ANALYSIS RESULTS OF THE MICRO-PARTICLES CAPTURER AND SPACE ENVIRONMENT EXPOSURE DEVICE (MPAC&SEED) EXPERIMENT ON THE INTERNATIONAL SERVICE MODULE

Yugo KIMOTO (1) *1, Junichiro ISHIZAWA(1), Eiji MIYAZAKI(1) , Hiroyuki SHIMAMURA(1) , Riyo YAMANAKA(1),

(1) Space Materials Section, Electronic Devices and Materials Group, Aerospace Research and Development Directorate, Aerospace Exploration Agency (JAXA), 2-1-1 Sengen, Tsukuba, Ibaraki 305-8505, Japan. *1Phone: +81-29-868-2317, E-mail: kimoto.yugo@.jp

ABSTRACT Three identical SM/MPAC&SEED units MPAC&SEED is a JAXA-owned experiment for (SM/MPAC&SEED #1, #2, and #3) were attached to the particle capture and material exposure mounted on the SM and flown. The SM/MPAC&SEED was launched Russian Service Module (SM) and KIBO Exposed aboard a Progress M-45 on 21 August, 2001. Facility (EF) of the International Space Station (ISS). SM/MPAC&SEED was unpacked and assembled by This report particularly describes analysis results of the inner vehicle activity (IVA). At 09:17 UT on 15 October MPAC&SEED on the SM (hereinafter 2001, all three units were mounted on the handrail SM/MPAC&SEED). This was a first experiment that outside the SM by extra-vehicular activity (EVA). On 26 prepared three sets of the same samples and evaluated August 2002, the first unit of the SM/MPAC&SEED, the relation between the material deterioration and the SM/MPAC&SEED #1, was retrieved by EVA after 315 exposure period. The MPAC is a passive experiment days (ten months (0.9 years)) of on-orbit exposure. designed to sample micrometeoroids and space debris. Subsequently, SM/MPAC&SEED #2 was retrieved on 26 The SEED is a passive experiment designed to expose February 2004 (after 865 days (28 months (2.4 years))). materials. The MPAC experiment succeeded in capturing Later, SM/MPAC&SEED #3 was repositioned to the dust and the componential analysis was done. The impact location that had been occupied by MPAC&SEED #2. flux from captured dust and debris models were Finally, SM/MPAC&SEED #3 was retrieved on 18 compared. From the SEED experiment, space August 2005 (after 1403 days (46 months (3.8 years))) demonstration data were acquired and the materials [1]. proposed by JAXA, universities, and companies in Japan proved to have high space environment durability. This 2. EXPERIMENTS paper summarizes the results from the 2.1 MPAC[1] SM/MPAC&SEED experiment. The MPAC is a passive experiment designed to sample micrometeoroids and space debris. Three types of 1. INTRODUCTION samples were prepared to capture and measure The SM/MPAC&SEED experiment is the space micro-particles for the MPAC. Silica-aerogel (hereafter, exposure experiment on the exterior of the Russian aerogel) is a transparent and porous solid with nanosized Service Module of the ISS. The most unique aspect of holes. It is used to capture dust particles intact and to the SM/MPAC&SEED experiment is that three identical estimate impact parameters (incident direction, particle components were manufactured. All three were exposed diameter, and impact velocity) based on the impact track at the same time and each was individually retrieved morphology. Polyimide foam was prepared to capture after varying periods of time. It was the world's first such large debris. The densities of aerogel and polyimide trial and this method can compare material aging foam used are 0.03 g cm-3 and 0.011 g cm-3, respectively. deterioration at virtually the same position. Another An aluminum plate was used to measure the number of unique feature is that samples capture micrometeoroids impacts from space debris or micrometeoroids. and space debris. This MPAC is a passive experiment designed to sample the micrometeoroid and space 2.2 SEED[2] debris environment and to capture particle residue for The SEED is a space material exposure experiment. later chemical analysis using aerogel, polyimide foam, The SEED consists of 28 samples, outlined in Table 1. and 6061-T6 aluminum. Another point is that the same Samples were proposed by JAXA, universities, and samples were arranged on both ram and wake sides. This companies in Japan and were selected by JAXA based on method should demonstrate the effect of AO, which their frequency of use and prospective future use. The collides with and erodes materials on the front and back SEED experiment also includes space environment of the [1]. monitoring samples which monitor the total dose of AO, UV, space radiation, and maximum temperature. possible impact residues, one just beyond the point after the track narrows, and one at the end. EDX analysis was Table 1. SEED samples conducted on both. The first one revealed Ag, Al and S, Sample name Organization Main Use and could be orbital debris. The second revealed only CF/Polycyanate, CF/Polyimide Heavy Industries Structural materials Ltd. background elements, and is thought to be altered PEEK (loaded & unloaded) Hokkaido University Inflatable structures aerogel. AlN Structural and Tokyo Institute of SiC (reaction sintering / Hot pressed) functional materials Technology TiN-coated Al / Al2O3 Ball-bearing (3 types) Tohoku University Mechanism application

TiN/MoS2/CuBN/Cu/-coated National Institute for Lubrication SUS304 Materials Science

MoS2 bonded film on Ti alloy IHI Aerospace Lubrication Loaded & unloaded polyimide film Inflatable structures (UPILEX-S) Modified polyimide film Thermal control Japan Aerospace Flexible OSR Exploration Agency White paint Silicone potting compound Potting Silicone adhesive Adhesion

3. RESULTS AND DISCUSSION 3.1 MPAC SM/MPAC results have been analyzed by some authors. Neish et al. described the inspection and analysis of the MPACs retrieved first and second (hereafter MPAC#1, #2) [2]. They showed the number of expected hits and the impacts on the aerogels, polyimide Figure 2. A typical carrot-shaped MPAC#2 impact. film, and aluminum plate on the MPAC#1 and #2. For The first fragment revealed Ag, Al and S, the second example, the largest feature found in MPAC#1 aerogel is background elements only [2]. shown in Fig. 1 together with the result of the EDX analysis performed on the inner wall of the impact track, Kitazawa et al. reported a detailed post-flight as highlighted by the white oval. Aluminum was analysis (PFA). They showed the numbers of identified, which Raman spectroscopy subsequently impact-induced features of the first quality level (Class I), indicated was in the pure form (as opposed to an oxide). which is a category that meets all of three criteria (<1> This suggests a debris impact. the feature has a crater-like rim and/or central peak, <2> the feature has radial cracks and/or ejecta, <3> the feature has a shape similar to those induced by hypervelocity impact experiments). Class II (the second quality level) has probable hypervelocity impact-induced features which meet one or two of the criteria. Class III has no hypervelocity impact-induced features. The number of impact-induced features was almost directly related to the exposure period (Fig. 3). The impact rate was almost constant with the sum of Class I and Class Ⅱ events running at about 15 impacts per year [3]. They also compared the calculated impact fluxes of the three environment models with the impact fluxes on aerogels estimated from inspection. They concluded that it is difficult to inspect small tracks and it was not possible to estimate the fluxes of diameter >10 μm of the aerogels retrieved second and third. Impact fluxes of aerogel appear inversely proportional to the exposure period and the fluxes are greater than the model results Figure 1. MPAC #1 impact (the largest found in [4]. MPAC#1 aerogel; entry hole diameter: ~2 mm, track length: ~5 mm). The inner wall was subjected to EDX analysis, revealing Al. Si and O are constituents of aerogel, and C was used as a coating [2].

An MPAC #2 impact is shown in Fig. 2. This large, typical carrot-shaped track (Fig. 7), over 1 mm long and almost perpendicular to the surface, revealed two cause of the crack formation on the silica aerogel surface; moreover, ultraviolet rays cause browning of the silica aerogel surface [7].

Fig. 3 Number of impact features of the first quality level (Class I) on SM/MPAC&SEED versus exposure [3].

Particle Diameter > 10μm Particle Diameter > 20μm 2.E+03 7.E+02 MPAC MPAC 2.E+03 ORDEM2000 6.E+02 ORDEM2000 /year] MASTER2001 2 MASTER2001 /year] 2 1.E+03 MASTER2005 MASTER2005 5.E+02 1.E+03

4.E+02 1.E+03

MPAC 8.E+02 MPAC (NO DATA ) (NO DATA ) 3.E+02

6.E+02

Cumulative Impact Flux [number/m Flux Impact Cumulative 2.E+02 Cumulative Impact Flux [number/m Flux Impact Cumulative 4.E+02

1.E+02 2.E+02

0.E+00 0.E+00 SM1 SM2 SM3 SM1 SM2 SM3 （315 days’ exposure） （865 days’ exposure）（1403 days’ exposure） （315 days’ exposure） （865 days’ exposure）（1403 days’ exposure） Fig. 4 Comparison of impact flux of MPAC aerogels and Fig. 5 Photomicrographs of tracks investigated in the calculated results of three models (Left: Particle diameter study. Terminal particles are indicated by black arrows. >10 µm, Right: Particle diameter >20 µm ) [4]. (A) SM1_3RA1, (B) SM2_4WD1, (C) SM3_3WD1, (D) an enlarged image of the terminal particle in the track Noguchi et al. discovered that the number density of shown in (C) [5]. craters on the surfaces of the WAKE facing tiles is smaller than those of the RAM facing aerogels. However, their depth/crater diameter ratios are larger than those facing towards the RAM side. We investigated three terminal particles found at the ends of tracks in silica aerogels. Combined SEM, transmission electron microscope (TEM), micro Raman spectroscopy, and synchrotron radiation X-ray diffraction analyses revealed that they are space debris, secondary debris and a micrometeoroid, respectively. Photomicrographs of the tracks are shown in Figure 5 [5]. A new type meteoroid was found by means of an oxygen isotopic composition analysis from in this study. This work will be published Micron- size particles soon. 100μm Many unexpected micron-size particles that did not Fig. 6 The micron-size particles on the top surface of the impact with hypervelocity were found. An example is aerogel [6]. shown in Figure 6. The terminal micron-size particles are almost transparent and colorless. They are so small that we have not succeeded in analyzing them. We have examined the analysis technique [6]. Meanwhile, the surface degradation of silicon aerogel by the space environment was suggested. The RAM side aerogels became whitened with many craters. On the other hand, the wake side aerogels were yellow and showed countless cracks as islands, which became smaller with exposure period. Figure 7 shows retrieved (#1, #2, and #3) aerogel surfaces and preflight aerogel surface images obtained using optical microscopy with reflected light. We evaluated the effects of AO and UV on silica aerogels. The results suggest that AO is the unexposed surfaces or those subjected to EB and UV irradiation. Cross-sectional examination was conducted in order to research the internal quality such as micro-cracks and delamination. Delamination was observed in all CF/PC including the unexposed specimen. Cracks were not observed in CF/PC. This result supports the characteristic of this material that susceptibility to form micro-cracks was very low. On the other hand, many cracks were observed in the CF/PX made using thermoplastic polyimide. Resin damage and carbon fiber without resin were observed on the surface of the specimens in Fig. 8.

Fig. 7. Surface change of SM/MPAC aerogel (Images of optical microscope with reflected light) [7].

3.2 SEED

3.2.1 CF/Polycyanate and CF/Polyimide[8] (a) CP/PX 46 months exposed Composite materials such as CFRP were applied to spacecraft especially reusable launch vehicle (RLV) more and more because of their high specific strength and specific elasticity. However effective data have not obtained even though the polymer matrix is degraded by space environment factors such as electron beam, ultra violet ray, high vacuum, thermal cycle and atomic oxygen. Effective data were virtually non-existent due to the difficulty of simulating space conditions. Two heat-resistant composite materials were evaluated in the experiment. One is YS90A/RS-3 consisting of pitch-based carbon fiber and polycyanate resin (CF/PC). The other is IM600PIXA consisting of PAN-based carbon fiber and the thermo plastic resin (CF/PX). The specimen size is 17 square mm with a thickness of 3 mm. Ground experiments applying space environment factors such as electron beam, ultraviolet ray, and atomic oxygen were conducted separately as reference experiments. From visual observation, surface fading of the space-exposed specimens and atomic-oxygen-irradiated (b) AO for 1 year ground irradiated CP/PX specimens was confirmed. The surface of the exposed Fig. 8. Cross section of CP/PX [8] area is rougher than that of the area for sample holder fitting. The rough surface was also seen by microscope Ultrasonic observation was conducted to detect examination in both of the materials exposed to space foreign objects and delamination. No indication was environment conditions and atomic oxygen irradiation. observed in either CF/PC and CF/PX. Fourier Transform On the other hand, no change was observed in the Infrared (FT-IR) spectroscopic analysis is a chemical analysis method and can get the resin degradation by matter. The detail of this feature is shown in Fig. 11. The measuring a organic composition change. The same pale whitish region was observed in about 70~80% of the trend was confirmed in the results from CF/PC and exposed area regardless of the applied stress and CF/PX. No change of organic composition was found. exposure period (the arrow A in Fig. 11). The EPMA analyses showed that Si and O elements were mainly 3.2.2 Poly-ether-ether-ketone (PEEK) [9] detected; therefore, this region can result from the In recent years, deployable structures have been chemical reaction between Si-derived contamination and acknowledged to be a leading edge technology for the atomic construction of space facilities. Their use is expected in oxygen. many space structures, such as large space antennas, photovoltaic generation systems, solar sails, etc. Polymeric films are one of the candidate materials for these deployable structures. In these applications, polymer films have to withstand a certain amount of load in space over a long term. Therefore, the reliability of mechanical properties in space is one of the most important factors for them. The material used was PEEK film with a thickness of 0.4 mm (FS-1100C, Sumitomo Bakelite Co., Ltd.). PEEK is not a commonly used spacecraft material compared to polyimide. However, it has many attractive properties due to its thermoplastic characteristics. Three specimens with different widths were machined as shown in Fig. 9. The axial direction of the specimens was set to the extruded direction of the film. The specimens were loaded by a tension spring with initial setting stresses of 0, 1.6, and 4.7 MPa during the experiments. These values corresponded to about 0, 2 and 5% of the yield strength (≈85 MPa).

Fig. 10. Surface appearances after space exposure under no tension [9].

Fig. 11. Magnified photograph of the 28-month flight sample under 1.5 MPa. The arrow A shows the pale Fig. 9 Flight test specimens [9]. whitish region [9].

Fig. 10 shows the surfaces of the flight samples The mass of the specimens decreased with an after space exposure. The color of the exposed areas increasing exposure period, which resulted from a clearly changed into brown. The browning became thickness reduction caused by AO attack. The average deeper with and increasing exposure period. However, thickness decrease, ∆h, was calculated by using the the amount of browning had no relation to tensile stress. equation below: A similar phenomenon was not seen in AO and EB Δm irradiation but in the UV irradiation of ground tests. The Δh = (1) ρ ⋅ A UV constituent in a LEO environment can be the main where ∆m is the mass loss, ρ is the density of reason for the browning of the sample surface. In 3 addition to the browning of the sample, specimen PEEK=1.3g/cm , and А is the exposed area. Fig. 12 surfaces were slightly covered with the pale whitish shows the relationship between ∆h and the exposure period. ∆h increased with increasing exposure period. However, it tended to deviate from a straight line. Stresses during exposure seemed to have no significant effect on ∆h.

Fig. 14. Relationship between the yield strength and exposure period of the flight samples [9].

Fig. 12. Thickness decrease of the flight samples [9].

The relation between the elastic modulus and exposure period is shown in Fig. 13. Although each elastic modulus had a wide distribution, the range of the flight samples was within that of the pristine ones. Fig. 14 shows yield strength. Yield strengths of flight samples were almost equal to the pristine one regardless of the exposure period or tensile stress. Fig. 15 shows the relationship between elongation and exposure period. The elongation at break decreased to about 15% of the pristine sample after 46 months of exposure. The UV irradiations in the ground tests showed similar tendencies. Fig. 15. Relationship between the elongation at break X-ray Photoelectron Spectroscopy (XPS) and and exposure period of the flight samples [9]. Differential Scanning Calorimeter (DSC) analyses suggested that crosslinking induced by the UV in LEO Based on the ground tests, it became clear that the was the main factor for the change of elongation. As for EB constituent in the ISS orbit hardly influenced the the 10-month and 28-month exposures, the decrease in tensile properties. Details of these analyses are described elongation of the stressed samples was less than that of in Nakamura, et al. [9, 10] the unstressed samples. However, the phenomena diminished after 46 months of exposure. This indicated 3.2.3 Loaded and unloaded polyimide film that tensile stress had a tendency to prevent (UPILEX-S) [11] embrittlement in the early stage of exposure. Polyimide has been applied as a base film for deployable structures of spacecraft, such as large flexible solar arrays, large deployable antenna and solar sails, because of its high specific strength and rigidity, high dimensional accuracy, and low rate of thermal expansion. Base films are used under tensile stress to maintain their structural shapes at low gravity when deployable structures are fully extended. The initial tensile stress is expected to be altered by orbital thermal cycling. For application as a structural material, it is indispensable to evaluate the effects of the space environment on mechanical properties. The effects of tensile stress should also be understood if they hasten the degradation of mechanical properties in a space environment. The tested polyimide films were 125-µm-thick UPILEX-S Fig. 13. Relation between the elastic modulus and (UBE Industries Ltd.). UPILEX-S has been applied as a exposure period of the flight samples [9]. base film for the flexible solar array of the (SFU) and the Advanced Earth Observing Satellite-I and II (ADEOS-I and II). The sample dimensions are presented in Fig. 16. The sample has a dog-bone shape which resembles the “Type IV” specimen of the American Society for Testing and Materials (ASTM) Standard D-638-03, punched from a sheet using a die. The tensile stress applied to the samples was set to 0 MPa, 1.4 MPa, and 7.0 MPa by adjusting the spring elongation. The tensile stress level of 1.4 MPa was based on the nominal stress of the base films of the ADEOS-I solar paddles; 7.0 MPa was chosen to be five times as large as 1.4 MPa.

Fig. 17. Tensile strength of the flight samples [11].

Fig. 15. Schematic view of the sample [11].

The thickness loss of flight samples is expected to become larger with increased exposure duration. There is, however, no remarkable difference of the thickness loss among flight samples, as presented in Fig. 16. In addition, the thickness loss of Flight #3 was about 8 µm, a much lower than expected value of approximately 250 µm for three years of space exposure. In ground reference tests, only AO irradiation decreased the sample thickness; there was almost no variation in the mass of UV and EB Fig. 18. Elongation of the flight samples [11]. irradiated samples. Tensile strength and elongation changes of the flight Ground reference tests demonstrated that the AO samples versus exposure duration are portrayed in Fig. irradiated samples underwent considerable degradation 17 and 18. The tensile strength of the flight samples had of tensile strength and elongation. The samples exhibited decreased slightly from the value of control samples. A no marked degradation of mechanical properties by major reduction in elongation was detected, showing a either UV or EB. This result clearly suggests that AO maximum of 70% loss compared to the control samples. attack is the main cause of decreased tensile strength and Samples of Flights #2 and #3 showed more serious elongation in a LEO environment. A tensile stress of less degradation than Flight #1 in both tensile strength and than 7.0 MPa, which was applied to the samples during elongation. the space exposure and irradiation tests, had little effect on the degradation of the mechanical properties of any sample. The tensile strength and elongation were decreased with increased AO fluence. In addition, the surface roughness development was proportional to the increased AO fluence. The flight sample surfaces that had been exposed to the space environment were examined using an SEM, as presented in Fig. 19. The samples were tilted at 45 degrees to facilitate viewing of the surface topography. Consequently, the rough surface was regarded as one cause of the degradation of mechanical properties. It is generally assumed that excessive stress concentrates at a concave region on the rough surface during deformation leading to the formation of surface cracks and initiation points of destruction. The flight samples’ surfaces showed a rough texture from AO erosion. Additionally, some extremely Fig. 16. Thickness loss of the flight samples [11]. deep, compared to the vicinity, concavities were found on the surface. These deep concavities have the potential to reduce mechanical properties considerably.

3.2.4 Aluminum Nitride (AlN), Hot-Pressed (HP) and Reaction-Sintered (RS) Silicon Carbide (SiC), Titanium Nitride (TiN) ion-plated on aluminum plate and alumina plate (ceramic) [12] Ceramic materials have beneficial properties such as high specific strength, tolerance to high temperature, great hardness and low friction. Such ceramic materials are useful for heat shield and/or insulator materials in space exploration where there exist many environmental factors such as particles, radiations, gravity, pressure, temperature. The research regarding the characteristic change in ceramics due to a long period of exposure in the space environment has been limited, although several space environment exposure missions have been Fig. 20. Roughness of HP-SiC and RS-SiC samples performed in the past by NASA and JAXA. In the [12]. present work, five kinds of tests have been carried out using three ceramic materials: AlN, HP-SiC, RS-SiC, In the case of space exposure, a slight increase in TiN/Al, and TiN/Al2O3 at 17 mm x 1 7 mm x 2 mm. All average roughness was observed after 0.9 year of samples were the same size. exposure but it decreased to less than the blank sample The change in average roughness of the SiC sample after 2.4 years of exposure, and increased again after 3.8 surfaces after space exposure and AO irradiation is years of exposure. The average surface roughness of the shown in Fig. 20. Irradiation of AO increased average RS-SiC samples was generally greater than that of the surface roughness with increasing AO irradiation, HP-SiC samples. although no correlation was found between the effect of According to the SEM examination of the HP-SiC AO irradiation and that of space exposure. and the RS-SiC samples, the polishing damage which existed on the SiC surface before space exposure gradually diminished with increasing space exposure duration. As shown in Fig. 21, the tiny particles which were thought to be SiC particles were observed for RS-SiC ceramics at the grain boundary silicon phase of the space-exposed samples. But it was rarely or never observed at the grain boundary of the sample irradiated by AO on the ground. In the cases of TiN/Al and TiN/Al2O3 tests, the morphologies exposed in space were much different from those irradiated by AO as shown in Fig. 22.

Fig. 21. Surface morphology changes of the RS-SiC samples during the exposure periods [12].

Fig. 19. SEM images of (a) the control sample, (b) Flight #1 and (c) Flight #2, and (d) Flight #3. The flight samples were under no tensile stress during space exposure [11]. The O/Si ratio versus the change in depth of the space-exposed sample and the ground AO-irradiated sample was compared using a Secondary Ion Mass spectrometer (SIMS).

3.2.5 Ball-bearing (lubricant) [13] Lubrication is one of the most important key technologies for the reliability and long life of space systems. In the present space systems, solid lubricants such as lead, indium, silver, gold and molybdenum disulfide are pre-coated with a certain thickness before Fig. 22. Surface morphology changes of the TiN/Al assembly. The life of present space systems is, therefore, and TiN/Al2O3 samples [12]. determined by the wear life of the lubricant coating. If solid lubricants could be replenished in-situ and Figs. 23(a) and (b) show the oxygen concentration on-demand during operation even when exposed to the of the surface of the HP-SiC and RS-SiC samples real space environment, the life of the space systems detected using a Wavelength-Dispersive-type X-ray could be extended. It is believed that “a self-replenishing spectrometer (WDX). The oxygen content of the surface lubrication system with in-situ and on-demand of the space-exposed samples (detection range: ~1 mm in controllable lubrication method” is needed for the future depth) not covered by a fixture jig increased markedly space systems to improve reliability and overcome compared to those of the blank samples. The surface unexpected tribological problems in space. To satisfy the oxygen content of the samples irradiated by AO also requirements, a new solid lubrication method called increased. Therefore, it is thought that AO is a principal "Tribo-coating" has been proposed. It has also been factor of the oxidation of the sample. To the contrary, the clarified that this method gives the unique advantage of oxygen content of the surface covered by the fixture jig excellent tribological performance such as low friction was comparable or slightly higher to that of the blank and a semi-permanent lifetime by the in-situ and samples and far lower compared to the uncovered parts on-demand formation of an optimum tribo-coating film. of the space-exposed samples. Therefore, direct The goals of this experiment are to clarify the effects of bombardment of AO with an energy of 5 eV (8 km/s) the real space environment, such as radiation and atomic appears to be essential for the surface oxidation of SiC. oxygen, on the friction properties of ball bearings Carbon should become volatile component and be lost lubricated by the tribo-coating of indium and to show the from the sample surface. possibility of replenishing the lubricant that might be damaged by exposure to the real space environment. Two types of ball bearings coated with a tribo-coating of indium and conventional vacuum ball bearing coated by sputtered MoS2 with a special retainer of PTFE as shown in Table 2 were used as one set of specimens.

Table 2 Specification of three types of ball bearings

Friction curves of the BR1 and BR3 ball bearings before and after exposure to the space environment for one year are shown in Figs. 24 and 25, respectively. It is Fig. 23 Oxygen depth profiles from the surfaces of the clearly seen that friction coefficients of both bearings HP-SiC and RS-SiC tested samples [12]. after one year of exposure are similar to those from before exposure. It is, therefore, clear that a tribo-coating bearings under space exposure conditions is less than 2.4 of indium is a useful lubricant for space systems and years. Indeed, the friction property of conventional works for times of less than one year. vacuum ball bearing (BR3) appears quite unstable with a large variation in the friction coefficient after two years exposure. In Fig. 27, indium was evaporated when the friction coefficient of a ball bearing lubricated by tribo-coating (BR2) starts to increase due to its lifetime after exposure for one year. The friction coefficient is kept low by re-tribo-coating (in-situ tribo-coating).

Fig. 24 Friction curves of the BR1 ball bearings before and after exposure in the space environment for one year [13].

Fig. 27 Friction curves of conventional vacuum ball bearings (BR3) after exposure in the space environment for one and two years [13].

This shows the possibility of in-situ and on-demand replenishment of solid lubricant in lubrication systems with tribo-coating for mechanical systems in the real space environment. This is a unique advantage of tribo-coating with in-situ and on-demand replenishment of lubricant for space mechanisms possibly damaged by Fig. 25 Friction curves of the BR3 ball bearings before exposure to the real space environment. and after exposure in the space environment for 0.9 year

[13].

(1 year)

Fig. 28 Possibility in-situ restoration of lubricant for Fig. 26 Friction curves of the ball bearings lubricated ball bearings after exposure to the real space by a tribo-coating (BR1) after exposure in the space environment [13]. environment for 0.9, 2.4 and 3.8 years [13]. 3.2.6 Titanium Nitride (TiN), Molybdenum Figures 26 and 27 show the effect of the exposure disulfide (MoS2), a mixture of Copper and Boron period on friction properties of two types of ball bearings Nitride (Cu/BN) and Copper (Cu) coated SUS304 (BR1 and BR3), respectively. Although the friction (lubricant) [14] coefficients are low after 0.9 year of exposure as The materials of moving components in an orbital mentioned before, those values increase with exposure of environment require surface modification with stable more than two years. It is certain that the lifetime of both lubrication to prevent an increase in the friction caused by oxidation and irradiation damage. A stainless steel substrate and four kinds of lubricating coatings such as TiN, MoS2, a mixture of Cu and BN and Cu were prepared. Commercial type 304 austenitic stainless steel sheet substrates (size: 14 mm x 14 mm x 1 mm) have been coated to a film thickness from 50 nm to 250 nm with four lubricants using an RF magnetron sputter deposition system. Figure 29 shows the effect of vacuum annealing on the vacuum friction coefficients of coatings on substrates with a variety of exposure test years. A one-year exposure test shows that substrates kept a low coefficient of friction (μ) even after vacuum annealing. However, a test with 2.4 years of exposure caused some substrates to increase μ. An exposure test of 3.8 years demonstrates the obvious difference in the increase in μ among substrates. The SUS304 substrate, TiN coating and Cu coating show a large increase in μ, a mixture coating of Cu and BN shows a small increase in μ and the MoS2 coating shows little increase in μ retaining good lubrication despite the long exposure on orbit. A comparison between surface analytical results, performed by X-ray photoelectron spectroscopy (XPS), of exposed and screened area of TiN coated stainless steel sheet showed the following results. The TiOx layer formed during TiN film preparation. Silicon evaporated from silicone adhesives used for space station structures might react with atomic oxygen to form SiO2 during exposure in orbit. The mixed structure of SiO2 in coating films is considered to form a good lubricant.

3.2.5 MoS2 bonded film on Ti alloy (lubricant) IHI Aerospace (formerly NISSAN Motor Co., LTD.) has proposed development of the Payload Attach Mechanism (PAM) to carry the external payloads of the Japanese Experiment Module (JEM) on the ISS by the

Space Shuttle. The HMB-34 MoS2 bonded film was selected for the JEM ELM-ES PAM from in-house research and development conducted by IHI Aerospace. HMB-34 is a kind of MoS2 with organic binder coated onto Ti-6Al-4V, CRES and aluminum alloys on the JEM ELM-ES PAM[15]. The solid lubricant evaluated in this experiment is MoS2 bonded film currently used for the above application. An organic binder, polyamide-imide, is used for the film. That material constitutes about 65% of the film. The film thickness was about 10 μm. Titanium alloy was used as the test specimen substrate [16,17]. Photographs of the film surfaces and wear tracks taken after friction tests using an optical microscope are presented in Fig. 30. The film surfaces as shown by SEM examination are also depicted in Fig. 30. Changes in the Fig. 29 Effect of vacuum annealing on the vacuum film color were observed for the flight and AO samples friction coefficients (μ) of substrates versus the exposure in comparison to the reference [16]. term; (a) SUS304 substrate, (b) TiN coating, (c) MoS2 Adhesive property, friction property and surface coating, (d) Cu coating, (e) mixture coating of Cu and property evaluations were conducted on all test articles BN. [14]. including those with 10, 28 and 46 months of exposure to the space environment and the same effects were measured.

Figure 31 presents a comparison of peaks of Mo 3d According to 46 months of exposure to the ISS from XPS analysis. Peaks of Mo (VI) were observed on environment, the following conclusions are reached: the film surface for the flight and AO samples, although a) Initial friction of space exposed HMB-34 MoS2 the peaks of molybdenum for the UV, EB and reference bonded film has been decreased compare to ground samples were Mo (IV). Molybdenum seemed to change initial condition (control) and has not changed. MoS2 to MoO3 because of atomic oxygen during orbit b) Overall friction has decreased and life time of and the ground test. The Mo (VI) peak remained on the HMB-34 bonded film seems to increase. rubbing track of only the AO sample after 20,000 strokes c) This property change might have some relation to [16, 17]. the SiO2 contamination which also proved not to change the MoS2 bonded film lubrication property. This space exposure has proved that the HMB-34 bonded film provide enough performance for JEM-exposed mechanisms such as the ELM-ES PAM SLM for 10-year operation in the ISS space exposed environment [15].

3.2.6 Modified Polyimide (Polyimide-siloxane (a) Photographs of film surfaces and rubbing tracks coated UPILEX-R) [18] It is important to improve the durability of thin polymer thermal-control films against AO erosion. JAXA has developed new polyimide films with a high tolerance against AO attack. This 50 μm thick UPILEX-R® ([biphenyl tetra carboxylic dianhydride]-[4,4’- oxydianiline] BPDA-ODA) film is coated with a 3-μm-thick polyimide siloxane layer. Two (b) SEM observation Modified PI specimens with different treatments (coated Fig. 30 Surface observation of the 46-month exposure and uncoated) were exposed. The uncoated specimen sample [16]. was equivalent to normal UPILEX-R®. The specimen exposure area for flight and ground experiments was a 30 20-mm-diameter circle. (a) Film surfaces Figure 32 portrays the second retrieved Modified PI, 25 both coated and uncoated. The exposed area of the coated sample was browned slightly; that of the uncoated 20 sample was lusterless, appearing to be eroded by AO. A 15 B

10 C

5 D E 0 245 240 235 230 225 220 Binding Energy 30 A Flight sample (b) Rubbing tracks B AO sample Coated Uncoated 25 C UV sample D EB sample 20 E Reference Fig. 32 Second retrieved Modified PI [18] A 15 B The changes in mass, solar absorptance (αS), and infrared emittance (εN) of the retrieved specimens were 10 C measured. The αS of the Modified PI was increased, but 5 D it was less than that of the uncoated sample. The

E Modified PI sample demonstrated its high AO tolerance. 0 An inorganic SiO2 layer was observed on an 245 240 235 230 225 220 uncoated sample. Figure 33 is an FE-SEM image of the Binding Energy uncoated specimen’s surface with a porous layer and Fig. 31 Peaks of Mo 3d by XPS analysis: (a) film carpet-like matrix. The carpet-like geometry is typically surfaces, (b) rubbing tracks after 20,000 sliding strokes found in AO-irradiated polyimide. [16]. The siloxane layer’s effective AO protection was confirmed in the coated sample. The results of the uncoated sample indicate that a SiO2 subjected to this material was exposed to a space environment in the outgassing siloxane contamination provided inadequate cargo bay of the 85 for 10 days. High protection. stability was confirmed, but the exposure period was too short to evaluate its stability in a space environment. For SM/SEED experiments, white paint specimens were coated to a thickness of 65μm onto aluminum alloy plates and a 21-square-mm area was exposed. Critical damage, such as peeling, cracks, and notable changes in thermo-optical properties, was not observed in any retrieved sample. Mass loss induced by AO erosion was quite small. Solar absorptance increased during the 315–865-day-exposure period; after 865 days of exposure, it was stable. Figure 35 presents FE-SEM 5 μm surface images of white paint specimens. White pigmentation was obviously identified on non-exposed samples. Silicon contamination deposited on the white

Fig. 33 FE-SEM image of the uncoated Modified PI paint surface probably affected the solar absorptance and surface retrieved second [18] increased AO protection.

Figure 34 presents TEM images of the cross-section (a) No exposure (c) 865-day exposure for the samples retrieved second. The second-retrieved uncoated sample’s SiO2 layer created by outgassing siloxane contamination was thicker and more uniform than the first-retrieved one. For that reason, the contaminant silicone layer might increase during exposure and become a more effective barrier against AO. However, that depends on several factors that affect contamination condensation, such as the atmosphere (b) 315-day exposure (d) 1403-day exposure (including outgassing), temperature, and active constituents of materials (AO, UV, EB, etc.). The contaminant barrier is uncertain and unreliable depending on individual missions. Uncoated samples were inferior to the coated sample in solar absorptance stability in the SM/SEED experiment. A highly AO-tolerant material is necessary for the LEO environment. Fig. 35 FE-SEM images of a white paint surface in SM/SEED retrieval samples [18].

3.2.8 Flexible Optical Solar Reflector (F-OSR) [19] The F-OSR provides low solar absorptance (αS) and high infrared emittance (εN) with flexibility. The F-OSR is a five-layered and second-mirrored film with a polyetherimide (PEI) base film. It has a UV protection 500nm 500nm layer to protect the base film against UV degradation. Fig. 34 TEM images of cross-sections for the first- and The SM/SEED experiment can reveal whether the second-retrieved Modified PI (uncoated specimen) [18] function of F-OSR can be maintained for a period on orbit on the order of one year. 3.2.7 White paint (NOVA 500 AstroⓇ White) Results indicate a mass increase. Atomic oxygen [18] does not erode F-OSR in low earth orbit. Thermo-optical JAXA has developed silicone-based white paint in properties show no marked change. In fact, F-OSR was collaboration with Nippon Paint Co. Ltd. The white paint verified as retaining its initial properties after exposure (NOVA 500 Astro® White) was designed as a on ISS orbit for 46 months. low-outgassing type for space use and qualified as Cross-sectional TEM images of the near exposed NASDA-1049/101-S. It comprises a silicone-resin surface are presented in Fig. 36. In those images, the matrix with mainly zinc oxide (ZnO) as a white pigment. dark area at the bottom is the ITO/CeO2 layer A bright JAXA has demonstrated its stability with regard to space area appearing in the upper half is mounting material for environments using a ground test facility and space the TEM examination and is not a part of the F-OSR. exposure experiments: Evaluation of Space Environment These areas are visible in both blank sample and and effects on Materials (ESEM). In ESEM experiments, retrieved samples. A gray area is apparent between the dark area and the bright area only in images of the retrieved samples. The gray area would be a new layer built during exposure and covering the flight samples. This new layer thickness increased concomitantly with the exposure duration. A Scanning TEM and Energy Dispersion X-ray Analysis (STEM-EDX) analysis reveals that the new layer consists mainly of Si and O suggesting that the new layer is silicon dioxide produced chemically from silicon contamination and AO. (a) Blank sample 3.2.9 Silicone potting compound [20] To evaluate the tolerance of a silicone potting compound and adhesive developed by JAXA to a space environment, specimens of the developed silicone potting compound and adhesive were exposed to space outside the ISS as SM/SEED samples. The silicone materials are fundamentally designed for inside use, exposed environments were not considered. For promotion of applications of the developed materials to outside use, this experiment was performed. The purpose of the SM/SEED mission was to evaluate whether the (b) #1 retrieved materials have a tolerance to the space environment outside the ISS or not. The sample is a two-part silicone potting compound “KE-101A/B” manufactured by the Shin-Etsu Chemical Co., Ltd. The potting compound was applied 3-mm thick onto aluminum substrates of 20 mm × 20 mm × 1 mm (thickness) as an SM/SEED sample. The samples are set in holes made in the SM/MPAC&SEED hardware, held by holding plates with a 16 mm × 16 mm window. The area within the window is the actual exposed area. The retrieved samples were evaluated by macroscopic examination, mass measurement, measurements of the electrical properties (relative permittivity ε and dielectric tangent tan δ) and (c) #2 retrieved cross-sectional microscopic observation near the surface. The results of this experiment verify that the JAXA-developed silicone potting compound can retain its properties within the requirement for 46 months of exposure on the ISS. A new layer of SiO2 was observed on the retrieved sample surface, similar to other silicone materials that collided with AO. The results of this investigation show that the mass loss is mainly the result of outgassing.

3.2.10 Silicone adhesive [20] The sample is a one-part silicone adhesive (KE4908SC-T; Shin-Etsu Chemical Co. Ltd.). The adhesive of 13.2 mm (L) × 25.4 mm (W) × 2 mm (t) is cured between two 83.2 mm × 25.4 mm × 1 mm (t) aluminum plates before flight. The specimen was tested for adhesive properties according to the standard test (d) #3 retrieved specification (JIS K 6850 – Adhesive – Determination of tensile lap-shear strength of rigid-to-rigid bonded Fig. 36 Cross-sectional TEM images of F-OSRs for assembly). SM/MPAC&SEED [19]. Degradation in elongation and loss of ductility were found, although the required properties are verified to be retained for up to 46 months of exposure on the ISS.

3.2.11 Space environment Monitoring samples fluence was 1020 atoms/cm2 from Vespel and 1021 [21] atoms/cm2 from PAMDEC. The MPAC&SEED #1 data The SM/MPAC&SEED was a passive experiment showed higher values than those of MPAC&SEED #2, that used neither a power source nor communication. although MPAC&SEED #2 had longer exposure than #1 Therefore, in-situ information was not telemetered from in the AO fluence. A similarly unexpected result also space. Samples for monitoring the total dose of AO, UV, occurred in the UV monitoring samples. The measured space radiation, and temperature were situated on board. intensity of UV in the wake face was greater than that The space environment monitoring samples are shown in from the ram face. The TLD data were dependent on the Table 3. shield thickness. Orbital and attitude flight information of the ISS during this experiment period was analyzed. It Table 3. Space environment monitoring samples appears that both the ram and wake faces of the Function Sample name SM/MPAC&SEED were pointed in the flight direction. UV monitor Urethane Sheet A comparison of flight data with data from the AO monitor VESPEL space-environment model was performed. Some PAMDEC discrepancies between the flight data and model Dosimeter RADFET calculation is discussed. ALANINE Dosimeter One reason for the discrepancies between the flight data TLD and the model calculation was considered that both ram Maximum temperature Thermo label and wake faces were pointed in the flight direction for AO fluence. A layer which is produced by contamination

was grown in flight. It protected the surface of the Table 4 presents results derived from sample monitoring sample from erosion. For UV fluence, the monitoring. The first-retrieved monitoring sample data ISS itself or some components in the field of view of the are labeled as #1, the second-retrieved are labeled as #2 MPAC&SEED trays suggested shading of the UV and the third-retrieved are labeled as #3. The maximum irradiation in the ram direction. temperatures in the three trays were 50–90°C. The AO

Table 4 Derivation results from the monitoring samples [21] Ram Face Wake Face #1 #2 #3 #1 #2 #3 Maximum Temperature [°C] 50a 50a 60a － － － 60b 90b 90b AO Vespel 2.04 × 1020 2.57 × 1020 2.70 × 1020 1.61 × 1020 2.05 × 1020 3.09 × 1020 [atoms/cm2] PAMDEC 2.41 × 1021 1.36 × 1021 1.37 × 1021 1.93 × 1021 1.22 × 1021 4.82 × 1021 UV [ESD] 18.1 15.8 13.4 122.2 201.0 205.5 TID [Gy] Alanine 1.95 15.3 32.0 3.5 21.9 58.3 dosimeter c RADFETd 0.44 5.99 8.59 0.27 4.92 14.9 TLDe 1.46 × 10-3 0.12 0.29 3.41 × 10-3 0.09 0.04 a At approximately 5 mm depth in Tray Nos. 1 and 2. b At approximately 1 mm depth in Tray Nos. 3 and 4. Shield thickness; c 0.04 [g/cm2], d 0.2 [g/cm2],e 1.2 [g/cm2]

for the ram side of SM/MPAC&SEED are a good gauge 3.2.12 Contamination [22] for qualitative and quantitative comparison with The induced contamination predictions and predicted contamination. XPS measurements for the measurements were compared. Material outgassing and wake side are a good gauge for qualitative comparison thruster plume induced contamination were calculated but have limitations with regard to quantitative using analytical and semi-empirical models developed by comparison. Nevertheless, consistent XPS results the Boeing Space Environments Team in Houston. showing the most prominent presence of nitrogen on the Quantitative comparisons of the measured and predicted wake face from all 3 units provide great confidence in levels of contamination are provided in Table 5. The the predictions for plume contamination on the wake calculated depth of contamination on the ram side side. surfaces is within a factor of 3 of the measured contamination. Plume contamination can be more Table 5 Contamination depth predicted and measured difficult to quantify with XPS measurements than (unit, nm) outgassing induced contamination. Whereas the #1 #2 #3 Measured* Predicted Measured* Predicted Measured* Predicted outgassing contamination was dominated by Ram 31,32 10.6-13.575,75 30.3-35.4 93,93 45.9-53.3 silicon-based outgassing sources, thruster plumes have Wake 2,55 8.6-10.3 0,4 18.6-23.7 2,2 31.7-41.4 * The contamination layer thickness is obtained by XPS depth profile of SiO2. Two points was multiple byproducts. Whereas outgassing yields a fairly measured. uniform molecular contamination layer, thruster plume induced contamination is dominated by the liquid phase, producing droplet features and a non-uniform distribution of contaminants. Hence, XPS measurements 4. CONCLUSIONS Engineering Review Vol. 40, No. 1, 31-41, (2008) The results of SM/MPAC&SEED are as follows: [5] Proc. of International Symposium on - MPAC “SM/MPAC&SEED Experiment”, Japan, 2008, It succeeded in capturing dust, and the JAXA-SP-08-015E (2009) pp. 59-66. componential analysis was done. The impact flux [6] Yugo Kimoto, Riyo Yamanaka and Takaaki Noguchi” from captured dust and debris models was compared. Analysis results of dusts from Micro-Particles Capturer - SEED on board Service Module of the International Space Space demonstration data were acquired and Station”, Proc. of the 27th International Symposium on were proven to have high space environment Space Technology and Science (2009), 2009-r-2-12. durability for the following materials: [7] Riyo Yamanaka and Yugo Kimoto,” Effect of atomic CF/Polycyanate, CF/Polyimide, PEEK, ceramics, oxygen and ultraviolet rays irradiation to Aerogel”, Proc. lubricants, Modified polyimide, white paint, F-OSR, of the 26th International Symposium on Space silicone potting compound and adhesive. Technology and Science (2008), 2008-r-2-24. The elongation of loaded PEEK was degraded [8] Proc. of International Symposium on according with the exposure time. It is suggested “SM/MPAC&SEED Experiment”, Japan, 2008, crosslinking induced by UV in LEO was the main JAXA-SP-08-015E (2009) pp. 67-71. factor for the change of elongation. [9] Proc. of International Symposium on The tensile strength and elongation of loaded “SM/MPAC&SEED Experiment”, Japan, 2008, polyimide film deteriorated. The degradation factor JAXA-SP-08-015E (2009) pp. 73-80. is atomic oxygen. [10]Takashi Nakamura, Hiroshi Nakamura and Hiroyuki Silicone rich contamination layers were Shimamura, “Effects of LEO Environment on Tensile observed on some materials. Properties of PEEK Films”, Proc. of the 9th International The difference of the induced space Conference on “Protection of Materials and Structures environment data from the space environment from Space Environment,”, Canada, 2008,(2009) pp. monitoring samples and space environment model 127-136. was acquired. [11] Proc. of International Symposium on “SM/MPAC&SEED Experiment”, Japan, 2008, Acknowledgments JAXA-SP-08-015E (2009) pp. 81-90. The authors would like to thank Mr. Mitsuyasu Kato [12] Proc. of International Symposium on in JAXA, Ms. Chie Saito, of JAXA at that time, Mr. “SM/MPAC&SEED Experiment”, Japan, 2008, Yoshiaki Tachi, Mr. Toshihiko Inoue and Mr. Ichiro JAXA-SP-08-015E (2009) pp. 135-138. Yamagata, who are currently with the Japan Atomic [13] Proc. of International Symposium on Energy Agency and Mr. Akitoshi Yokota of JGC Corp. “SM/MPAC&SEED Experiment”, Japan, 2008, for their support and progress on this project. The JAXA-SP-08-015E (2009) pp. 121-125. authors would also like to thank Ms. Yukiko Yamaura in [14] Proc. of International Symposium on IHI Aerospace Co., Ltd. for development and primary “SM/MPAC&SEED Experiment”, Japan, 2008, check out. 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