Kim K. de Groh

Glenn Research Center, Cleveland, Ohio

Bruce A. Perry

Ohio Aerospace Institute, Brook Park, Ohio

Jelila S. Mohammed

Goddard Space Flight Center, Greenbelt, Maryland

Bruce A. Banks

Science Application International Corporation, Cleveland, Ohio

Analyses of Hubble Space Telescope

Aluminized-Tefl on Multilayer Insulation

Blankets Retrieved After 19 Years of

Space Exposure

NASA/TM—2015-218476

February 2015
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Kim K. de Groh

Glenn Research Center, Cleveland, Ohio

Bruce A. Perry

Ohio Aerospace Institute, Brook Park, Ohio

Jelila S. Mohammed

Goddard Space Flight Center, Greenbelt, Maryland

Bruce A. Banks

Science Application International Corporation, Cleveland, Ohio

Analyses of Hubble Space Telescope

Aluminized-Tefl on Multilayer Insulation

Blankets Retrieved After 19 Years of

Space Exposure

NASA/TM—2015-218476

February 2015

National Aeronautics and

Space Administration

Glenn Research Center

Cleveland, Ohio 44135
Acknowledgments

The authors would like to thank Ben Reed and the HST Project Offi ce at NASA GSFC for their support of this project. We greatly 

appreciate the opportunity to analyze this unique material and provide the results to the space community. We would also like to

thank the following people from GSFC for their contributors to testing and analysis: Bryan Abbamonte, Josh Abel, David Hughes, 

Kristin McKittrick, Aparna Boddapati, Debbie Thomas and Mollie Powell.

Available from

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This report is available in electronic form at http://www.sti.nasa.gov/ and http://ntrs.nasa.gov/
NASA/TM—2015-218476 1 

Analyses of Hubble Space Telescope Aluminized-Teflon Multilayer 

Insulation Blankets Retrieved After 19 Years of Space Exposure 

Kim K. de Groh 

National Aeronautics and Space Administration 

Glenn Research Center 

Cleveland, Ohio 44135 

Bruce A. Perry 

Ohio Aerospace Institute 

Brook Park, Ohio 44142 

Jelila S. Mohammed 

National Aeronautics and Space Administration 

Goddard Space Flight Center 

Greenbelt, Maryland 20771 

Bruce A. Banks 

Science Application International Corporation 

Cleveland, Ohio 44135 

Abstract 

Since its launch in April 1990, the Hubble Space Telescope (HST) has made many important 

observations from its vantage point in low Earth orbit (LEO). However, as seen during five servicing 

missions, the outer layer of multilayer insulation (MLI) has become increasingly embrittled and has cracked 

in many areas. In May 2009, during the 5th servicing mission (called SM4), two MLI blankets were 

replaced with new insulation and the space-exposed MLI blankets were retrieved for degradation analyses 

by teams at NASA Glenn Research Center (GRC) and NASA Goddard Space Flight Center (GSFC). The 

retrieved MLI blankets were from Equipment Bay 8, which received direct sunlight, and Equipment Bay 5, 

which received grazing sunlight. Each blanket was divided into several regions based on environmental 

exposure and/or physical appearance. The aluminized-Teflon (DuPont, Wilmington, DE) fluorinated 

ethylene propylene (Al-FEP) outer layers of the retrieved MLI blankets have been analyzed for changes in 

optical, physical, and mechanical properties, along with chemical and morphological changes. Pristine and 

as-retrieved samples (materials) were heat treated to help understand degradation mechanisms. When 

compared to pristine material, the analyses have shown how the Al-FEP was severely affected by the space 

environment. Most notably, the Al-FEP was highly embrittled, fracturing like glass at strains of 1 to 8 

percent. Across all measured properties, more significant degradation was observed for Bay 8 material as 

compared to Bay 5 material. This paper reviews the tensile and bend-test properties, density, thickness, solar 

absorptance, thermal emittance, x-ray photoelectron spectroscopy (XPS) and energy dispersive spectroscopy 

(EDS) elemental composition measurements, surface and crack morphologies, and atomic oxygen erosion 

yields of the Al-FEP outer layer of the retrieved HST blankets after 19 years of space exposure. 

1.0 Introduction 

The HST was launched into LEO during STS-31 on April 25, 1990 as the first mission of NASA’s 

Great Observatories program. It is capable of performing observations in the near-ultraviolet, visible, and 

near-infrared wavelengths and was designed to be serviced on-orbit to upgrade scientific capabilities. 

Five servicing missions (SM) have taken place, with the last mission occurring in May 2009 after 19 years 

in space. 
NASA/TM—2015-218476 2 

Multilayer insulation (MLI) blankets covering over 80 percent of the surface of the HST are 

important for passive temperature control on-orbit. Aluminized-Teflon (DuPont, Wilmington, DE) 

fluorinated ethylene propylene (Al-FEP) is used as the outer layer of the MLI thermal control insulation 

because the FEP outer layer has excellent optical properties (low solar absorptance (Ds) and high thermal 

emittance (H)) and the vapor deposited aluminum backing reflects incident solar radiation. 

In general, polymers such as Teflon FEP experience degradation in mechanical, electrical, and 

thermal properties due to exposure to radiation (Ref. 1). Hubble Space Telescope surfaces are exposed to 

various types of radiation including solar ultraviolet (UV) radiation, x-rays from the Sun, and electron and 

proton particle radiation. Additionally, the temperature cycling experienced by HST insulation has been 

found to have an effect on mechanical properties of FEP, especially when passing through specific 

transition temperatures (Ref. 2). Furthermore, space-exposed FEP can be eroded by atomic oxygen, which 

is formed in low Earth orbit (LEO) from the photodissociation of molecular oxygen by short wavelength 

energetic UV radiation. By Servicing Mission 2 (SM2, during STS-82), in February 1997, severe 

cracking of the 5 mil (127 μm) thick Al-FEP outer layer of the MLI blankets was observed on the light 

shield, forward shell and equipment bays of the HST (Refs. 3 and 4). The worst cracks in MLI on 

Equipment Bays 8 and 10 were covered during the mission with two patches of 2 mil (50.8 μm) thick 

Al-FEP on each bay (Fig. 1). 

In May 2009 during the 5th servicing mission (called SM4, during STS-125), two degraded MLI 

blankets that were originally installed on the telescope and had 19 years of space exposure were replaced 

with new insulation pieces, called New Outer Blanket Layers (NOBLs). The space exposed MLI blankets 

were brought back for analyses by researchers at NASA Glenn Research Center and NASA Goddard 

Space Flight Center. No other spacecraft material has been retrieved and analyzed after having this length 

of space exposure. The two blankets retrieved during SM4 included Equipment Bay 8 MLI, which 

received direct sunlight, and Equipment Bay 5 MLI, which received grazing sunlight. Also retrieved were 

remnants of the two patches that were placed over cracked areas on Bay 8 during SM2 and were exposed 

to space for 12.3 years. The Al-FEP outer layers of the retrieved MLI blankets were found to be highly 

degraded, and have been analyzed for changes in optical, physical, and mechanical properties, along with 

space induced chemical and morphological changes and heat induced changes. This paper reviews the 

tensile and bend-test properties, density, thickness, solar absorptance, thermal emittance, x-ray

(a) (b) 

Figure 1.—On-orbit photos of Equipment Bay 8 at SM2 (a) Before with cracked areas 

circled, and (b) After placement of the 2 mil Al-FEP patches. 
NASA/TM—2015-218476 3 

photoelectron spectroscopy (XPS) and energy dispersive spectroscopy (EDS) elemental composition 

measurements, surface and crack morphologies, and atomic oxygen erosion values of the outer Al-FEP 

layer of the retrieved HST blankets after 19 years of space exposure. 

2.0 Materials and Environmental Exposure 

2.1 HST SM4 Bay 5 and Bay 8 MLI 

The Bay 5 and 8 MLI blankets originally installed on HST (exposed to space for 19.1 years), and 

2 mil Al-FEP patches installed on Bay 8 on February 18, 1997 during SM2 (exposed to space for 

12.2 years) were retrieved by astronaut John Grunsfeld on May 18, 2009 during SM4, as shown in 

Figure 2. As can be seen in the illustration in Figure 3, Bay 8 is located 15° from the +V3 solar facing 

axis, and therefore is almost directly facing the sun. Bay 5 faces towards the +V2 solar array drive arm 

direction, and at 75° from the +V3 direction, receives grazing sunlight. 

Figure 2.—Astronaut John Grunsfeld placing a NOBL on 

Bay 8 after he removed the damaged Bay 8 MLI and 

placed it folded into one of the EVA storage bags on his 

left along with the retrieved patch material. The second 

EVA bag contains the retrieved Bay 5 MLI blanket. 

Figure 3.—Locations and orientations of Bays 5 

and 8 on HST (+V3 is the solar facing axis). 

(Sun direction) 
NASA/TM—2015-218476 4 

2.2 Environmental Exposure 

The sun exposure on each bay was determined by the science pointing profile of HST. The attitude 

profile between January 1, 2000 and SM4 was analyzed to determine the “typical” attitude of HST. The 

results were then extrapolated to provide estimates of the sun exposure since launch. Equivalent hours of 

sun exposure (ESH) of the telescope from deployment to SM4 are estimated as 111,000 hr, based on time 

in orbit, average time exposed to the sun per orbit, and length of each orbit. Based on the attitude profile 

from 2000 through SM4, and extrapolating over mission life, Bay 5 was exposed to ~24,300 ESH and the 

unpatched and patched areas of the Bay 8 MLI were exposed to ~89,300 and ~30,300 ESH, respectively. 

The HST underwent an estimated 110,000 thermal cycles overall, from deployment to SM4. The 

range of temperatures seen by Bay 5 and Bay 8 vary greatly, because they are dependent on HST attitude 

and environmental heating variables. Thermal Desktop was used to model the Bay 5 and Bay 8 MLI and 

simulate the general thermal cycling behavior. Bay 5 MLI temperatures were estimated to range from 

–175 to 0 °C and the Bay 8 temperatures were estimated to range from –175 to 40 °C, for the attitude and 

orbit configurations modeled. A region of Bay 8 had been patched with a thin layer of Al-FEP in 1997, 

which resulted in decreased exposure to some space environmental effects such as radiation and atomic 

oxygen, but may have increased the maximum on-orbit temperature for the original top layer of the MLI 

blanket in that region. Additionally, a section of Bay 8 had undergone stress cracking causing a large 

piece to curl up while on orbit. This section may also have received higher maximum on-orbit 

temperatures than the flat regions of Bay 8 that remained intact and flat, while receiving some protection 

from radiation and atomic oxygen effects. 

The X-ray fluence for solar facing surfaces was computed to be 641.1 J/m2 between 1-8 Å and 

43.1 J/m2

 between 0.5-4 Å. Data for x-ray fluence are based on x-ray flux data from the Geosynchronous 

Operational Environmental Satellites (GOES) for the time period of launch (April 1990) through SM2 

(February 1997) (Ref. 5). For SM2 through SM4, the x-ray fluence was estimated assuming an average 

11-year solar cycle (Ref. 5). 

Electron and proton fluence from solar wind particles trapped in Earth’s magnetic field have been 

calculated from deployment until December 9, 1999 using NASA’s proton and electron models, AP-8 and 

AE-8, respectively (Ref. 5). The data for SM4 were extrapolated from these prior data, and indicated the 

Bay 5 and Bay 8 MLI received an electron fluence of 5.6u1013 particles/cm2

 for electrons >40 keV, and a

proton fluence of 5.4u1010 particles/cm2 for protons !40 keV. 

The atomic oxygen fluence was computed for the duration of the HST mission over the period of time 

from deployment to SM4. This was accomplished by adding: (1) the prediction over the period of time 

from deployment to SM1 as predicted by SAIC’s version 5.0 Environmental Work Bench, which uses 

MSIS-86 atmospheric model, to (2) a prediction based on orbital and atmospheric data from NASA 

Goddard for SM1 to SM4. The parameters used in the SM1 to SM4 fluence calculation included altitude, 

atmospheric density, and orbital velocity corrected for the Earth’s atmosphere co-rotation. Based on these 

calculations the total ram fluence was 2.61u1021 atoms/cm2

. If the surfaces of Bay 8 are 15° from solar 

facing, then the fluence would be ~25.8 percent of the ram fluence or ~6.73u1020 atoms/cm2

. But, as the 

surfaces are solar facing but randomly tipped (always with Bay 8 somewhat towards the sun), then the 

fluence is decreased an additional factor of 2/S resulting in the probable fluence of ~4.28u1020 atoms/cm2

. 

Because the Bay 5 faces towards the +V2 direction and is 75° off the solar facing axis, the fluence was 

estimated as ~28.0 percent of the ram fluence, or 7.30u1020 atoms/cm2

. As the surfaces are randomly 

tipped, the fluence would be decreased by 2/S resulting in a probable fluence of ~4.65u1020 atoms/cm2

. 

Environmental exposure for patched and curled regions was calculated based on the time periods 

when these regions were uncovered. The patches were assumed to protect a region until the midpoint 

between the last servicing mission when the patch was viewed covering the region and the first servicing 

mission when the region was observed unprotected. A table of the exposure values for all regions is 

provided in Section 4.1. 
NASA/TM—2015-218476 5 

3.0 Experimental Procedures 

3.1 Scanning Electron Microscopy 

Either a Hitachi S-4700 field emission scanning electron microscopes (FESEM) operated at an 

accelerating voltage of 6 kV, or a Hitachi S-3500 operated at an accelerating voltage of 10 kV, were used 

at GRC to document various surface features, such as erosion morphologies, cracks and impact sites. The 

samples were sputter-coated with thin gold-palladium (Au-Pd) films to make them conductive for 

imaging. Energy dispersive spectroscopy (EDS) was conducted using either EDAX CDU Leap Detector 

or IXRF Detector systems. 

3.2 Thickness Measurements 

Thickness measurements were taken at three locations along the cross section of each sample, using a 

LEO Supra 50VP scanning electron microscope (SEM) at GSFC, with an accelerating voltage of 20 kev. 

Samples for SEM thickness measurements were mounted in epoxy, and then SiC paper and diamond 

slurry were used to polish the cross sections. These thickness measurements were used to calculate atomic 

oxygen erosion yield and bend-testing strain. 

A Heidenhain MT12 drop gauge was used to measure the thickness of all regions at GRC. Thickness 

was measured in three locations on the grip area of each sample used for tensile testing (4 to 17 samples 

per region). These thickness measurements were used to determine cross sectional area individually for 

each sample when calculating tensile stress. The average thickness in each region was also used to 

calculate bend-test strain and atomic oxygen erosion yield. The drop gauge thickness measurements were 

used for these calculations because thickness was found to vary significantly within each region, and 

SEM measurements were only made at one or two locations within each region. 

3.3 Tensile Properties 

A DDL Inc. Model 200Q Electromechanical Test System was used to measure load and displacement 

during tensile elongation. From this data, the percent elongation at failure and ultimate tensile strength 

(UTS) were determined. The tensile samples were sectioned to the specifications defined in American 

Society for Testing and Materials (ASTM) Standard D-638 for Type V tensile specimens (Ref. 6). Care 

was taken to avoid cracks and impact sites. The initial sample length was set at 25.4 mm and the 

elongation rate was 12.7 mm/min. As an update from previously reported results of tensile testing SM4 

material (Ref. 7), tensile stress was calculated based on the thickness of each sample individually 

measured using the Heidenhain drop gauge. 

3.4 Bend-Testing 

Bend-testing was conducted on samples from select regions to assess tensile strain induced surface 

cracking. The bend-test procedure induced a surface tensile strain without adding an overall tensile load 

to the samples. The strain necessary to induce surface cracking was determined by bending 1.3u0.5 cm 

samples over successively smaller mandrels until cracking was observed. The test apparatus used a 

pliable platform to push the samples against the mandrel without applying overall tension, as illustrated in 

Figure 4. Each sample was placed with the space-exposed surface face-down on the platform and then 

bent into a U-shape, with the space exposed surface on the outside under tension and the back surface 

under compression. 
NASA/TM—2015-218476 6 

Figure 4.—Illustration of the bend-test 

configuration showing a cradle platform 

used to bend the sample around the 

mandrel. 

Optical microscopy was used to document surface features in the bend-test area prior to testing, and 

the same area was examined after bending around each mandrel to identify and document any induced 

surface cracks. The samples were examined at magnifications up to approximately 50X with an Olympus 

SMZ stereo-zoom optical microscope outfitted with a Canon digital camera. This procedure was 

continued with samples bent around successively smaller mandrels until cracks were visible or until the 

sample did not experience any cracking with the smallest mandrel, in which case the sample was recorded 

as having not cracked. Samples were bent over up to 69 mandrels with diameters ranging from 1.253 to 

0.052 cm in order of decreasing diameter (increasing tensile strain). The percent strain at failure, E, is 

given by 

 ̧ u100 1

·  ̈

©

§

 t d

t E (1) 

where t is the thickness of the sample and d is the diameter of the mandrel that caused cracking. Sample 

thickness t was taken as the average of all drop gauge thickness measurements from that region. Bendtesting is well suited to very brittle material because no overall tension is applied to the samples and small 

strains can be accurately measured. 

3.5 Density 

Density was measured using calibrated density gradient columns. The density gradient was 

established in a 50 mL buret using two solutions of carbon tetrachloride (CCl4, 1.594 g/cm3

) and 

bromoform (CHBr3, 2.899 g/cm3

). A less dense solution (1.803 g/cm3) was siphoned into a continuously 

stirred beaker initially containing only a more dense solution (2.299 g/cm3

), from which an equal flow 

was directed into the buret. Because the amount of lighter solution in the effluent from the beaker 

increased over time, solution density continuously decreased as a function of height. A quadratic 

calibration curve was developed for each column based on the equilibrium vertical position of three to 

four standards of known density (r0.0001 g/cm3

). Subsequently, density values were calculated from the 

vertical positions of small ( 2 mm) samples placed into the column and allowed to settle for 2 hr. Data 

were obtained from four separate density columns, with the density of each exposure region being 

measured at least twice. The very thin vapor-deposited aluminum backing on the FEP was previously 

determined to have no measureable effect on density measurements, and therefore was not removed prior 

to density measurements (Ref. 8). 
NASA/TM—2015-218476 7 

3.6 Optical and Thermal Properties 

Solar absorptance and thermal emittance measurements were taken at both GSFC and GRC. Cary 

5000 spectrophotometers equipped with Spectralon integrating spheres were used to measure total 

reflectance from 250 to 2500 nm at an 8° angle of incidence at both locations. Absorptivity data were 

integrated with respect to the air mass zero solar spectrum to obtain solar absorptance. Samples at GRC 

were measured with an Al spacer backed by a single layer of sample of the quilted Al/Kapton/Al inner 

film (Region 8.15) replicating the on-orbit configuration. 

Emissivity measurements were obtained at GSFC using a Gier-Dünkle DB-100 InfraRed 

Reflectometer following the ASTM E408-71 standard test method. The normal emittance (İn) of the 

surface was measured from 5 to 40 Pm while at room temperature. Three measurements were taken for 

each sample, with the exception of Bay 8 Region 3, which was highly delaminated, hence only one 

measurement was taken. Emissivity measurements were obtained at GRC using a Surface Optics 

Corporation Model SOC 400T Reflectometer. The SOC 400T measures the directional reflectance of 

surfaces over a large spectral range, 2 to 25 Pm, to obtain the directional thermal emittance over a large 

temperature range. Automatic integration of reflectivity data in the infrared with respect to blackbody 

curves is used to calculate total emittance for a selectable temperature range. Data were obtained at 293, 

313, and 393 K. The samples were placed in the sample holder face-down with an aluminum spacer and 

were backed by a sample of the quilted Al/Kapton/Al MLI inner layer. 

3.7 Atomic Oxygen (AO) Erosion Yield (Ey) 

The volume eroded per incident atomic oxygen atom (cm3/atom), or erosion yield (Ey), of the HST 

samples was calculated in two distinct ways, by mass loss and by thickness loss. Some of the HST 

material experienced a density change on orbit, so part of the observed thickness loss was caused by the 

density change, not atomic oxygen erosion. Therefore, while erosion yield can typically be calculated 

simply by dividing thickness loss by atomic oxygen fluence, in this case it was calculated using the 

following Equation (2) 

F

t t i  f U f Ui / Ey (2)

where ti is the thickness of pristine material, tf is the thickness of the space-exposed material, ȡi is the 

density of pristine material, ȡf is the density of the space-exposed material, and F is the atomic oxygen 

fluence (atom/cm2

). In this equation, the numerator represents the thickness loss that can be attributed to 

AO erosion. This calculation is an update from previously reported atomic oxygen erosion yield data for 

the Bay 5 and Bay 8, which was calculated assuming constant density (Ref. 7). 

To determine mass loss Ey, the mass of 12.7 mm diameter circular samples were measured using a 

Mettler M3 balance and compared to the mass of pristine material. Mass loss Ey was calculated using the 

following Equation (3) 

AF

M

Ui

' Ey (3)

where 'M is the measured mass loss and A is the exposed surface area of the sample. The result of this 

expression is directly comparable to Ey calculated using Equation (3). 
NASA/TM—2015-218476 8 

3.8 X-ray Photoelectron Spectroscopy (XPS) 

Samples were analyzed for elemental composition using a M-Probe X-Ray Photoelectron 

Spectrometer. Three locations on samples from various regions were run using a general survey scan, 

with a spot size of 800 Pm, to determine atomic percent composition. 

3.9 Static Heat Treatment 

The solar absorptance, thermal emittance, tensile properties, and density of heat treated samples were 

measured in addition to the as-retrieved material. Density measurements were taken after both 120 and 

200 qC heat treatments. A heat treatment at 200 qC was used for tensile strength samples and a heat 

treatment at 120 qC was used for solar absorptance and thermal emittance samples. Heating at 200 qC 

corresponds to the maximum temperature experienced on-orbit by very tightly curled insulation 

(Refs. 4 and 5). The 120 qC treatment was chosen to represent a more moderate on-orbit heating 

condition. Heating was conducted in air for 96 hr as prior data indicate a minimum of 72 hr is necessary 

for heating effects to be stabilized (Ref. 8). Bend-tests of the 200 qC heated HST material were planned 

but not completed because the insulation curled significantly on heating. 

4.0 Results and Discussion 

To retrieve the MLI blankets, astronaut John Grunsfeld unpeeled each blanket from its Velcro secured 

border, then folded the blankets and placed them into EVA storage bags (Fig. 2). Although great care was 

taken in retrieval planning and execution, the retrieval process did cause cracking at fold seams, 

introducing handling cracks that were not present on-orbit prior to retrieval. 

The Bay 5 and Bay 8 MLI blankets are shown in the on-orbit photos taken during SM4 in 

Figures 5(a) and (b), respectively. The Bay 5 MLI has several large cracks extending from the two 

rectangular radiator areas. The Bay 8 MLI also has several very large cracks, and one side of the largest 

crack on Bay 8 curled up into a cone-like roll, as seen in Figure 5(b). Also, the Bay 8 patches were highly 

degraded with the majority of the Al-FEP completely gone. Upon post-retrieval visual inspection, the 

Bay 8 MLI appeared significantly more degraded than the Bay 5 MLI, as the Bay 8 MLI had broken into 

many fragments after removing it from the EVA bag, unfolding it and laying it flat again. In fact, the 

Bay 8 Al-FEP outer layer needed to be reassembled post-flight like a jigsaw puzzle. 

 (a) (b)

Figure 5.—On-orbit photos of Equipment Bays 5 and 8 just prior to retrieval 

during SM4: (a) Bay 5, and (b) Bay 8 with large cracks circled. 
NASA/TM—2015-218476 9 

4.1 Exposure Regions 

The retrieved Bay 5 and Bay 8 MLI blankets contained a range of unique regions that were identified 

based on environmental exposure and/or physical appearance and assigned numbers. For example, the 

Bay 8 MLI that was not covered by patch material and was exposed to the space environment for the full 

19.1 years had areas with three distinct appearances: shiny areas (“nominal”), hazy-white areas and areas 

where the Al was delaminated. These regions were defined as Regions 8.1, 8.2 and 8.3 (R8.1, R8.2 and 

R8.3), respectively. Examples of these regions are provided in Figure 6. Also, as mentioned, one of the Bay 

8 outer layer Al-FEP cracked areas (not covered by patch material) curled up into a cone-like roll. In the 

curled region, the backside Al was exposed to the space environment. This material was heated to a higher 

temperature on-orbit than the nominal space-facing FEP because of the lower emittance of the aluminized 

surface. The curled area was divided into two regions: a tightly curled region (R8.7) and a loosely curled 

region (R8.8), as those two regions may have heated differently on-orbit. This curling was not observed at 

SM3B, so the insulation was in the curled state for no more than 7.1 years. Some areas of the patches 

installed during SM2 eventually degraded exposing the underlying MLI to the space environment once 

again as observed during SM3B and SM4. Therefore, the MLI under the patched areas is divided into two 

regions: the region patched during SM2 and exposed by SM3B (SM2-SM3B, patched 5.1 years, called 

R8.13), and the region patched during SM2 and still covered at SM4 (SM2-SM4, patched 12.2 years, called 

R8.11).

A total of 10 different regions were identified on Bay 5 (R1-R10), and 15 different regions were 

identified on Bay 8 (R1-R15) and 4 regions identified on the retrieved Patches (R16-R19). The regions 

selected for testing are defined in Table 1 and shown in Figure 7. During sectioning of samples from the 

blankets, care was taken to avoid any cracks or imperfections in the samples which would affect the tests. 

Figure 6.—Bay 8 nominal (shiny), hazy and delaminated MLI fragments. 

 

Nominal 

(Shiny) 

Delaminated 

Hazy 
NASA/TM—2015-218476 10 

TABLE 1.—HST SM4 BAY 5 AND BAY 8 REGION DEFINITIONS AND ENVIRONMENTAL EXPOSURES 

HST 

Material 

Bay.Region 

Description 

AO 

Fluence 

(E+20 

atom/cm2) 

Equivalent 

Sun Hours 

Max. 

OnOrbit 

Temp 

(qC)* 

Proton 

Fluence 

(E+10 

particles 

/cm2) 

Electron 

Fluence 

(E+13 

particles 

/cm2) 

X-ray 

Fluence 

(J/cm2) 

1-8 Å 0.5-4 

Å 

Pristine 

(P) 

Pristine 5 mil 

Al-FEP - - - - - - 

5.1 Nominal/Shiny 4.65 24,300 0 5.4 5.60 641.1 43.1 

5.4 Nominal/AlDelaminated 4.65 24,300 0 5.4 5.6 641.1 43.1 

8.1 Nominal/Shiny 4.28 89,300 40 5.4 5.6 641.1 43.1 

8.2 Nominal/White 

Hazy 4.28 89,300 40 5.4 5.6 641.1 43.1 

8.3 Nominal/AlDelaminated 4.28 89,300 40 5.4 5.6 641.1 43.1 

8.7**+ Cone/ 

Tight Curl 3.92 72,400 >120 

<200 4.4 4.5 519.8 34.9 

8.8**+ Cone/ 

Loose Curl 3.92 72,400 >120 <200 4.4 4.5 519.8 34.9 

8.11**+ Patched 

SM2-SM4 1.64 30,300 >40 <120 1.8 1.9 217.5 14.6 

8.13**+ Patched 

SM2-SM3B 2.69 70,800 40 4.3 4.4 508.3 34.2 

8.15 

Protected 

Inner Layer 

(Al/Kapton/Al) 

- - - - - - - 

* All HST material had 110,000 total thermal cycles and a minimum temperature of –175 qC 

** Proton, electron, and x-ray fluences for these regions were calculated assuming constant flux over the exposure period of 

the insulation. 

+ Temperature ranges for these regions were determined based on density measurements of as-retrieved and 

ground-heated samples (see section 4.5). 

 (a) (b)

Figure 7.—Photographs of Bay 5 and Bay 8 with exposure regions identified: (a) Bay 5, and (b) Reassembled 

Bay 8 (red regions were tested, blue regions were not). 
NASA/TM—2015-218476 11 

Figure 8.—Comparison of Bay 5 and Bay 8 shiny and delaminated regions: (a) A 

section of material cut from Bay 5, and (b) A cracked fragment of Bay 8 material 

(back lit photos). 

The delaminated regions of Bay 5 and Bay 8 have very different physical appearances; therefore the 

delamination may not have been caused by the same factor. Bay 5 delaminated samples (an example is 

provided in Fig. 8) have a checkerboard-like pattern of delaminated areas on a surface with predominately 

retained Al backing. There is no tendency for more Al to flake off. In contrast, Bay 8 delaminated 

samples (an example is provided in Fig. 8) have small patches of Al distributed throughout a mostly 

delaminated surface and additional Al easily spalls off from the FEP on handling the samples. 

4.2 Scanning Electron Microscopy 

Figures 9 and 10 compare the space-exposed (FEP) surfaces of Bay 8 Region1 (nominal/shiny) and 

Bay 8 Region 2 (nominal/hazy) using the Hitachi S-4700. Atomic oxygen erosion has resulted in distinct 

textures in these two regions. The shiny material has fine, rill-like directional texturing, whereas the hazy 

material has larger rounded features. The difference in texture may be attributed to the hazy region being 

located near the top edge of the bottom patch, where it may have received additional shadowing and/or 

scattering of AO from the patch at various times in the orbit. EDS spectra of Regions 1 and 2 are shown 

in Figures 11(a) and (b), respectively. The spectra indicate that only C and F peaks, and very small Au 

and Pd peaks from the Au-Pd conductive coating, are present. There is no presence of Si, hence no 

indication of silicone contamination in the nominal/shiny region, nor in the hazy region. However, a 

barely detectable Si peak was detected on a Bay 8 Region 1 sample imaged with the Hitachi S-3500. 

Figure 12 shows the space-exposed (FEP) surface of Bay 8 Region 3 (delaminated). The FEP is clearly 

cracked in the delaminated region, and the texture appears to be less pronounced as compared to 

Regions 8.1 and 8.2. It is not known if the cracking occurred on-orbit or due to retrieval handling. The 

space-exposed FEP from Bay 8 Regions 7 and 11, and Bay 5 Region 1, are shown in Figures 13 to 15, 

respectively. Bay 8 Region 7, which became tightly curled on-orbit, looks similar in texture to Bay 8 

Region 1 surfaces, with perhaps slightly less texture. The Bay 8 Region 11 surface, which was patched 

1 mm

2 cm

1 mm

Nominal Delaminated

(a)

(b)
NASA/TM—2015-218476 12 

between SM2-SM4, has more rounded rill features, possibly due to receiving scattered AO from the 

patch. Both these surfaces had barely detectable Si peaks. As can be seen in Figure 15, the Bay 5 

Region 1 surface has a different appearance than the Bay 8 surfaces. Whereas the Bay 8 surfaces eroded 

with a texture consistent with receiving grazing, or sweeping, AO exposure, the Bay 5 Region 1 surface 

showed little evidence of erosion with no notable erosion texture. This is consistent with thickness 

measurements, discussed later, which indicate that the Bay 5 material was eroded the least. The Bay 5 

surface did show a larger Si EDS peak, indicating that there is some silicone contamination on this 

surface. This is consistent with the XPS analyses reported later. 

 (a) (b)

Figure 9.—Space exposed FEP surface of Region 8.1 (nominal/shiny): (a) 1KX magnification 

(45q tilt), and (b) 5KX magnification (45q tilt). 

 (a) (b) 

Figure 10.—Space exposed FEP surface of Region 8.2 (hazy): (a) 1KX magnification (45q tilt), 

and (b) 5KX magnification (45q tilt). 
NASA/TM—2015-218476 13 

 (a) (b) 

Figure 11.—EDS spectra and corresponding SEM image areas (0q tilt) of space exposed Teflon FEP surfaces: 

(a) Region 8.1 Nominal/Shiny, and (b) Region 8.2 Hazy. 

(a) (b) 

Figure 12.—Space exposed FEP surface of Region 8.3 (delaminated): (a) 1KX magnification 

(0q tilt), and (b) 5KX magnification (0q tilt). 

 (a) (b) 

Figure 13.—Space exposed FEP surface HST SM4 Region 8.7 (tightly curled region): (a) 1KX 

magnification (45q tilt), and (b) 2.5 KX magnification (45q tilt). 

F F

C

C

Au Pd Au Pd
NASA/TM—2015-218476 14 

 (a) (b) 

Figure 14.—Space exposed FEP surface HST SM4 Region 8.11 (patched SM2-SM4): (a) 1KX 

magnification (45q tilt), and (b) 2.5 kX magnification (45q tilt). 

 (a) (b) 

Figure 15.—Space exposed FEP surface of Region 5.1 (nominal/shiny): (a) 1KX magnification 

(0q tilt), and (b) 1KX magnification (45q tilt). 

Figures 16 to 18 compare the aluminum surfaces of pristine Al-FEP, Bay 8 Region1 (nominal/shiny) 

and Bay 8 Region 2 (nominal/hazy) materials, respectively. The aluminum surface of the pristine Al-FEP 

is smooth with no cracks, as shown in Figure 16(a). At very high magnification (100KX) a grain structure 

can be seen as shown in Figure 16(b). As can be seen in Figures 17(a) and 18(a), the aluminum layer of 

the HST material is cracked, with more cracks (smaller islands) present in the hazy region (18a). Cracking 

of the Al layer can be attributed to the stress caused during thermal cycling because of the difference in 

the coefficient of expansion between Al and FEP (Ref. 4). Another interesting observation is the apparent 

growth of microscopic crystals in the aluminum surface, as shown in Figures 17(b) and 18(b), as 

compared to the pristine Al-FEP in Figure 16(b). 

Images of the aluminum backing of Bay 8 Region 7, Bay 8 Region 11 and Bay 5 Region 1 are shown 

in Figures 19(a) to (c), respectively. The Bay 8 Region 7 surface has more mud-tiling than Bay 7 

Region 1, and appears similar to the extent of mud-tiling of Bay 8 Region 2. The Bay 8 Region 11 

surface, patched between SM2 and SM4, had the most extensive, finiest, mud-tiling. It was determined 

that this region received a higher temperature on-orbit with the patch in place. Bay 8 Region 7 also 

experienced a higher temperature on-orbit, but the curling was estimated to begin after SM3B, and hence 

had excessiving heating for a maximum of 7.1 years, and the tightly curled region would have been 

heated the least (at the bottom of the curled region). Bay 8 Region 11 would have received intense heating 

for approximately 12.2 years. The Al backing of the Bay 5 Region 1 material had more mud-tiling 

compared to Bay 8 Region 1. It appeared similar to the extent of mud-tiling of Bay 8 Region 2 or Bay 8 

Region 7. The maximum on-orbit temperature of the Bay 5 Region 1 MLI was estimated to be less (0°) 

than that of the Bay 8 Region 1 MLI (40°), so the extent of mud-tiling is not easily explained. 
NASA/TM—2015-218476 15 

 (a) (b) 

Figure 16.—Aluminum backing of pristine Al-FEP: (a) 500X magnification (0q tilt), and (b) 100KX 

magnification (0q tilt). 

 (a) (b) 

Figure 17.—Aluminum backing of HST SM4 Region 8.1 (nominal/shiny): (a) 250X magnification 

(0q tilt), and (b) 100KX magnification (0q tilt). 

 (a) (b) 

Figure 18.—Aluminum backing of HST SM4 Region 8.2 (hazy): (a) 250X magnification (0q tilt), 

and (b) 100KX magnification (0q tilt). 
NASA/TM—2015-218476 16 

 (a) (b)

(c) 

Figure 19.—Aluminum backing (250X, 0q tilt) of: (a) HST SM4 Region 8.7 (tightly curled region), 

(b) HST SM4 Region 8.11 (patched SM2-SM4), and (c) HST SM4 Region 5.1 (nominal/shiny). 

 
NASA/TM—2015-218476 17 

Several micrometeoroid and orbital debris (MMOD) impact sites were observed on the HST material. A 

typical impact site is shown in Figure 20, with fine cracking of the FEP layer visible near the impact site. 

SEM micrographs show very clear differences in the morphologies of on-orbit formed cracks 

(Fig. 21) and post-flight handling formed cracks (Fig. 22). Handling cracks have fibrous features typical 

of ductile failure. However, the fracture surfaces of on-orbit cracks were very smooth. Similar 

observations were made on material retrieved during SM2, and this supports a slow crack growth 

mechanism in which on-orbit fractures propagate slowly under relatively low stresses, influenced by 

radiation and thermal cycling (Ref. 4). 

 Figure 20.—An MMOD impact site in Bay 8 Region 1. 

 (a) (b) 

Figure 21.—Edge on view of an on-orbit formed crack in HST SM4 Region 8.1 FEP (nominal/shiny) 

showing a smooth texture: (a) 1KX magnification, and (b) 20KX magnification. 

 (a) (b) 

Figure 22.—Edge on view of a post-flight handling formed crack in HST SM4 Region 8.1 FEP 

(nominal/shiny) showing a fibrous texture: (a) 1KX magnification, and (b) 20KX magnification. 
NASA/TM—2015-218476 18 

4.3 Thickness Measurements 

Thickness measurements from the SEM and drop gauge were largely similar, as shown in Table 2. 

However, thickness measurements using the drop gauge showed significant variability across different 

locations in a given region. The drop gauge measurements are better represent each region’s over all 

thickness because thickness measurements were taken at many more locations within each region, as 

indicated in the table under the number of samples tested. 

4.4 Tensile Properties and Bend-Test Results 

The average ultimate tensile strength (UTS) and percent elongation at failure of the HST and pristine 

Al-FEP are provided in Table 3 and graphed in Figure 23. Because of the degree of embrittlement of the 

HST samples, many broke while being punched out or handled prior to testing. Others were tested but 

failed at pre-existing cracks, causing a premature break. Data from premature breaks were not included in 

the average. Five samples from Region 8.7 (tightly curled region) were tested, and all broke prematurely, 

thus no data are reported. For all other regions, usable data was obtained from 3 to 8 samples depending 

on material availability. 

All HST material was extremely embrittled. Elongation at failure was reduced from 255 percent for 

pristine material to 8.5 percent for Bay 5 nominal/shiny material and to 1.5 percent for Bay 8 

nominal/shiny material. The Bay 8 Al-FEP, which had a significantly higher solar exposure and on-orbit 

thermal cycling temperature, was found to fracture very easily with handling, like thin brittle glass. All 

regions also experienced large decreases in UTS, from 22.4 MPa for pristine Al-FEP to 13.3 and 5.9 MPa 

for Bay 5 and Bay 8 nominal/shiny material, respectively. Although the standard deviation is large 

compared to some of the absolute values, both the UTS and the elongation at failure appear to correlate 

with the amount of solar exposure received. For example, the solar facing Bay 8 material is more 

embrittled than the solar grazing Bay 5 material, and the elongation at failure of Bay 8 material is 

proportional to the amount of time that it was patched. 

The decrease in both elongation and UTS indicates that the material has undergone significant 

scission of the polymer bonds. If the material had cross-linked with space irradiation, then the material 

should have experienced increased strength with the decreased elongation, which it did not experience. 

TABLE 2.—COMPARISON OF DROP GAUGE AND SEM THICKNESS MEASUREMENTS* 

HST 

Material 

Bay.Region 

Drop Gauge (GRC) SEM (GSFC) 

No. of 

Samples 

Avg. 

Thickness 

(ȝm) 

Std. 

Dev. 

(ȝm) 

 No. of 

Samples 

Avg. 

Thickness 

(ȝm) 

Std. 

Dev. 

(ȝm) 

Pristine 5 125.6 1.2 1 130.77 3.15 

5.1 8 120.6 5.8 1 124.13 3.06 

5.4 6 104.4 15.4 0 - - 

8.1 17 75.9 3.3 1 72.19 0.18 

8.2 8 88.1 15.7 1 88.61 0.36 

8.3 9 73.6 4.1 2 75.40 2.01 

8.7 5 77.1 3.5 1 84.49 1.13 

8.8 5 76.8 3.0 3 83.00 3.03 

8.11 4 109.3 1.8 2 110.58 2.25 

8.13 5 82.4 3.8 1 94.80 3.08 

*For both techniques, thickness measurements were made three times on each sample. 
NASA/TM—2015-218476 19 

TABLE 3.—TENSILE PROPERTIES OF PRISTINE AND HST Al-FEP 

HST Material 

Bay.Region 

No. of 

Samples 

Tested 

No. of 

Cracked 

Samples 

Excluded 

Avg. 

Thickness 

(ȝm) 

Ultimate Tensile 

Strength Elongation at Failure 

Avg. 

(MPa) Std. Dev. Avg. (%) Std. Dev. 

Pristine (P) 13 0 125.6 22.4 0.8 254.7 9.7 

P-200 °C 8 0 125.6 21.0 1.6 253.6 20.3 

5.1 8 1 120.6 13.3 1.0 8.5 8.9 

5.1-200 °C 8 1 123.0 9.4 1.9 2.3 0.3 

5.4 6 2 104.4 9.5 2.0 2.3 0.4 

8.1 17 5 75.9 5.9 1.5 1.4 0.3 

8.2 8 3 88.1 8.9 3.5 2.7 0.9 

8.3 9 4 73.6 8.4 3.5 2.4 1.1 

8.7 5 5* - - - - - 

8.8 5 2 76.8 6.0 3.2 1.8 1.4 

8.11 4 0 109.3 9.9 1.2 2.6 0.4 

8.13 5 1 82.4 6.9 0.9 1.7 0.2 

*All five samples broke prematurely, hence there is no tensile data

 
NASA/TM—2015-218476 20 

(a) 

(b) 

Figure 23.—Tensile properties of pristine and HST Al-FEP: (a) Average UTS and 

(b) Average elongation at failure.

22.4 21.0

13.3

9.4 9.5

5.9

8.9

8.4

6.0

9.9

6.9

0

5

10

15

20

25

UTS (MPa)

255.0

253.6

8.5

2.3 2.3 1.4 2.7 2.4 1.8 2.6 1.7

0

50

100

150

200

250

300

Elongation at Failure (%)
NASA/TM—2015-218476 21 

TABLE 4.—COMPARISON OF ELONGATION AT FAILURE (E) FOR BEND AND TENSILE TESTS 

HST SM4 

Material 

Bay.Region 

Bend-Test Tensile Test 

Number 

of Trials 

Avg. E 

(percent) 

Number 

of Trials 

Avg. E 

(percent) 

Pristine 1 No Failure* 8 255 

5.1 3 6.69 7 8.51 

8.1 3 0.82 12 1.51 

*The maximum strain tested was 19.7 percent. 

 (a) (b)

Figure 24.—A sample from Bay 5 Region 1 (a) before and (b) after bend-testing with the bend-test-induced 

crack circled. 

Heating pristine Al-FEP marginally decreased its strength and had no effect on elongation. The Bay 8 

MLI was too embrittled to conduct studies to assess the effect of heating, but heating Bay 5 Region 1 

material at 200 °C for 96 hr further reduced both the elongation and strength of the samples. This supports 

the mechanism proposed for degradation in which solar radiation and solar heating play synergistic roles. 

To determine whether space-exposed Al-FEP continues to degrade after being returned to Earth, 

11 samples from Bay 8 Region 1 were tested one year after recovery in 2010 and 6 samples were tested 

four years after recovery in 2013. There was no significant difference between these samples, so the 

results were combined. 

The elongation at failure measured by bend-testing is shown and compared to the results of the tensile 

data in Table 4. Because bend-testing is designed for brittle samples, the maximum strain applied in the 

bend-test was insufficient to cause failure in pristine material. The elongation at failure measured by 

bend-testing qualitatively agrees with elongation at failure measured by tensile testing, with Bay 8 

nominal material much more embrittled than Bay 5 nominal material. Therefore, bend-test results also 

support the conclusions from tensile testing. However, elongation at failure measured by bend-testing is 

slightly lower than elongation at failure measured by tensile testing. This discrepancy likely occurs 

because failure is defined as initial cracking for bend-testing, but it is defined as a complete break for 

tensile testing. A representative image of bend-test induced cracking failure is shown in Figure 24. Bendtests to verify the effect of heating on Bay 5 Region 1 and Bay 8 Region 1 were not run because all 

samples curled significantly while being heated, which would affect the strains induced during testing.

Due to the brittle properties of the MLI and the high incidence of cracking, the data for Regions 5.1, 

8.1, 8.11 and pristine Al-FEP were supplemented with microscale mechanical testing at GSFC. 

Microscale mechanical testing uses very small specimen sizes and offers the advantage that pre-existing 

macroscale cracks in the material do not promote premature failure of the specimen. These test specimens 

could experience strains and ultimate loads reflective of the environmental degradation seen in the 

material properties; however, the microscale test results are less conservative than traditional mechanical 

testing in that they do not account for the macroscale defects such as tears or cracks that may limit the 

material on the bulk scale. Detailed experimental methods and results for these microscale mechanical 

0.5 mm 0.5 mm
NASA/TM—2015-218476 22 

tests have been previously reported by Jones et al. (Ref. 9). When comparing results between the two 

tensile test methods, the magnitude of the UTS and the elongation at failure values for the microscale tests 

differ from those of the traditional tensile tests due to the reasons stated above; however, the general trend 

with regard to strength in the different exposure regions is consistent between the two test methods. 

4.5 Density 

The densities for all measured regions of HST Al-FEP and pristine material are provided in Table 5 

and graphed in Figure 25. With the exceptions of R8.7 (tightly curled), R8.8 (curled), R8.11 (patched), 

the density of as-retrieved material did not vary significantly from pristine material (2.140 g/cm3

). Most 

exposure regions (besides R8.7, R8.8 and R8.11) only reached an estimated maximum temperature of 40 

qC, which is too low to allow for crystallization, so no change in density was observed. A large change in 

density (to 2.174 g/cm3

) was observed for the curled regions, which has previously been shown to be a 

result of the higher on-orbit temperatures experienced in these regions (Refs. 8 and 10). At elevated 

temperatures chain mobility increases, allowing more polymer chains to align into the compact, 

crystalline state, increasing both percent crystallinity and density, as further discussed (Refs. 8 and 10). 

The density of the Bay 8 Region 11 material was only marginally higher (2.146 g/cm3

) than pristine 

material, indicating that this region may have experienced more moderate heating. These samples were 

taken from directly beneath the Velcro border of the patch, which, as will be discussed, was observed to 

have contamination and an increase in solar absorptance. This would cause a higher temperature maximum 

during thermal cycling, and may account for the density increase. 

Heating at 120 qC for 96 hr resulted in substantial increases in density (~0.015 g/cm3

) for most spaceexposed material, but did not have a significant effect on pristine material or material from the curled 

regions, and the increase was much smaller (~0.004 g/cm3

) for the patched region. 

The increases in density are again explained by the increase in crystallinity that occurs when polymer 

chains become more mobile while heated. The glass transition temperature (Tg) (D relaxation) of pristine 

Teflon FEP is listed from |83 to 150 qC in the literature, depending on hexafluoropropylene (HFP) 

content. Eby and Wilson report transition temperatures for FEP with densities similar to the pristine FEP 

examined in this report to be |150 and |127 qC for 10.7 and 17.7 mol% HFP, respectively (Ref. 11). 

Commercially available FEP is reported to be 20 mol% HFP (Don Farrelly, DuPont, personal 

communication (1999)), which would indicate that the Tg for pristine FEP would be close to 120 qC 

TABLE 5.—DENSITY OF AS-RECEIVED, 120 AND 200 qC HEATED SAMPLES 

HST Material 

Bay.Region 

No. of 

Samples 

As Retrieved 120 qC Heated 200 qC Heated 

Density 

(g/cm3) Std. Dev. Density 

(g/cm3) Std. Dev. Density (g/cm3) Std. Dev. 

Pristine Al-FEP 4 2.1397 0.0028 2.1415 0.0020 2.1597 0.0037 

5.1 3 2.1361 0.0035 2.1485 0.0009 2.1840 0.0061 

5.4 3 2.1438 0.0014 2.1509 0.0013 2.1823 0.0126 

8.1 3 2.1378 0.0025 2.1545 0.0012 2.1970 0.0095 

8.2 3 2.1403 0.0026 2.1558 0.0024 2.1901 0.0058 

8.3 2 2.1408 0.0013 2.1539 0.0016 2.1863 0.0008 

8.7 3 2.1714 0.0080 2.1718 0.0078 2.1879 0.0058 

8.8 2 2.1744 0.0061 2.1729 0.0062 2.1852 0.0011 

8.11 3 2.1458 0.0005 2.1500 0.0004 2.1809 0.0047 

8.13 3 2.1385 0.0024 2.1523 0.0010 2.1875 0.0031 
NASA/TM—2015-218476 23 

Figure 25.—Density of as-received, 120 and 200 qC heated samples. Error 

bars indicate the standard deviation. 

based on the Eby study. In a study by de Groh and Martin, the change-of-slope temperature for density 

versus vacuum heat treatment temperature for pristine FEP was found to be a126 °C (Ref. 12). Therefore, 

heating to 120 qC is likely to be insufficient to induce a change in crystallinity of non-irradiated FEP, as 

indicated in the de Groh and Martin study (Ref. 12).The density increase for HST material on heating at 

120 qC indicates that solar radiation induced bond scission, increasing the mobility of polymer chains. 

The curled regions were already heated above 120 qC on orbit, so further heating at this temperature did 

not have an effect on the density of these regions. The slight density increase observed for Bay 8 

Region 11 indicates that although it experienced an increase in maximum temperature, this region did not 

reach 120 qC on orbit. 

Heating at 200 qC for 96 hr resulted in a substantial increase in density for both pristine Al-FEP and 

all regions of space-exposed material as compared to both pristine and 120 qC heated material. It should 

be noted that de Groh et al. found the on-set of melt of pristine 5 mil Al-FEP to be 244.6r3.0 °C based on 

DSC studies (Ref. 10), and the melt temperature was 268 °C, which is consistent with values reported in 

the literature. The density of 200 qC heated pristine material (2.158 g/cm3

) was significantly less than the 

heated material from the HST (2.174 to 2.190 g/cm3

), again indicating that space irradiation-induced 

scission of bonds increased mobility and crystallization on heating as compared to non-irradiated FEP. Of 

the HST material, the largest density increase occurred for the Bay 8 nominal material, and the smallest 

increases occurred for the Bay 5 nominal material, demonstrating that increased solar exposure leads to 

increased scission. Because there was a significant increase in density for the curled regions, it can be 

inferred that they did not reach 200 qC on orbit. 

In the 2001 de Groh et al. study, x-ray diffraction (XRD) data was taken on pristine and retrieved 

HST FEP from the first three servicing missions for as-retrieved material and after heating at 200 °C. 

The as-retrieved space exposed FEP was found to be only very slightly higher in crystallinity 

(28 to 32 percent) than pristine FEP (28 percent), consistent with density data. With the exception being 

the servicing mission 2 (SM2) material, which has curled up tightly and was heated to 200 °C on-orbit. 

This material displayed both a higher density and XRD crystallinity (46 to 47 percent). The study also 

showed that the crystallinity of 200 °C heated samples increases, with greater increases occurring for the 

space exposed FEP after 200 °C. These results are consistent with the density data in this study, and 

provide further evidence for space irradiation-induced scission of bonds leading to increased mobility and 

crystallization on heating as compared to non-irradiated FEP. 

2.100

2.120

2.140

2.160

2.180

2.200

2.220

Pristine 5.1 5.4 8.1 8.2 8.3 8.7 8.8 8.11 8.13

Density (g/cm3)

As-Retrieved Heated 120 °C Heated 200 °C
NASA/TM—2015-218476 24 

4.6 Optical and Thermal Properties 

The as-retrieved and 120 °C-heated solar absorptance values obtained at GSFC and GRC are 

provided in Table 6 and graphed in Figure 26. It should be noted that the pristine Al-FEP absorptance 

obtained at GRC (0.15) was slightly higher than the value obtained at GSFC (0.13). This is attributed to 

calibration differences between the two instruments; hence comparisons should only be made between 

data taken on the same instrument. As-retrieved samples from both GSFC and GRC experienced an 

increase in solar absorptance compared to pristine Al-FEP, with the exception of Bay 8 Region 3 

(delaminated region), which was measured at GSFC without an inner layer. The GSFC absorptance of 

Bay 5 was slightly higher (0.16) than pristine Al-FEP (0.13), while Bay 8 had a wide range of 

absorptance values, with the greatest absorptance for Bay 8 Region 13 (0.27). This is the region that was 

patched between SM2 and SM3B, and then exposed again to space. The data trends for absorptance 

changes are consistent for both GSFC and GRC, with the Bay 8 hazy-white region (Region 2) and the 

patched regions (Regions 11 and 13) having the greatest increases in solar absorptance. 

As stated previously, the GRC samples were measured with a piece of the inner layer embossed 

Al/Kapton/Al to best replicate the optical properties in space. As might be expected, the reflectance of the 

inner layer was found to have the biggest impact on the solar absorptance of the delaminated Region (R8.3), 

increasing the absorptance significantly as compared to the GSFC data without the inner layer. 

Heating was found to increase the solar absorptance of the pristine Al-FEP by 0.09. The Bay 8 

Al-FEP experienced similar increases due to heating (0.09-0.11), with the regions that were covered by 

patches increasing even more (0.12-0.14). Although the Bay 5 thermal cycled to a lower maximum 

temperature on-orbit (0 °C) than Bay 8 (40 °C), the Bay 5 Al-FEP experienced very small increases in 

absorptance with heating as well. The reason for this is currently unknown. 

The as-retrieved and heated thermal emittance values obtained at GSFC and GRC are provided in 

Table 7 and graphed in Figure 27. The GSFC thermal emittance values were found to decrease for all 

samples except Bay 5 Region 1 and Bay 8 Region 11 (the region patched between SM2 and SM4). 

Emittance loss is typically associated with thickness loss of FEP, but for Bay 8 Regions 7 and 8 (the 

curled regions), samples may have experienced additional emittance change due to excessive heating onorbit. Similar to the GSFC emittance values, the GRC emittance was found to decrease for all samples 

except Bay 5 Region 1 (also Bay 5 Region 4) and Bay 8 Region 11, the region patched between SM2 and 

SM4. Heating had little impact on the emittance values of the retrieved HST MLI. 

TABLE 6.—SOLAR ABSORPTANCE VALUES FOR HST AND PRISTINE Al-FEP 

HST Material 

Bay.Region 

GSFC (Without Inner Layer) GRC (With Inner Layer) 

No. of 

Samples 

As-Retrieved Std. Dev. No. of 

Samples 

As-Retrieved Std. Dev. Heated Std. Dev.

Pristine 3 0.13 - 1 0.154 0.000 0.246 0.016 

5.1 3 0.16 - 2 0.217 0.021 0.225 0.040 

5.4 - - - 2 0.253 0.043 0.299 0.063 

8.1 3 0.18 0.02 2 0.189 0.038 0.301 0.035 

8.2 3 0.22 0.02 1 0.249 - 0.337 - 

8.3 1 0.13 - 1 0.214 - 0.326 - 

8.7 3 0.19 0.01 - - - - - 

8.8 3 0.23 0.01 - - - - - 

8.11 3 0.26 0.05 1 0.240 - 0.359 - 

8.13 3 0.27 0.01 1 0.213 - 0.349 - 

8.15 - - - 1 0.141 - 0.301 - 
NASA/TM—2015-218476 25 

Figure 26.—As-retrieved and 120 qC heated solar absortpance values. 

TABLE 7.—THERMAL EMITTANCE VALUES FOR HST AND PRISTINE Al-FEP 

HST Material 

Bay.Region 

GSFC GRC 

AsRetrieved Std. 

Dev. 

AsRetrieved Std. 

Dev. Heated Std. Dev. 

Change from 

heating 

Pristine 0.79 0.01 0.879 0.890 0.011 

5.1 0.80 0.01 0.877 0.008 0.895 0.001 0.018 

5.4   0.875 0.021 0.886 0.018 0.011 

8.1 0.75 0.01 0.831 0.006 0.836 0.009 0.004 

8.2 0.76 0.01 0.841 - 0.861 - 0.020 

8.3 0.74 0.800 - 0.805 - 0.005 

8.7 0.75 0.02 - - - - - 

8.8 0.73 0.01 - - - - - 

8.11 0.79 0.01 0.868 - 0.874 - 0.006 

8.13 0.77 - 0.830 - 0.851 - 0.021 

8.15 Ͳ Ͳ  0.030 - 0.065 - 0.035 

0.0

0.1

0.2

0.3

0.4

Pristine 5.1 5.4 8.1 8.2 8.3 8.7 8.8 8.11 8.13 8.15

Solar Absorptance (

Ƚs)

As-Retrieved (GSFC) As-Retrieved (GRC) Heated (GRC)
NASA/TM—2015-218476 26 

Figure 27.—As-retrieved and heated emittance values. 

TABLE 8.—XPS RESULTS FOR ATOMIC PERCENT COMPOSITION 

HST SM4 

Material 

Bay.Region 

Atomic % 

F/C Ratio C N O F Si 

Pristine 33.27 0.00 0.00 66.73 0.00 2.01 

5.1 19.44 1.12 36.36 22.56 20.52 1.16 

8.1 39.19 2.31 6.35 50.87 1.29 1.30 

8.2 39.50 2.12 5.43 52.55 0.41 1.33 

8.11 59.41 4.10 18.70 15.14 2.65 0.25 

4.7 X-Ray Photoelectron Spectroscopy (XPS) 

The XPS analyses for pristine Al-FEP and Regions 5.1, 8.1, 8.2, and 8.11 are provided in Table 8. As 

expected, the pristine Al-FEP is comprised of C (33.4 percent) and F (66.7 percent). The Bay 8 nominal 

regions (1 and 2) were similar with an increase in C (|39 percent), a decrease in F (|51 percent) and the 

presence of O (|6 percent), N (|2 percent) and Si (0.4 to 1.2 percent). Bay 8 Region 11, which was 

patched from SM2-SM4 had a high concentration of C, possibly due to contamination from the Velcro 

used to attach the patch. Surprisingly, Bay 5 was found to have a significant amount of silicone 

contamination with the presence of 20 percent Si and 36 percent O. The back of the solar arrays are 

coated with DC 93-500 silicone and because Bay 5 faces towards the +V2 solar array direction, the 

DC 93-500 is most likely the source of contamination. As only one area of Bay 5 was tested, additional 

tests could determine if silicone contamination is wide spread or localized on the blanket. It would be 

useful to determine if this is the case, as silicone contamination could impact Ey values along with optical 

and thermal results. 

 

0.0

0.2

0.4

0.6

0.8

1.0

Pristine 5.1 5.4 8.1 8.2 8.3 8.7 8.8 8.11 8.13 8.15

Thermal Emittance (İ)

As-Retrieved (GSFC) As-Retrieved (GRC) Heated (GRC)
NASA/TM—2015-218476 27 

In a 2008 study by de Groh et al., samples sectioned from Al-FEP circular thermal shields covering 

the bi-stem booms on the second set of HST solar arrays (retrieved after 8.25 years of space exposure) 

were examined with scanning electron microscopy for surface morphology and with energy dispersive 

spectroscopy (EDS) for surface chemistry. Pristine Al-FEP and solar-facing and anti-solar-facing thermal 

shield surfaces did not have any evidence of contamination, with only C and F peaks observed. However, 

a crazed surface texture was observed on one of the solar-grazing samples. The EDS data indicated the 

crazed texture is due to silicone contamination. Hence, the Bay 5 silicone contamination is consistent with 

contamination found on the solar-grazing surface of the retrieved solar array thermal shields, which is 

probably caused by contamination from the DC 93-500 silicone coating on the anti-solar side of the 

solar arrays. 

4.8 Atomic Oxygen Erosion Yield (Ey) 

Atomic oxygen Ey values determined through mass or thickness loss for the probable AO fluences 

are provided in Table 9. Erosion yield values were determined for Bay 8 Regions 7, 8, 11, and 13 based 

on the fluence only during time periods when these regions were exposed. Table 9 also provides the 

average Al-FEP thickness and the ratio of thickness loss (TL) Ey to mass loss (ML) Ey. The two Ey 

methods gave generally consistent results with each other. 

The ML based Ey value of the HST SM4 Bay 8 nominal shiny FEP (Ey = 1.17u10–23 cm3

/atom) was 

found to be an order of magnitude greater than for the Bay 5 nominal shiny FEP (Ey = 1.27u10–24

cm3

/atom). This is attributed to the significantly higher dose of solar radiation, combined with the higher 

on-orbit temperature, for the solar facing blanket. It should be noted that silicone contamination on Bay 5 

may have contributed somewhat to the decreased the rate of erosion. But, it is not expected that the 

silicone contamination would contribute an order of magnitude difference in Ey. A thick, protective layer 

of silicone typically is seen as a crazed surface layer, which was not at all evident in the SEM images of 

Bay 5 MLI (Ref. 13). 

TABLE 9.—ATOMIC OXYGEN EROSION YIELD COMPARISON

HST 

Material 

Bay.Region 

AO 

Fluence 

× 10-20 

(atom/cm2) 

Avg. 

Mass* 

(g) 

Mass Loss Ey 

(cm3/atom) 

Avg. 

Thickness 

(cm) 

Thickness 

Loss+

Ey (cm3/atom) 

Ratio of Ey 

(TL)/ Ey 

(ML) 

Pristine - 36.837 - 0.01256 - - 

5.1 4.65 35.108 1.27E-24 0.01206 1.12E-24 0.88 

8.1 4.28 22.158 1.17E-23 0.00759 1.16E-23 0.99 

8.2 4.28 26.544 8.22E-24 0.00881 8.72E-24 1.06 

8.3 4.28 22.810 1.12E-23 0.00736 1.21E-23 1.08 

8.7 3.92 23.630 1.15E-23 0.00771 1.21E-23 1.05 

8.8 3.92 23.419 1.17E-23 0.00768 1.21E-23 1.03 

8.11 1.64 31.654 1.08E-23 0.01093 9.7E-24 0.90 

8.13 2.69 25.282 1.49E-23 0.00824 1.60E-23 1.08 

 *1.27 cm diameter sample

+ Drop-Gauge Thickness

 
NASA/TM—2015-218476 28 

The Ey data provide evidence that solar exposure plays a significant role in the AO erosion of FEP. 

Erosion yield values were similar across all regions of Bay 8, with the exception of Region 2, which had a 

lower Ey (mass loss Ey = 8.22u10–24 cm3

/atom). No silicone contamination was found in this region and 

the reduction in Ey may be due to shadowing from adjacent patch material. The thickness reduction of 

HST materials comes almost entirely from erosion, with only at most 3 percent (for Bay 8 Region 8) of 

the total thickness change being caused by densification. 

The Ey values of the HST SM4 FEP were found to be orders of magnitude greater than those 

determined from shuttle flight experiments such as the Evaluation of Oxygen Interaction with 

Materials III, reported from 5.0u10–26 cm3

/atom (Ref. 14) to 1.8u10–25 cm3/atom (Ref. 15), the Long 

Duration Exposure Facility, determined to be 3.37u10–25 cm3

/atom for ram facing surfaces (Ref. 16) and 

various ram facing Materials International Space Station Experiment (MISSE) samples, determined to 

range from 1.28u10–25 cm3

/atom for MISSE 4 to 2.11u10–25 cm3/atom for MISSE 2 (Refs. 17 and 18). 

Again, this is thought to be attributed to the effects caused by the significantly higher solar exposure of 

the HST materials. 

The Ey values of Teflon FEP from various missions including MISSE 2, MISSE 4, MISSE 6, the 

Long Duration Exposure Facility (LDEF) and space exposed multilayer insulation retrieved from HST 

during the second servicing mission (SM2) are provided in Table 10, along with the Ey values for HST 

SM4 Bay 5 and Bay 8. Included in the table are the space exposure duration, solar exposure, AO fluence 

and AO exposure arrival. These Ey values were plotted in numerous ways including: Ey versus AO 

fluence, Ey versus time, Ey versus ESH and Ey versus ESH/AO ratio. The best fit was found for Ey 

versus ESH, and the corresponding graph is shown in Figure 28. It should be noted that although the best 

fit was found for Ey versus ESH, it is likely that temperature plays a critical role in the erosion process.

TABLE 10.—TEFLON FEP FROM VARIOUS MISSE MISSIONS, LDEF AND HST 

Material Mission Duration 

(Yrs) 

Solar 

Exposure 

(ESH) 

AO 

Fluence 

(atom/cm2) 

AO 

Exposure 

Ey 

(cm3/atom) Ref. 

5 mil 

Ag-FEP MISSE 4 1.04 1,400 2.15E+21 Ram 1.28E-25 19 

5 mil 

Al-FEP MISSE 7 1.5 2,400 4.22E+21 Ram 1.81E-25 20 

5 mil FEP MISSE 6 1.45 2,600 1.97E+21 Ram 1.69E-25 18 

5 mil 

Al-FEP MISSE 2 3.95 6,100 8.51E+21 Ram 2.11E-25 21 

5 mil FEP MISSE 2 3.95 6,300 8.43E+21 Ram 2.00E-25 17 

5 mil 

Ag-FEP LDEF 5.8 11,155 8.99E+21 Ram 3.37E-25 16 

5 mil 

Al-FEP HST SM2 6.8 33,640 3.20E+20 Sweeping AO 3.10E-24 5 
NASA/TM—2015-218476 29 

Figure 28.—Erosion yield versus ESH for Teflon FEP flown from various missions. 

5.0 Summary and Conclusions 

Two MLI blankets retrieved from the HST during SM4 after 19.1 years in space have been analyzed 

for space-induced changes in thickness, density, mechanical, optical and thermal properties, and surface 

morphology and chemistry. In addition, atomic oxygen Ey values were determined. The blankets 

experienced different solar exposures and thermal cycling temperature ranges due to their positions on the 

telescope and each had regions with differing types of damage. 

The mechanical properties of the Al-FEP outer-layers of the blankets were extremely degraded. The 

Bay 8 Al-FEP, which had a significantly higher solar exposure and on-orbit thermal cycling temperature, 

was more embrittled than the solar-grazing Bay 5 insulation, and was found to fracture like thin brittle 

glass. Areas on Bay 8 which had been protected by a patch for a period of time on-orbit were slightly less 

embrittled than areas which had been exposed for all 19.1 years. However, even the least brittle samples 

(Bay 5) still had only 3 percent of the elongation at failure of pristine materials. Heating pristine Al-FEP 

marginally decreased its strength, and had no effect on elongation. The Bay 8 MLI was too embrittled to 

conduct studies to assess the effect of heating, but heating Bay 5 Region 1 material reduced both the 

elongation and strength of the samples. 

The curled and patched regions experienced increased heating on orbit, and as-retrieved material from 

these regions had increased density as compared to pristine material and the other regions of HST 

material. Heating to 120 and 200 qC resulted in larger increases in density for space-exposed material as 

compared to pristine material. Increases in density are correlated to increases in crystallinity. Therefore, 

these results indicate that irradiation in space induced polymer chain scission in the HST material, which 

increased the mobility of polymer chains. In turn, this increased the extent of crystallization that occurred 

on heating either on-orbit, or post-flight in ground studies. 

 

1.28EͲ25

1.81EͲ25

1.69EͲ25

2.11EͲ25

2.00EͲ25

3.37EͲ25

3.10EͲ24

1.27EͲ24

1.17EͲ23

y = 2.68EͲ29x1.09E+00

R2 = 8.94EͲ01

1.EͲ25

1.EͲ24

1.EͲ23

1.EͲ22

1,000 10,000 100,000

Erosion Yield (cm3/atom)

Equivalent Sun Hours (ESH)

Mission ESH Ey

MISSE 4 1,400 1.28EͲ25

MISSE 7 2,400 1.81EͲ25

MISSE 6A 2,600 1.69EͲ25

MISSE 2 E6 6,100 2.11EͲ25

MISSE 2 E5 6,300 2.00EͲ25

LDEF AgͲFEP 11,155 3.37EͲ25

HST SM4 Bay 5 24,300 1.27EͲ24

HST SM2 33,638 3.10EͲ24

HST SM4 Bay 8 89,300 1.17EͲ23
NASA/TM—2015-218476 30 

All space-exposed samples experienced an increase in solar absorptance as compared to pristine 

Al-FEP, with the exception of Bay 8 Region 3, which is the region where the Al has delaminated from the 

FEP. The greatest increase in absorptance occurred in the patched areas, likely due to contamination from 

the patches. The thermal emittance values were found to decrease for all samples except Bay 5 Region 1 

and Bay 8 Region 11 (the region patched between SM2 and SM4). 

The Ey value of the HST SM4 Bay 8 nominal shiny FEP (1.37u10–23 cm3

/atom) was found to be an 

order of magnitude greater than for the Bay 5 nominal shiny FEP (1.43u10–24 cm3

/atom). This is 

attributed to the significantly higher dose of solar radiation, combined with the higher on-orbit 

temperature, for the solar facing blanket. The silicone contamination on Bay 5 may have decreased the Ey 

somewhat, but further studies are needed to assess the extent. The Ey values of the HST SM4 FEP were 

found to be orders of magnitude greater than those determined from prior flight missions, such as shuttle 

experiments (0.5 - 1.8u10–25 cm3

/atom), LDEF (3.37u10–25 cm3/atom) and MISSE 2 experiments 

(2.00u10–25 cm3

/atom). These results support the theory that the effects of solar exposure play a 

significant role in the AO erosion of FEP. 

Overall, the results of these tests support the proposed model for on-orbit degradation of 

Al-FEP, in which radiation causes chain scission of the polymers, and heating to high temperature 

extremes causes the Al-FEP to become more embrittled. 

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