Effect of LEO Environment on Composite Materials: Comparison
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Space weather in terms of low earth orbits has been characterized into seven main elements, namely microgravity, residual atmosphere, high vacuum, atomic oxygen, ultraviolet and ionization radiation, solar radiation, and space debris. Each element effects on polymers and composite materials. Quantification of these effects can be evaluated by understanding the mechanisms of material degradation caused by each environmental factor along with its synergetic effect. Hence, the design elements to mitigate the material degradation can be identified. Finally, a cause-and-effect diagram (Ishikawa diagram) is designed to characterize the important design elements required to investigate while choosing a material for a satellite’s structure. This will help the designers to develop experimental methodologies to test the composite material for its suitability against the space environment. Some available testing facilities will be discussed. Some potential polymers will also be suggested for further evaluation.

  • low earth orbit environmental exposure
  • material degradation
  • composite material
  • space debris

1. Introduction

Low Earth orbit (LEO) is closest to the Earth at an altitude of 160 km (thermosphere) to 1000 km. The International Space Station (ISS) orbits at 400 km in this region. Satellites are also launched in this region. The time period of a satellite for one circumnavigation of the earth in this orbit is approximately 90 min. These satellites generally perform tasks like telecommunication, science exploration etc. Low earth orbit is actually the part of the earth’s atmosphere with very few air molecules in the form of ions and minimal gravity. Solar radiation ionizes the gases in the atmosphere from 50 km to 965 km, which overlaps mesosphere and thermosphere. Gases stay in the form of ions rather than in molecules, for example, oxygen, hydrogen, helium ions, etc. Aurora light occurs in this region due to the presence of ions, magnetic storms and solar winds. The environment of the orbit varies with altitude. Hence, the material degradation phenomenon and design strategies for space structures also vary. Therefore, to select a suitable material for low earth orbit, the LEO environment and its effect on the material must be studied thoroughly. The LEO environment can be categorized into seven sub-categories, according to the NASA researcher guide [1].

2. High Proximity to the Earth

LEO is the nearest orbit of Earth and a part of the thermosphere, as mentioned earlier. In fact, the gravitational force at the International Space Station ISS is 89% of the force of gravity at sea level [1]. However, the centrifugal force due to orbital velocity balances the forces, creating a feeling of ‘weightlessness.’ The remaining part of the gravity is termed micro-gravity.
Close proximity also has an effect on the orbital lifetime. Microgravity, as well as the effect of residual atmosphere and associated effects, reduce the orbital energy, resulting in de-orbiting. For example, ISS loses its altitude at a rate of 90 m per day. The use of composite materials for satellites in this region is beneficial as they burn out easily in the earth’s atmosphere.
In micro-gravity, the flow of the liquid is dominated more by surface tension rather than gravity. This characteristic can be utilized for the wetting of fibers with resin. In orbit, the manufacturing of composite structures has less porosity due to the uniform wettability of fibers. A woven glass fiber specimen was cured with acrylic-based resin with UV polymer in zero gravity of Novespace’s Zero-G aircraft laboratory. These specimens have 2.5% porosity, whereas the same specimens consolidated in the same condition on Earth have 12% porosity [2]. These specimens have also shown higher failure stress, stiffness and higher failure strain. The curing of resin in microgravity requires less pressure to ensure uniform impregnation, which reduces effort, cost and time. Due to less contamination and voids by the curing process, in-space consolidated material shows better mechanical properties.

3. Residual Atmosphere

LEO is still a part of the Earth’s atmosphere. Therefore, it is often referred to as residual atmosphere. The residual atmosphere in LEO consists of molecules of hydrogen, helium, nitrogen and oxygen. At LEO, the residual atmosphere consists of 80% oxygen [3]. Due to the ionizing action of solar radiation, oxygen is in the form of ions, which is hazardous for polymeric composites. The concentration of atomic oxygen depends upon solar activities, which also varies over a period of 11 years. The residual atmosphere and its constituent elements vary with altitude. This is well explained in the handbook of environmental degradation of materials, which describes the distribution of several gases in number density in atom/cm3 over an altitude above 100 km or, in other words, in near-earth space [4]. According to this handbook, at low earth orbit (range 160–1000 km), oxygen shares ~80% of the atmospheric gases. Until 400 km, which is considered a very low earth orbit and a hotspot for launching commercial small satellites, the residual atmosphere is dominated by oxygen and nitrogen, whereas the upper part of LEO orbit (>650 km) is dominated by hydrogen, helium and oxygen [5].
These gas molecules are disassociated in the presence of ultraviolet rays and solar radiation. They stay in as free radical atoms or ions. These ions are traveling in LEO at a speed greater than 7 km/s. Hence, they might be highly reactive to polymeric material, and they oxidize the material, leading to corrosion. Traveling speed increases the rate of hitting ions, which erodes the material. A high rate of hitting speeds up the gradual process of corrosion even further. The surface erosion of the material ultimately creates dimensional instability of the structure. Atomic oxygen also changes the chemical composition of the polymer, which eventually degrades the mechanical strength of the structure.
The ionized gas exhibits a collective response to the magnetic and electric fields of the earth, solar and cosmic radiation. They create a charged field referred to as plasma. Plasma flux modifies the molecular structure of the polymer, resulting in embrittlement. A charged environment alters the optical properties of the thin film polymer or coating used for solar panels, mirrors, antennas, cameras, electronic equipment, etc. Hence, the residual atmosphere at the target orbit requires analysis for the proper selection of the polymer.

4. High Vacuum

Even though LEO has a residual atmosphere, there is an ultra-high vacuum, which ranges from 10−9 to 10−11 torr (1.3 × 10−6 mbar to 13 × 10−9 mbar) [1]. This high vacuum causes the outgassing of volatiles and moisture entrapped in the composite during manufacturing. Reducing the void content is challenging, especially for out-of-autoclave manufacturing techniques. Therefore, highly precise manufacturing is required to produce composite space structures. Otherwise, the gases will escape from the structure in space, which will contaminate other associated components and distort its geometry, which results in inaccuracy in terms of dimension, shape, mass and strength [6][7]. It also deteriorates the absorption and transmittance [8]. Therefore, material that shows low outgassing tendencies is preferred for space applications.

5. High Thermal Cycling

LEO lies in the thermosphere and exosphere of the earth’s atmospheric layers, where the temperature rises with altitude. The temperature of this layer depends on solar radiation, and it can raise the temperature of gases up to 1700 to 2000 °C. This high radiation makes the gas molecules electrically charged. However, gas density is extremely low due to high vacuum. The atmosphere temperature is usually lower than 0 °C. The possible source for heat transfer to objects in LEO is heat transfer from the gas molecules and radiation. Hence, the temperature of the satellite depends upon its position with respect to both the Earth and the sun.
Source of heat for satellites in orbit:
  • Direct sunlight;
  • Its view to the sun, i.e., sun side or shadow side;
  • Reflected sunlight from earth (Albedo);
  • Position of the earth with respect to the sun, i.e., aphelion or perihelion and position of the satellite with respect to the earth, i.e., perigee and apogee;
    Infrared radiation from Earth;
  • Internal heat produced by satellite component.
Variation of temperature on the satellite surface is due to its position:
  • Its view to the earth, i.e., earth’s shadow or earth-facing.
When the satellite is facing solar radiation directly at the sun’s perihelion and earth’s perigee position, its distance from the sun and earth is minimal, as illustrated in Figure 1 [9]. Hence, the surface of the satellite absorbs the highest amount of heat, whereas, at the sun’s aphelion and earth’s apogee, it absorbs the least. Therefore, the temperature gradient varies from −150 °C to 150 °C within 90 min of its orbital period. Hence, one part of the satellite facing the sun could experience 150° C, whereas another part of the same satellite in the shadow could face −150 °C at the same time. Due to the extreme temperature variation experienced by the same structure, thermal stresses occur. This temperature profile changes every 90 min, which develops thermal fatigue. Fatigue may cause micro-cracks, which reduce the mechanical strength of the material. These cracks, in turn, increase the possibility for atomic oxygen to penetrate inside the structure and erode it. Furthermore, in the presence of these micro-cracks, volatiles and moisture entrapped inside the structure escape easily. Hence, the mass reduces until all volatile components are escaped from the structure [10][11][12][13][14][15][16].
Jcs 07 00515 g001
Figure 1. Source of heat for a satellite.
When the structure is made of composite material that consists of fiber and matrix, both have different thermal expansion characteristics. CTE (coefficient of thermal expansion) mismatch, high-temperature variation in the same structure and cyclic thermal loading create high amplitude thermal fatigue. This develops additional micro-damages to the structure. However, this can be minimized by an optimized selection of fiber and polymer matrix. For example, choosing carbon fiber and PEEK polymer creates the overall CTE of the composite in the range of 0 to 8 × 10−6/°C [17]. The overall CTE of graphite epoxy can also be altered from −0.06 to −2 × 10−6/°C by changing the fiber layup [18]. Therefore, proper fiber and matrix selection are important to design satellite structures.

6. Ultra-Violet and Ionization Radiation

Sunlight radiates in LEO at its full irradiance due to the absence of the ozone layer. As shown in Figure 2 cited from [19], sunlight radiates energy with wavelengths of 120 nm to 3000 nm. Total solar radiation is about 1366 W/m2, in which the wavelength below 400 nm shares 8% of total radiance [20]. The UV radiation due to solar rays accounts for the wavelength from 40–400 nm, which is further classified between the UVA of wavelengths 315–400 nm, the UVB of whose wavelength is in the range of 280–315 nm and UVC, whose wavelength is in the range 100–280 nm. The energy of the photons is inversely proportional to the wavelength. UVA and UVB contain photon energies greater than 3.9 eV. This energy is sufficient to disassociate the chemical bonds of polymeric chains and produce free radicals [21]. The most significant damage by UV radiation to the composite material occurs in the range of 10–200 nm, which is also referred to as the vacuum UV (VUV) range.
Jcs 07 00515 g002
Figure 2. Spectral irradiance at low earth orbit.
This accounts for 0.1% of total irradiance of UV radiation and 0.007% of total solar radiation [19][22]. The energy of a photon in VUV radiation varies from 6 eV to 124 eV.
UV radiation is also absorbed by the residual atmosphere, which influences the plasma belt, high-energy protons and electrons whose energy can be up to several MeV [1]. Organic compounds are strong UV rays absorbents. It creates photochemical reactions. The thermal energy provided by the radiation breaks the molecular bonds, which eventually alters the molecular structure. These broken bonds can be disassociated from main molecular chains, then re-associate to another polymer chain or can become free radicals [23]. Ultra-violet rays and ionization can affect FRP in the following ways:
  • Material discoloration occurs due to the absorption of UV rays. The surface color of the material gets yellowed or dark. The polymer absorbs UV rays and undergoes several chemical reactions, such as dissociation of bonds, isomerization, free radical polymerization and contamination or recombination with other free radicals. This changes the mechanical properties of the structure.
  • It erodes the surface, which increases the surface roughness. Discoloration, in addition to surface roughness, deteriorates the thermo-optical properties, which decreases the efficiency of the thermal control surface. Hence, low solar absorption and high thermal emittance are required in the material to reduce this phenomenon.
  • The polymeric chain is de- and re-associated. Such cross-linking modifies the chemical structure of the polymer. Such modification leads to loss of mechanical, optical and chemical properties [11][23][24][25][26][27][28][29]. The polymeric chain with aromatic rings or phenyl rings may have higher erosion compared to polymers having long polymeric chains [30]. Aromatic rings or phenyl groups have more C=C bonds, which makes them stronger UV absorbents. Thus, there is a higher chance of chain scissioning and polymer fragmentation.
  • Polymers undergo embrittlement or chain scission in which polymeric bonds de-associate from polymeric chains and perform bonds with free ions. This free radical polymerization creates volatiles, which sublimate in the space. This phenomenon erodes the surface. The presence of a high vacuum increases this erosion even further. Reactive functional groups in the polymer chain sublimate faster in the environment [30].
  • Changes in the chemical structure of the polymer degrades the viscoelastic properties. UV radiation increases the glass transition temperature of epoxy-based shape memory polymer by 2.9% after exposition to UV radiation with a wavelength of 250–400 nm for 80 h [11]. It also increases the stiffness of the material by 41% after 80 hours of exposure. However, 80 h of exposure is too short to conclude the variation of glass transition temperature and stiffness. Hence, more polymers and longer exposure times are required for the investigation.

7. Atomic Oxygen

Due to UVC radiation from the sun, the oxygen atoms get disassociated into oxygen ions. These atomic oxygen ions travel in LEO at speeds higher than 7–8 km/s. As described earlier, 80% of the residual atmosphere consists of oxygen at LEO. Atomic oxygen (AO) flux is the most important factor in the degradation of spacecraft materials. The concentration of oxygen atoms at LEO ranges from 108 to 109 atoms/cm3 [31]. However, ISS receives an AO flux of about 5.23 × 1013 atoms/cm2/s, whose altitude is 400 km [32]. Atomic oxygen attacking the spacecraft at orbital velocity initiates several material degradation processes. For example, elastic scattering of the oxygen atoms leads to the removal of material through impact load. Kinetic energy possessed by AO at LEO is about 5 eV, which is higher than the bond dissociation energy of polymeric compounds [19][33]. Hence, AO reacts with the polymeric matrix and produces some gaseous oxides like CO2, H2O2, H2O, HCO2, CH3 and other organic compounds, which can be both volatile and non-volatile in nature. Due to huge heat transfers, it directly sublimates, leaving a crater at the surface, which leads to surface erosion [31][32][33][34][35]. Some possible degradation mechanisms that affect the composite material due to atomic oxygen are [31][32][33][34][35][36][37][38]:
  • Change in the chemical composition of the polymer, which leads to loss of mechanical properties;
  • Embrittlement and chain scission of polymeric chains;
  • Material erosion, which leads to loss of material, hence creates dimensional instability;
  • Loss of thermo-optical properties of the material (i.e., absorptance and emittance) due to material discoloration.
Selection of a suitable polymer that has less susceptibility to AO is very necessary in order to design an FRP structure for space [35]. Erosion yield is a criterion to determine suitable polymers.

8. Space Debris

Discarded rocket bodies and non-functional satellites dwell in space for many years. Due to collision, these objects are fragmented into even smaller parts, and the number of parts increases even further [39]. There are 34,000 objects larger than 10 cm, 900,000 larger than 1 cm, and 128 million larger than 1 mm floating in LEO orbit with a speed greater than 11 km/s. A total of 26,600 of them have identified objects evolving according to ESA [40]. Each launch introduces two times the amount of debris [41]—once at the start of a mission in the form of a rocket and another at the end of its operational life in the form of defunct satellites. With the emergence of the new space age with small satellites, the number of launches has increased at an exponential rate. In the report from Euroconsult 2020, it is mentioned that 1000 small satellites of mass less than 500 kg (20 satellites/week) have been launched in 2020 despite the coronavirus pandemic [42]. Consequently, in terms of rising object counts, the probability of catastrophic collisions will also grow. Despite the 25-year end-of-life (EOL) guidelines, 90% of satellites weighing 500–1000 kg and 60% of satellites weighing 100–500 kg would fail the 25-year EOL regulation based on an EOL estimation of satellites launched between 2015–2020 [43]. Thus, the low earth orbit is highly polluted with space debris. Seventy percent of space debris belongs to the low earth orbit [40][41][44][45][46][47].
Apart from space debris, there are micro-meteoroids that travel at speeds higher than 11 km/s. They can cause impact damage and sometimes, catastrophic failure if they strike functional satellites. In 2009, one non-functional satellite, Cosmos 2251 (space debris), collided with the functional satellite Iridium 33. This collision has created more than 1500 fragments. These fragments, in turn, increase the growth rate of space debris by 320 objects per year [48][49].
Under such a high impact velocity, the failure mechanism of the composite material transits between ductile to brittle fractures. Due to high impact velocity, there is less time for energy absorption for plastic deformation. This creates severe surface pitting, spalling of coating, cracking and delamination. Impact damage may short out the solar cells and sometimes destroy the functionality of the satellite. High velocity is associated with high kinetic energy, which transits into thermal energy. Hence, the viscoelasticity of the composite material changes due to the temperature rise. Consequently, the polymer softens, which reduces its stiffness. In thermoplastic material, the plastic flow in a softened polymer matrix near the delaminated region reduces the tendency to widen the matrix cracks. This inhibition of matrix cracks reduces the delamination area [50]. However, in thermoset materials, thermal energy reduces the structural integrity. Therefore, thermoplastic composites like PEEK/CF show much less pitting and delamination compared to epoxy/CF [51]. Similarly, the fabricating method also affects the impact damage. Braided composite structures provide more resistance against impact damage compared to unidirectional laminates due to the interlocking of two interlaced fiber bundles. This interlocking provides resistance against crack propagation. The network of interlaced fibers distributes the impact load evenly throughout the structure. This even stress distribution reduces the severity of local damage [52][53][54][55].

9. Conclusion

Even though polymers have been used since the beginning of the space age as polymer thin films in protective layers, coating, insulation, space suits, etc., researchers have limited material data. Since 2000, the new space age industries have focused more on reducing weight. They have already listed the benefits of composite structures in space. Nowadays, composite materials are used in space structures such as antennas, hinges, morphing wings, booms, solar arrays, struts and trusses, frames of satellites, battery casings, holding cases of satellite equipment like lenses, sensors, thrusters, reaction wheels, cameras, etc. Surprisingly, these composite parts are launched into space without a thorough examination of material data dependent on the space environmental conditions. Space regulations for cube satellites are also relaxed compared to those for large satellites. One possible reason behind this could be a lack of assessment of the risk factors that could lead to catastrophic failures of the satellites. Material erosion and degradation due to long-term exposure are not well understood, and there is a lack of data on material performance in relation to space environmental factors.

Studying material degradation in the space environment has many challenges. The initial challenge is to develop testing chambers equivalent to the orbit environment, which can replicate approximately vacuum, radiation and thermal cycles experienced in the orbit. Very few institutions have these facilities, but they are mostly in separate chambers situated in different locations. The costs and time required to obtain experimental data from these specialized chambers are significant, which further extends project durations and costs. Synergetic climatic chambers that combine multiple environmental factors could be a valuable addition, which not only reduces the time but also improves the accuracy of material data, as a synergetic environment is more detrimental to material degradation. However, it is a second challenge to combine all the factors of the orbit environment in one chamber, even with modern technology. The third challenge is to accelerate the process so that long-term material data could be generated in the shortest time period and at the lowest cost.

The synergistic effects of the LEO environment, including electron/proton and ultraviolet radiation combined with atomic oxygen and thermal cycling, can significantly accelerate mass erosion compared to individual factors. For example, the double-sided aluminized Kapton thermal blanket used in the ISS was completely damaged after one year of exposure even though the expected lifetime was 15 years, based on ground laboratory tests in which the specimen was irradiated by atomic oxygen with fluence level equivalent to 15 years [68,71]. The erosion rate of Kapton film in space (in orbit of the ISS) was 18 times higher than the results obtained in a ground test facility on Earth [2]. Therefore, material degradation should be investigated thoroughly for more polymer and composite materials. A comprehensive material database should be provided to the designers to create a responsible space structure that will not fail during the functional life of the satellite as well as does not stay longer in space as space debris.

References

  1. Finckenor, M.M.; Groh, K.K.D. Reseacher’s Guide to ISS: Space Environmental Effects; NASA: Washington, DC, USA, 2015.
  2. Sarantinos, N.; Loginos, P.; Charlaftis, P.; Argyropoulos, A.; Filinis, A.; Vrettos, K.; Adamos, L.; Kostopoulos, V. Behavior of photopolymer fiber structures in microgravity. SN Appl. Sci. 2019, 1, 1693.
  3. Banks, B.A.; Miller, K.R.S. Overview of Space Environment Effects on Materials and GRC’s Test Capabilities. In Proceedings of the 2005 NASA Seal/Secondary Air System Workshop, Cleveland, OH, USA, 8–9 November 2005; Glenn Research Center: Sandusky, OH, USA NASA/CP—2006-214383/VOL1.
  4. Banks, B.A. Atomic Oxygen. In Proceedings of the LDEF Materials Data Analysis Workshop, Merritt Island, FL, USA, 13–14 February 1990; p. 290.
  5. Samwel, S.W. Low Earth Orbital Atomic Oxygen Erosion Effect on Spacecraft Materials. Space Res. J. 2014, 7, 1–13.
  6. Walter, N.A.; Scialdone, J. Outgassing Data for Selecting Spacecraft Materials; NASA Reference Publication 1124; NASA: Washington, DC, USA, 1997; 444p.
  7. Anwar, A.; Elfiky, D.; Hassan, G.; Albona, M.; Marchetti, M. Outgassing Effect on Spacecraft Structure Materials. Int. J. Astron. Astrophys. Space Sci. 2015, 2, 34–38.
  8. Tribble, A.C. Designing for the space environment. Aerosp. Am. 1993, 31, 26–29.
  9. Maini, A.K.; Agrawal, V. Satellite Technology: Principles and Applications, 3rd ed.; Wiley: Chichester, WS, USA, 2014; ISBN 978-1-118-63647-3.
  10. Gao, Y.; He, S.; Yang, D.; Liu, Y.; Li, Z. Effect of vacuum thermo-cycling on physical properties of unidirectional M40J/AG-80 composites. Compos. Part B: Eng. 2005, 36, 351–358.
  11. Tan, Q.; Li, F.; Liu, L.; Liu, Y.; Yan, X.; Leng, J. Study of low earth orbit ultraviolet radiation and vacuum thermal cycling environment effects on epoxy-based shape memory polymer. J. Intell. Mater. Syst. Struct. 2019, 30, 2688–2696.
  12. Mahdavi, S. Thermal Cycling of Out-of-Autoclave Thermosetting Composite Materials. Master’s Thesis, Concordia University, Montreal, QC, Canada, 2017.
  13. Hegde, S.R.; Hojjati, M. Effect of Thermal Cycling on Composite Honeycomb Sandwich Structures for Space Applications; ASTRO Cancadian Aeronautics and Space Institute: Montréal, QC, Canada, 2018; p. 7.
  14. Gao, Y.; Dong, S.L.; Wang, H.S.; Lu, S.W.; Bao, J.W. Effect of Vacuum Thermal Cycling on Mechanical and Physical Properties of an Epoxy Matrix Composite. Adv. Mater. Res. 2011, 415–417, 2236–2239.
  15. Park, S.Y.; Choi, H.S.; Choi, W.J.; Kwon, H. Effect of vacuum thermal cyclic exposures on unidirectional carbon fiber/epoxy composites for low earth orbit space applications. Compos. Part B Eng. 2012, 43, 726–738.
  16. Shimokawa, T.; Katoh, H.; Hamaguchi, Y.; Sanbongi, S.; Mizuno, H.; Nakamura, H.; Asagumo, R.; Tamura, H. Effect of Thermal Cycling on Microcracking and Strength Degradation of High-Temperature Polymer Composite Materials for Use in Next-Generation SST Structures. J. Compos. Mater. 2002, 36, 885–895.
  17. Abdullah, F.; Okuyama, K.; Tapia, I.F. In-Situ Monitoring of Carbon Fiber/Polyether Ether Ketone (CF/PEEK) Composite Thermal Expansion in Low Earth Orbit. In Proceedings of the 8th International Symposium on Space Technology and Science, Tokyo, Japan, 17–21 June 2019.
  18. Lukez, R. The Use of Graphite/Epoxy Composite Structures in Space Applications; DigitalCommons@USU: Logan, UT, USA, 1987; p. 11.
  19. The Extraterrestrial (AM0) Solar Spectrum. Available online: https://www2.pvlighthouse.com.au/resources/courses/altermatt/The%20Solar%20Spectrum/The%20extraterrestrial%20%28AM0%29%20solar%20spectrum.aspx (accessed on 22 March 2023).
  20. ASTM E490-00a(2019); E21 Committee Standard Solar Constant and Zero Air Mass Solar Spectral Irradiance Tables. ASTM International: West Conshohocken, PA, USA, 2019.
  21. Yu-Ran Luo Bond Dissociation Energies 2010. Available online: http://staff.ustc.edu.cn/~luo971/2010-91-CRC-BDEs-Tables.pdf (accessed on 22 April 2021).
  22. ASTM E512-94(2020); E21 Committee ASTM E512 Practice for Combined, Simulated Space Environment Testing of Thermal Control Materials with Electromagnetic and Particulate Radiation. ASTM International: West Conshohocken, PA, USA, 2020.
  23. Dever, J.; Bruckner, E.; Rodriguez, E. Synergistic effects of ultraviolet radiation, thermal cycling and atomic oxygen on altered and coated Kapton surfaces. In Proceedings of the 30th Aerospace Sciences Meeting and Exhibit, Reno, NV, USA, 6–9 January 1992; American Institute of Aeronautics and Astronautics: Reston, VA, USA, 1992.
  24. Liau, W.B.; Tseng, F.P. The effect of long-term ultraviolet light irradiation on polymer matrix composites. Polym. Compos. 1998, 19, 440–445.
  25. Groh, K.K.D.; Smith, D.C. Investigation of Teflon FEP Embrittlement on Spacecraft in Low Earth Orbit; NASA Lewis Research Center: Cleveland, OH, USA, 1997.
  26. Ishizawa, J.; Mori, K. Space Environment Effects on Cross-linked ETFE Polymer. In Proceedings of the International Symposium on Materials in a Space Environment, Aix-en-Provence, France, 15–18 September 2009; p. 5.
  27. Grossman, E.; Gouzman, I. Space environment effects on polymers in low earth orbit. Nucl. Instrum. Methods Phys. Res. Sect. B Beam Interact. Mater. At. 2003, 208, 48–57.
  28. Awaja, F.; Moon, J.B.; Zhang, S.; Gilbert, M.; Kim, C.G.; Pigram, P.J. Surface molecular degradation of 3D glass polymer composite under low earth orbit simulated space environment. Polym. Degrad. Stab. 2010, 95, 987–996.
  29. Sznajder, M.; Renger, T.; Witzke, A.; Geppert, U.; Thornagel, R. Design and performance of a vacuum-UV simulator for material testing under space conditions. Adv. Space Res. 2013, 52, 1993–2005.
  30. Suliga, A. Resistance of CFRP Structures to Environmental Degradation in Low Earth Orbit. Ph.D. Thesis, University of Surrey, Guildford, UK, 2017.
  31. Bitetti, G.; Marchetti, M.; Mileti, S.; Valente, F.; Scaglione, S. Degradation of the surfaces exposed to the space environment. Acta Astronaut. 2007, 60, 166–174.
  32. De Groh, K.K.; Banks, B.A.; Mccarthy, C.E.; Rucker, R.N.; Roberts, L.M.; Berger, L.A. MISSE 2 PEACE Polymers Atomic Oxygen Erosion Experiment on the International Space Station. High Perform. Polym. 2008, 20, 388–409.
  33. Dooling, D.; Finckenor, M.M. Material Selection Guidelines to Limit Atomic Oxygen Effects on Spacecraft Surfaces; NASA Marshall Space Flight Center: Huntsville, AL, USA, 1999; p. 50. Available online: https://s3vi.ndc.nasa.gov/ssri-kb/static/resources/19990064119.pdf (accessed on 7 December 2021).
  34. Miller, S.K.R.; Banks, B. Degradation of Spacecraft Materials in the Space Environment. MRS Bull. 2010, 35, 20–24.
  35. Cool, G.R. Atomic Oxygen Erosion Resistance of Organic Polymers for Low Earth Orbit Spacecraft; National Library of Canada-Bibliothèque Nationale du Canada: Ottawa, ON, Canada, 1996.
  36. Devapal, D.; Packirisamy, S.; Korulla, R.M.; Ninan, K.N. Atomic oxygen resistant coating from poly(tetramethyldisilylene-co-styrene). J. Appl. Polym. Sci. 2004, 94, 2368–2375.
  37. Banks, B.A.; Backus, J.A.; Manno, M.V.; Waters, D.L.; Cameron, K.C.; Groh, K.K.D. Atomic Oxygen Erosion Yield Prediction for Spacecraft Polymers in Low Earth Orbit. In Proceedings of the 11th International Symposium on Materials in a Space Environment (ISMSE) 2009, Symposium and Exhibition NASA/TM--2002-211360, Aix en Provence, France, 15–18 September 2009.
  38. Banks, B.A.; Demko, R. Atomic Oxygen Protection of Materials in Low Earth Orbit. In Proceedings of the 2002 Symposium and Exhibition NASA/TM--2002-211360; Long Beach, CA, USA, 12–16 May 2002. Available online: http://gltrs.grc.nasa.gov/GLTRS (accessed on 5 June 2023).
  39. Kessler, D.J.; Johnson, N.L. The Kessler Syndrome: Implications to Future Space Operations. In Proceedings of the 33rd Annual AAS Guidance and Control Conference, Breckenridge, CO, USA, 6–10 February 2010.
  40. ESA UNOOSA ESA and UNOOSA Illustrate Space Debris Problem. Available online: https://www.esa.int/Safety_Security/Space_Debris/ESA_and_UNOOSA_illustrate_space_debris_problem (accessed on 18 May 2022).
  41. Self-cremating Satellite: A solution to reduce lifetime of space debris. In Proceedings of the European Interparliamentary Space Conference (EISC), Paris, France, 16 September 2022.
  42. EuroConsult. Prospects for Small Satellite Market Forecasts to 2029; EuroConsult: Washington, DC, USA, 2020; p. 7.
  43. Palla, C.; Kingston, J. Forecast analysis on satellites that need de-orbit technologies: Future scenarios for passive de-orbit devices. CEAS Space J 2016, 8, 191–200.
  44. TEC-QI, ESA. ESA Requirements on_End of Life to De-orbit. In Proceedings of the Technical Day on De-Orbit Strategies, Noordwijk, The Netherlands, 17 March 2015.
  45. ESA. ESA Technology Strategy 2019+; European Space Agency: Paris, France, 2019; p. 74.
  46. ESA Space Debris Office. ESA’s Annual Space Environment Report; ESA Space Debris Office; European Space Agency: Paris, France, 2020; GEN-DB-LOG-00288-OPS-SD.
  47. Flegel, S.K. Multi-layer insulation model for MASTER-2009. Acta Astronaut. 2011, 69, 911–922.
  48. Mejía-Kaiser, M. Collision Course: The 2009 Iridium -Cosmos Crash. In Proceedings of the 52nd Colloquium on the Law if Outer Space, Daejeon, Korea, 12 October 2009; Volume E8.3.9, p. 11. Available online: https://ssrn.com/abstract=3350010 (accessed on 7 December 2021).
  49. Walker, R.; Martin, C.; Strokes, H.; Wilkinson, J.; Sdunnus, H.; Hauptmann, S.; Beltrami, P.; Klinkrad, H. Update of the ESA Space Debris Mitigation Handbook; ESA: Paris, France, 2002.
  50. Dorey, G.; Bishop, S.M.; Curtis, P.T. On the Impact Performance of Carbon Fibre Laminates with Epoxy and PEEK Matrices. Compos. Sci. Technol. 1985, 23, 221–237.
  51. Graves, M.; Koontz, J. Initiation and extent of impact damage in graphite/epoxy and graphite/PEEK composites. In Proceedings of the 29th Structures, Structural Dynamics and Materials Conference, Williamsburg, VA, USA, 18–20 April 1988; American Institute of Aeronautics and Astronautics: Reston, VA, USA, 1988.
  52. Braid-Enhanced Composite Overwrapped Pressure Vessels |A&P Technology. Available online: https://www.braider.com/Case-Studies/Braid-Enhanced-Composite-Overwrapped-Pressure-Vessels.aspx (accessed on 7 December 2021).
  53. Kelkar, A.D.; Whitcomb, J.D. Characterization and Structural Behavior of Braided Composites; U.S. Department of Transportation: Washington, DC, USA, 2009.
  54. Malhotra, A.; Guild, F.J. Impact Damage to Composite Laminates: Effect of Impact Location. Appl. Compos. Mater. 2014, 21, 165–177.
  55. Metzner, C.; Gessler, A.; Weimer, C.; Beier, U.; Middendorf, P. Performance assessment on unidirectional braided CFRP materials. In Proceedings of the Low cost Composite Processing, from Aerospace OOA to Automotive Thermoplastics, Paris, France, 10–11 March 2014; SAMPE Europe: Paris, France, 2014.
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