propylene), a common
spacecraft thermal control material, from the exterior layer of the
Hubble Space Telescope (HST) has become embrittled and suffers from
extensive cracking. Teflon samples retrieved during Hubble servicing
missions and from the Long Duration Exposure Facility (LDEF) indicate
that there may be continued degradation in tensile properties over
time. An investigation has been conducted to evaluate the effect of air
and vacuum storage on the mechanical properties of x-ray exposed FEP.
Aluminized-FEP (Al-FEP) tensile samples were irradiated with 15.3 kV Cu
x-rays and stored in air or under vacuum for various time periods.
Tensile data indicate that samples stored in air display larger
decreases in tensile properties than for samples stored under vacuum.
Air-stored samples developed a hazy appearance, which corresponded to a
roughening of the aluminized surface. Optical property changes were
also characterized. These findings indicate that air exposure plays a
role in the degradation of irradiated FEP, therefore proper sample
handling
and storage is necessary with materials retrieved from space.
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.
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.
Portable Infrared Reflectometer for
Evaluating Emittance
Optical methods are frequently used to evaluate the emittance
of candidate spacecraft thermal control materials. One new optical
method utilizes a portable infrared reflectometer capable of obtaining
spectral reflectance of an opaque surface in the range of 2 to 25
microns using a Michelson-Type FTIR interferometer. This miniature
interferometer collects many infrared spectra over a short period of
time. It also allows the size of the instrument to be small such that
spectra can be collected in the laboratory or in the field. Infrared
spectra are averaged and integrated with respect to the room
temperature black body spectrum to yield emittance at 300 K.
Integrating with respect to other black body spectra yields emittance
values at other temperatures. Absorption bands in the spectra may also
be used for chemical species identification. The emittance of several
samples was evaluated using this portable infrared reflectometer, an
old
infrared reflectometer equipped with dual rotating black body cavities,
and a bench top thermal control vacuum chamber. Samples for evaluation
were purposely selected such that a range of emittance values and
thermal control material types would be represented, including polished
aluminum, Kapton®, silvered Teflon®, and the inorganic paint Z-93-P.
Results indicate an excellent linear relationship between the room
temperature emittance calculated from infrared spectral data and the
emittance obtained from the dual rotating black body cavities and
thermal vacuum chamber. The prospect of using the infrared spectral
date for chemical species identification will also be discussed.
Correlation of Predicted and Observed Optical
Properties of Multi-Layer Thermal Control Coatings
Thermal control coatings on spacecraft will be increasingly important
as spacecraft grow smaller and more compact. New thermal control
coatings will be needed to meet the demanding requirements of next
generation
spacecraft. Computer programs are now available to design optical
coatings,
and one such program was used to design several thermal control
coatings
consisting of alternating layers of WO3 and SiO2.
The coatings were subsequently manufactured with electron beam
evaporation
and characterized with both optical and thermal techniques. Optical
data
were collected in both the visible region of the spectrum and the
infrared.
Solar absorptance values were predicted in the range of 0.177-0.196 and
were observed in the range of 0.155-0.228. Infrared emittance values
were
predicted in the range of 0.074-0.083 and were observed optically in
the
range 0.048-0.093 and calorimetrically in the range of 0.069-0.100.
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.
Synchrotron VUV and Soft X-Ray Radiation
Effects on Aluminized Teflon® FEP
Surfaces of the aluminized Teflon® FEP multi-layer
thermal insulation on the Hubble Space Telescope were found to be
cracked
and curling in some areas at the time of the second servicing mission
in February 1997, 6.8 years after HST was deployed in low Earth orbit
(LEO). As a part of a test program to assess environmental conditions
which
would produce embrittlement sufficient to cause cracking of Teflon®
on HST, samples of Teflon® FEP with a backside layer of
vapor
deposited aluminum were exposed to vacuum ultraviolet (VUV) and soft
x-ray radiation of various energies using facilities at the National
Synchrotron Light Source, Brookhaven National Laboratory. Samples were
analyzed for ultimate tensile strength and elongation. Results will be
compared to those of aluminized Teflon® FEP retrieved from
HST
after 3.6 years and 6.8 years on orbit and will be referenced to
estimated
HST mission doses of VUV and soft x-ray radiation.
A Comparison of Space and Ground Based
Facility Environmental Effects on FEP Teflon®
Fluorinated Ethylene Propylene (FEP) Teflon® is widely used
as a thermal control material for spacecraft, however, it is
susceptible to erosion, cracking, and subsequent mechanical failure in
low Earth orbit. One of the difficulties in determining whether FEP
Teflon® will survive during a mission is the wide disparity
of erosion rates
observed for this material in space and in ground based facilities.
Each
environment contains different levels of atomic oxygen, ions, and
vacuum
ultraviolet (VUV) radiation in addition to parameters such as the
energy
of the arriving species and temperature. These variations make it
difficult
to determine what is causing the observed differences in erosion rates.
This paper attempts to narrow down which factors affect the erosion
rate
of FEP Teflon® through attempting to change only one
environmental
constituent at a time. This was attempted through the use of a single
simulation facility (plasma asher) environment with a variety of
Faraday
cages and VUV transparent windows. Isolating one factor inside of a
radio
frequency (RF) plasma proved to be very difficult. Two observations
could
be made. First, it appears that the erosion yield of FEP Teflon®
with respect to that of polyimide Kapton is not greatly affected by the
presence or lack of VUV radiation present in the RF plasma and the
relative
erosion yield for the FEP Teflon® may decrease with
increasing
fluence. Second, shielding from charged particles appears to lower the
relative erosion yield of the FEP to approximately that observed in
space,
however, it is difficult to determine for sure whether ions, electrons,
or some other components are causing the enhanced erosion.
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 metalized Teflon® FEP, the outer layer of the
mulit-layer insulation (MLI), in many locations around the telescope.
Also, the emissivity of the bonded metalized 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 embnttlement. Further, strong
evidence
was found indicating that chain scission (reduced molecule weight) is
the
dominant form of damage to the metalized Teflon® FEP.
Evaluation and Selection of Replacement
Thermal Control Materials for the Hubble Space Telescope
The mechanical and optical properties of the metalized Teflon®
FEP thermal control materials on the Hubble Space Telescope (HST) have
degraded over the nearly 7 years the telescope has been in orbit. Given
the damage to the outer layer of the multi-layer insulation (MLI) that
was apparent during the second servicing mission (SM2), the decision
was made to replace the outer layer during subsequent servicing
missions. A
Failure Review Board (FRB) was established to investigate the damage to
the MLI and identify a replacement material. The replacement material
had
to meet the stringent thermal requirements of the spacecraft and
maintain
structural integrity for at least 10 years. Ten candidate materials
were
selected and exposed to ten-year HST-equivalent doses of simulated
orbital
environments. Samples of the candidates were exposed sequentially to
low
and high-energy electrons and protons, atomic oxygen, x-ray radiation,
ultraviolet
radiation, and thermal cycling. Following the exposures, the mechanical
integrity
and optical properties of the candidates were investigated using
Optical
Microscopy, Scanning Electron Microscopy (SEM), and a Laboratory
Portable
Spectroreflectometer (LPSR). Based on the results of these simulations
and
analyses, the FBR selected a replacement material and two alternates
that
showed the highest likelihood of providing the requisite thermal
properties
and surviving for 10 years in orbit.
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 metalized 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.
Ground Based Testing of Replacement Thermal
Control Materials for the Hubble Space Telescope
The mechanical and optical properties of the metallized Teflon FEP
thermal control materials on the Hubble Space Telescope (HST) have
degraded over the nearly seven years the telescope has been in orbit.
Given
the damage to the outer layer of the multi-layer insulation (MLI)
blanket
that was apparent during the second servicing mission (SM2), the
decision
was made to replace the outer layer during subsequent servicing
missions.
A Failure Review Board was established to investigate the damage to the
MLI and identify a replacement material. The replacement material had
to
meet the stringent thermal requirements of the spacecraft and maintain
mechanical
integrity for at least ten years. Ten candidate materials were selected
and exposed to ten-year HST-equivalent doses of simulated orbital
environments.
Samples of the candidates were exposed sequentially to low- and
high-energy
electrons and protons, atomic oxygen, x-ray radiation, ultraviolet
radiation,
and thermal cycling. Following the exposures, the mechanical integrity
and
optical properties of the candidates were investigated using optical
microscopy,
scanning electron microscopy (SEM), a laboratory portable
spectroreflectometer (LPSR) and a Lambda 9 spectroreflectometer. Based
on the results of these simulations and analyses, the Failure Review
Board selected a replacement material and two alternatives that showed
the highest likelihood of providing the requisite thermal properties
and surviving for ten years in orbit.
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 microhardenss 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.
Effects of Radiation and Thermal Cycling on
Teflon® FEP
Surfaces of the aluminized Teflon® FEP (fluorinated
ethylene propylene) multilayer thermal insulation on the Hubble Space
Telescope (HST) were found to be cracked and curled in some areas at
the time of the second servicing mission (SM2) in February 1997, 6.8
years after
HST was deployed into low Earth orbit (LEO). In an effort to understand
what elements of the space environment might cause such damage,
pristine second-surface aluminized Teflon® FEP was tested
for durability to various
types of radiation, to thermal cycling and to radiation followed by
thermal
cycling. Types of radiation included synchrotron vacuum ultraviolet and
soft x-ray radiation, electrons and protons. Thermal cycling was
conducted in various temperature ranges to simulate HST orbital
conditions for Teflon® FEP. Results of tensile testing of
the exposed specimens showed that
exposure to high fluences of radiation caused degradation in tensile
properties
of FEP. However, exposure to radiation alone in exposures comparable to
those experienced by HST did not produce reduction in ultimate tensile
strength and elongation of Teflon® similar to that observed
for HST-retrieved aluminized Teflon®. Synergism of radiation
exposure and thermal cycling was evident in the results of three
experiments: thermal cycling following electron and proton irradiation,
thermal cycling following x-ray exposure, and additional thermal
cycling of a sample retrieved from HST. However, irradiation and
thermal cycling with comparable HST SM2 exposure conditions did not
produce the degradation observed in the FEP material
retrieved during HST SM2.
Thermal Cycling-Caused Degradation of Hubble
Space Telescope Aluminized FEP Thermal Insulation
The Hubble Space Telescope (HST) was launched in April of 1990 and was
visited during service missions in December of 1993 and February of
1997. During the latter servicing mission, astronauts observed that the
top layer of multi-layer insulation, which consisted of second surface
aluminized FEP Teflon®, has occasional tears in its 0.127 mm
thick outer layer. A sample was retrieved which had torn and rolled up
under its own stress such that the aluminized layer was on the exposed
surface.
The sample was found to have an increase in solar absorptance and has
multiple
cracks in the aluminization in a mud-tile configuration. Tests
conducted
in a ground laboratory high-rate thermal cycling system indicate that a
signification portion of the observed increase in solar absorptance may
have been caused by cracks in the fatigued aluminum as a result of
approximately
40.000 thermal cycles it received in space.
Effect of X-Rays on the Mechanical
Properties
of Aluminized FEP Teflon
Pieces of the multilayer insulation (MLI) that is integral to
the thermal control of the Hubble Space Telescope (HST) have been
returned by two servicing missions after 3.6 and 6.8 years in orbit.
They reveal that the outer layer, which is made from 5 mil (0.13 mm)
thick aluminized fluorinated ethylenepropylene (FEP) Teflon®,
has become severely embrittled. Although possible agents of
embrittlement include electromagnetic radiation across the entire solar
spectrum, trapped particle radiation, atomic oxygen, and thermal
cycling, intensive investigations have not
yielded unambiguous causes. Previous studies utilizing monoenergenic
photons
in the 69-1900 eV range did not cause significant embrittlement, even
at much higher doses than were experienced by the HST MLI. Neither did
x-rays in the 3 to 10 keV range generated in a modified electron beam
evaporator. An antidotal aluminized FEP sample that was exposed to an
intensive dose from unfiltered Mo x-ray radiation from a rotating anode
generator, however, did show the requisite brittlement. Thus, a study
was undertaken to determine the effects of x-ray exposure on the
embrittlement of aluminized FEP in hopes that it might elucidate the
HST MLI degradation mechanism. Tensile specimens of aluminized 5 mil
thick FEP were exposed to a constant fluence of unfiltered x-ray
radiation from a Mo target whose maximum energy ranged from 20-60 kV.
Other samples were annealed, thermally cycled (100x) between 77-333 K,
or cycled and irradiated. Tensile tests and density measurements were
then performed on the samples which had been irradiated had the
drastically reduced elongation-to-break, characteristic of the HST
samples. Thermal cycling may accelerate the embrittlement, but the
effect was near the
scatter in the measurements. Annealing and thermal cycling had no
apparent
effect. Only the samples which had been irradiated and annealed showed
significant density increases, likely implicating polymer chain
scission and annealing.
Ground Laboratory Soft X-Ray Durability
Evaluation of Aluminized Teflon® FEP Thermal Control Insulation
Metallized Teflon® fluorinated ethylene propylene (FEP)
thermal control insulation is mechanically degraded if exposed to a
sufficient fluence of soft x-ray radiation. Soft x-ray photons (4 to 8
Å in wavelength or 1.55 to 3.2 keV) emitted during solar flares have
been proposed as a cause of mechanical properties degradation of
aluminized Teflon® FEP thermal control insulation on the
Hubble Space Telescope (HST). Such degradation can be characterized by
a reduction in elongation-to-failure of the Teflon® FEP.
Ground laboratory soft x-ray exposure tests of aluminized Teflon®
FEP were conducted to assess the degree of elongation degradation,
which would occur as a result of exposure
to soft x-rays in the range of 3 to 10 keV. Test results indicate that
soft x-ray exposure in the 3 to 10 keV range, at mission fluence
levels,
does not alone cause the observed reduction in elongation of flight
retrieved
samples. The soft x-ray exposure facility design, mechanical properties
degradation results, and implications will be presented.
Atomic Oxygen/Vacuum Ultraviolet Radiation
Exposure of Z-93 and Z-93-P Coatings
Laboratory testing was conducted in order to assess the long-term
atomic oxygen and vacuum ultraviolet radiation durability of the
thermal control coating Z-93-P to be used on the International Space
Station
radiator surfaces. This testing provided atomic oxygen equivalent to
approximately four years and vacuum ultraviolet radiation equivalent
to approximately twenty-five years on Space Station radiator surfaces.
Solar absorptance data were obtained in vacuo at various exposure
increments.
Facility limitations resulted in the inability to provide the
appropriate
atomic oxygen to vacuum ultraviolet radiation ratio that would be
experienced
by Space Station radiator surfaces, and unexpected sputtering of
components
in the vacuum chamber caused a contaminant layer to be deposited on the
samples. However, some conclusions can be made from the data. First,
Z-93-P
samples performed comparably to the Z-93 control sample assuring that
the
successful flight history of the original Z-93 formulation can be
applied
to the reformulated Z-93-P coating. Second, solar absorptance increases
of no more than 0.1 were calculated for the combined atomic oxygen and
vacuum
ultraviolet radiation exposure environment used in this test.
Combined Contamination and Space
Environmental Effects on Solar Cells and Thermal Control Surfaces
For spacecraft in low Earth orbit (LEO), contamination can occur from
thruster fuel, sputter contamination products, and from products of
silicone degradation. This paper describes laboratory testing in which
solar cell materials and thermal control surfaces were exposed to
simulated spacecraft environmental effects including contamination,
atomic oxygen, ultraviolet radiation and thermal cycling. The objective
of these experiments was to determine how the interaction of the
natural LEO environmental effects with contaminated spacecraft surfaces
impacts the performance of these
materials. Optical properties of samples were measured and solar cell
performance
data was obtained. In general, exposure to contamination by thruster
fuel
resulted in degradation of solar absorptance for fused silica and
various
thermal control surfaces and degradation of solar cell performance.
Fused
silica samples which were subsequently exposed to an atomic
oxygen/vacuum
ultraviolet radiation environment showed reversal of this degradation.
These
results imply that solar cells and thermal control surfaces which are
susceptible to thruster fuel contamination and which also receive
atomic oxygen exposure may not undergo significant performance
degradation. Materials which were exposed to only vacuum ultraviolet
radiation subsequent to contamination showed, slight additional
degradation in solar absorptance.
The Effects of Simulated Low Earth Orbit
Environments on Spacecraft Thermal Control Coatings
Candidate Space Station Freedom radiator coatings including Z-93,
YB-71, anodized aluminum, and SiOx-coated silvered Teflon
have been characterized for optical properties degradation upon
exposure
to environments containing atomic oxygen, vacuum ultraviolet (VUV)
radiation and/or silicone contamination. YB-71 coatings showed a
blue-gray discoloration, which has not been observed in space, upon
exposure in atomic oxygen facilities which also provide exaggerated VUV
radiation. This is evidence that damage mechanisms occur in these
ground laboratory facilities which are different from those which occur
in space. Radiator coatings exposed to an electron cyclotron resonance
(ECR) atomic oxygen source in the presence of silicone-containing
samples showed severe darkening form the intense VUV radiation provided
by the ECR and from silicone contamination. Samples exposed to atomic
oxygen from the ECR source and to VUV lamps, simultaneously, with in
situ reflectance measurement, showed that significantly greater
degradation
occurred when samples received line-of-site ECR beam exposure than when
samples were exposed to atomic oxygen scattered off of quartz surfaces
without line-of-site view of the ECR beam. For white paints, exposure
to air following atomic oxygen/VUV exposure reversed the darkening due
to VUV damage. This illustrates the importance of in situ reflectance
measurement.
Evaluation of Low Earth Orbit Environmental
Effects on International Space Station Thermal Control Materials
Samples of International Space Station (ISS) thermal control coatings
were exposed to simulate low Earth orbit (LEO) environmental conditions
to determine effects on optical properties. In one test, samples of the
white paint coating Z-93P were coated with outgassed products from
Tefzel® (ethylene tetrafluoroethylene copolymer) power cable
insulation as may occur on ISS. These samples were then exposed, along
with an uncontaminated Z-93P witness sample, to vacuum ultraviolet
(VUV) radiation to determine solar absorptance degradation. The Z-93P
samples coated with Tefzel® outgassing products experienced
greater increases in solar absorptance than witness samples not coated
with Tefzel® outgassing products. In another test, samples
of second surface silvered Teflon® FEP (fluorinated ethylene
propylene), SiOx (where x(2)-coated silvered Teflon®
FEP, and Z-93P witness samples were exposed
to the combined environments of atomic oxygen and VUV radiation to
determine optical properties changes due to these simulated ISS
environmental effects. This test verified the durability of these
materials in the absence of contaminants.
Optical and Calorimetric Evaluation of
Z-93-P
and Other Thermal Control Coatings
The hemispherical total emissivity of two thermal control coatings,
Z-93-P and black anodized aluminum, was calculated from hemispherical
total reflectivity measured in wavelength range of 2 to 40 m m. These
data were compared to hemispherical total emissivity values obtained on
the same samples measured in a thermal vacuum chamber with a
calorimetric
technique. The comparison showed close agreement in the vicinity of
room
temperature and above, with differing trends at lower temperatures.
Emittance of Thermal Control Materials
Between 100K and 400K
The hemispherical total emittances of several common thermal coatings
tapes were evaluated calorimetrically over a wide temperature range.
The calorimetric technique used here to evaluate thermal control
materials allows a thermally isolated sample to cool solely through
radiant heat transfer to a liquid nitrogen cooled cold wall. The
mechanism of cooling is similar to that found in space, providing a
functional evaluation of
hemispherical total emittance. The five samples that were evaluated
included
several first and second surface mirrored Kapton tapes, one
carbon-filled
Kapton tape, and one black film. These tapes were adhesively bonded to
an aluminum substrate, which provided the sensible heat for the
calorimetric
calculation. Temperature-time data were collected as heat flowed from
the
aluminum, through the adhesive layer, through the thermal control tape,
and ultimately through the surface to the surroundings. Hemispherical
total
emittances were calculated over the temperature range of 100 to 400 K.
Emissivity Characterization of Plastics,
Ceramics, and Coatings Using a Calorimetric Technique
The emissivity of several different materials, including commercially
available plastics, ceramics, and coatings, was evaluated by both
optical and calorimetric means to select appropriate materials of
construction for a microgravity combustion experiment. Samples of grit
blasted black anodized aluminum, grit blasted black oxidized stainless
steel, Bakelite®, Noryl®, Zelux®, Macor®,
and Eccocoat®, were all characterized in the range of 200 to
400 K. Several of these
materials exceeded the combustion experiment emissivity requirement of
0.8 over a wide range of temperatures, suggesting their promising use
as
materials of construction. Emissivity values from all of the materials
will
be summarized, and those materials selected for the microgravity
combustion
experiment will be identified.
Emittance Characterization of Thermal
Control
Paints, Coatings, and Surfaces using a Calorimetric Technique
Thermal control surfaces are used in every spacecraft thermal
management system to dissipate heat through heat transfer. This paper
describes the thermal performance of several thermal control paints,
coatings,
and surfaces, as characterized by a calorimetric vacuum emissometer.
The emissometer is designed to measure the functional emittance of a
surface
based on heat transfer from an underlying substrate to the surface and
from the surface or near surface to a surrounding cold wall. Emittance
measurements
were made between 200 and 350 K. Polished aluminum, used here as a
standard,
was found to have a total hemisperical emittance of 0.06, as expected.
A
velvet black paint, also used as a standard, was found to have an
emittance
of 0.94 at room temperature. Other surfaces of interest included a
polyurethane-based black paint designated Z-306, a highly polished 316L
stainless steel,
and an atomic beam-textured carbon-carbon composite.
Thermal Modeling of a Calorimetric Technique
for Measuring the Emittance of Surfaces and Coatings
A finite element analysis model of a transient technique used
to measure the emittance of surface and coatings was developed and used
to estimate the uncertainty in emittance. The dimensions used in the
model matched the dimensions used in the design of a low temperature
calorimetric vacuum emissometer being built to characterize the thermal
properties of space power materials in the temperature range 173-673 K.
Radiant energy from a quartz halogen lamp impinged on an aluminum
sample that was coated with a thermal control coating and suspended in
a liquid-nitrogen-cooled vacuum chamber by narrow gauge thermocouple
wires. After removing the heat source, the temperature of the sample
was monitored vs. time and the temperature-time curve was used to
calculate the emittance. Factors contributing to the uncertainty in the
emittance included uncertainties in time, temperature, area of the
sample, heat capacity of the sample, and heat loss from the
uncoated back of the sample. Heat losses from the thermocouple wires
were
found to be negligible. The total probable error in the emittance
obtained
from the low temperature calorimetric vacuum emissometer design was
estimated
to be less than 4% for emittance values greater than 0.5 at
temperatures
between 173 and 673 K.
Advances in Optical Property Measurements of
Spacecraft Materials
This report describes some of the instruments and experimental
approaches available for measuring optical properties of thermal
control materials. It also describes the instruments' uses in
laboratory studies of the effects of combined contaminants and the
space environment on these
materials, and in the qualification of hardware for spacecraft. In
recent
years, several instruments for measurement of solar absorptance (a )
and infrared emittance (e ) have been introduced. These instruments
offer
improved speed, accuracy, and data-handling, all of which substantially
improve the study of contaminated thermal control materials. A
transient method for directly measuring material e is also described,
and the results are compared with other instruments. In addition, our
understanding
of oxygen exposure effects on the ( of materials following
contamination
or exposure to simulated space conditions shows that oxygen exposure
before measuring of e should be avoided.
Reflectivity of Silver and Silver-coated
Substrates from 25°C to 800°C
A bench top facility was used to evaluate the reflectivity of
several candidate coating-substrate combinations in vacuum at elevated
temperatures. Silver was selected as the reflective coating of choice,
while copper, nickel, electroless nickel on copper, and 304 stainless
steel were selected as substrates. Pure silver, with no coating at all,
was also evaluated. An optically flat silver-coated sapphire substrate
was used as a standard. All metal substrates were either
metallurgically
polished or diamond turned to mirror finish prior to silver deposition.
Silicon dioxide was used as a protective coating in most cases.
Reflectivity
measurements were made at room temperature in the visible range with a
spectrophotometer,
and at elevated temperatures up to 800°C with a helium-neon laser at
632
nm. Results from the high temperature reflectivity measurements will be
present |
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