Microgravity and Orbital Environment

Low Earth orbit is a free-fall environment in which weight is replaced by a small residual acceleration, and in which a surface is exposed to atomic oxygen, hard vacuum, trapped and galactic ionizing radiation, hypervelocity particle flux and a thermal cycle once per orbit.

Reference orbit

Quantity	Value	Source
ISS inclination	51.6 deg	[7]
ISS altitude band, assembly to assembly complete	350 to 450 km	[7]
Orbital translational velocity	7.7 km/s	[4]
Altitude band in which atomic oxygen dominates surface chemistry	180 to 650 km	[5]
Ambient pressure measured on LDEF	1e-4 to 1e-7 torr	[6]

Residual acceleration

Free fall is not zero acceleration. Experiment and life support equipment, crew activity, aerodynamic drag, gravity gradient and vehicle rotation all contribute, and the resulting field is treated as three components separated by frequency [1].

Component	Band	Spatial behavior	Source
Quasi-steady	below 0.01 Hz	Predictable at any point in the vehicle from a single measurement, the center of gravity location and rigid-body kinematics	[1]
Vibratory	0.01 to 300 Hz	Must be measured near the point of interest; propagation through structure is not reliably predictable	[1]
Transient	non-periodic, content across the whole spectrum	Must be measured near the point of interest	[1]

The quasi-steady band is bounded above by the lowest natural structural frequency of the vehicle, and its content varies over periods longer than a minute [1]. Vibratory content begins at that same structural frequency. Orbiter structural modes were identified at 3.5, 4.7, 5.1, 6.2 and 7.4 Hz, and the Spacelab module at 8.5 Hz [2].

Measured disturbance magnitudes:

Event	Magnitude	Duration	Source
ATV initial docking contact, 3 April 2008	12.9 mg	impulsive	[1]
ATV reboost	21 min burn, delta-v just over 4 m/s	21 min	[1]
Thruster firing transient	~20 mg	minutes	[2]
Crew push-off or impact	~50 mg	short	[2]
Payload-induced shock	50 to 55 mg	tens of ms	[2]
Water pump at 10,000 rpm (166.7 Hz)	4 to 5 mg	continuous	[2]
General equipment operation	0.3 to 0.7 mg	continuous	[2]
Crew exercise	band 0.01 to 5 Hz	daily	[2]

Reboost is not optional: the vehicle decelerates continuously under drag and the orbit decays [1]. A manipulator working at a berthing interface therefore does so against a base whose acceleration state includes both a periodic thrust-and-coast history and crew-driven transients at tens of milli-g.

Instrumentation defines what can be resolved. The Space Acceleration Measurement System uses an adjustable anti-alias cutoff between 25 and 400 Hz [1], the Microgravity Acceleration Measurement System HiRAP head reaches 100 Hz, and the JAXA Microgravity Measurement Apparatus in Kibo covers 10 to 300 Hz. SAMS-II remote triaxial sensor heads cover 0.01 Hz to beyond 300 Hz, extending to 400 Hz for transients, store 10 hours of data from five sensors at full bandwidth, and had returned 3.4 TB by Expedition 14 [3].

Neutral atmosphere and atomic oxygen

Solar ultraviolet below 243 nm exceeds the 5.12 eV dissociation energy of O2, so the residual atmosphere at orbital altitude is predominantly atomic oxygen [5]. At 7.7 km/s each ram-facing surface meets that population at a collision energy of 4.5 eV and a number flux of 5.23e13 atoms/cm2/s at ISS altitude [4]. Erosion is therefore strongly directional: ram surfaces erode, wake surfaces do not.

Ram, wake, zenith and nadir directions

Ram, wake, zenith and nadir directions relative to the velocity vector. Atomic oxygen erosion is confined to ram-facing surfaces because the 4.5 eV collision energy comes from the vehicle velocity rather than from thermal motion of the gas [5]. Source: [5]. Public domain (NASA).

Erosion yield Ey is the volume removed per incident atom. Flight values from MISSE 2 (3.95 years, 16 August 2001 to 30 July 2005, ram fluence 8.43e21 atoms/cm2) [5]:

Material	Erosion yield (cm3/atom)	Source
Polyoxymethylene	9.14e-24	[5]
Polycarbonate	4.29e-24	[4]
Epoxy, Hysol EA956	4.21e-24	[4]
Mylar, PET	3.01e-24	[4]
Kapton H (reference material)	3.00e-24	[5]
PEEK	2.99e-24	[4]
Kapton HN	2.81e-24	[4]
ECTFE, Halar	1.79e-24	[4]
ETFE, Tefzel	9.61e-25	[4]
FEP Teflon	2.00e-25	[4]
PTFE	1.42e-25	[4]
DC 93-500 silicone	3.81e-27	[5]

At the MISSE 2 fluence, a Kapton H surface loses about 25 micrometers [4]. Silicones survive because atomic oxygen converts the surface to a glassy silicate that passivates the layer below, at the cost of mud-tile cracking of that layer under strain [5]. Kapton H density of 1.4273 g/cm3 is the conversion used between measured mass loss and thickness loss.

The ground facility against orbit, on the same 39 polymers

A radio-frequency plasma asher is the ground facility that stands in for orbit when a material has to be screened before flight, and the MISSE 2 PEACE experiment is the one case where the same 39 polymers were run in both [20]. The asher yield was higher than the space yield for every one of the 39, by a factor between 1.02 and 37.1 [20]. Both halves are measurements: the flight yields are mass loss on vacuum-dehydrated samples of stated density and exposed area at one common fluence, and that fluence, the 8.43e21 atoms/cm2 above, was itself determined from the mass loss of two Kapton H witness samples on the tray rather than taken from an atmospheric model. The asher yields are relative to Kapton H, with the samples in metal holders and the asher operated on air.

The direction is systematic and the size is not. Ground atomic oxygen testing overestimates erosion for every material in this set, but by a material-dependent factor, so the overestimate cannot be divided out: the fluoropolymers sit at 6.1 to 8.0, Kevlar at 24.0 and white Tedlar at 37.1 [20]. Two consequences follow for a designer. An asher screening pass is a conservative bound on erosion and nothing more. And two materials ranked against each other by asher yield can be ranked wrongly, because a 37-fold error and a 1.02-fold error appear in the same table [20]. Kevlar is a woven fabric rather than a film, and the source reads its position near the top of that list as weave geometry being what a ground facility fails to reproduce. That is an inference drawn from the ranking, not a separate experiment.

Mass loss is not a usable erosion metric for silicones at all. A silicone typically does not lose weight in atomic oxygen, because the surface converts to the glassy SiOx layer described above, so a mass-loss ranking places silicones among the most durable materials while that layer is shrinking and cracking to expose the substrate; nanomechanical surface hardness is used instead to set ground-facility equivalent fluences [20]. What changes optically is measurable. An uncovered DC 93-500 silicone sample on the MISSE 2 ram tray went from a solar absorptance of 0.005 before flight to 0.033 to 0.037 after four years [5], a factor of about seven, with total reflectance barely changed and the surface crazed into mud tiles.

Ram fluence scales with mission duration and solar activity: MISSE 4 accumulated 2.15e21 atoms/cm2 in 1.04 years and MISSE 8 accumulated 4.62e21 atoms/cm2 in 2.14 years [5]. On LDEF, inorganic coatings including aluminum, nickel and SiO2 protected T300/934 graphite epoxy from atomic oxygen with negligible mass penalty [6]. MISSE 2 samples also accumulated 1400 equivalent sun hours of solar exposure, and erosion yield for several polymers rose with combined solar exposure and temperature rather than with atomic oxygen fluence alone [5].

Erosion yield of FEP against solar exposure

Erosion yield of FEP Teflon against accumulated solar exposure in equivalent sun hours, across several flight experiments. Yield rises with solar exposure, so an atomic oxygen fluence alone does not predict thickness loss for this material. Source: [5]. Public domain (NASA).

Thermal cycling

An orbiting structure passes through eclipse roughly every 90 minutes [6]. LDEF accumulated about 34,000 thermal cycles in 5.8 years, over a range of -20 to 160 F (about -29 to 71 C) with a +/-20 F band on those limits [6]. The recorded failure modes were microcracking of composites and delamination of some dielectric and metallic coatings; the aluminized graphite-epoxy noted above survived the full cycle count without coating delamination. Cycling acts synergistically with atomic oxygen, ultraviolet, vacuum and contamination rather than independently.

Heat rejection is radiative because there is no convecting medium. The ISS US On-Orbit Segment rejects heat through the External Active Thermal Control System, two independent mechanically pumped single-phase liquid ammonia loops, each serving three radiators of eight panels, with each radiator carrying two flow paths [16]. Radiator orientation is controlled by the Thermal Radiator Rotary Joint, either parked at a fixed angle or in Autotrack, which adjusts the angle to maximize heat rejection against the moving sink direction. A robot on the outside of such a vehicle has the same constraint: heat generated in an actuator leaves only by conduction to a radiating surface.

Ionizing radiation

Total dose in a 51.6 deg, 400 km orbit originates primarily in the trapped belts, with one solar particle event assumed to occur per mission and add to it [13]. Design limits for the LEO-ISS orbit give a daily integral trapped proton fluence of 1.66e8 protons/cm2 above 0.1 MeV and 5.40e7 protons/cm2 above 0.5 MeV, external to the vehicle. LDEF returned an integrated trapped electron and proton dose of 2.5e5 rad at the vehicle ends against less than 10 rad from cosmic rays over the same 5.8 years [6].

Single event effects, not total dose, set the avionics problem in LEO. ISS screens assembled articles rather than de-lidded parts, in 200 MeV proton beams: protons of 200 to 500 MeV pass through packages, boards and even box enclosures at nearly full energy, and roughly one in 289,000 of them strikes a nucleus inside a device and makes the short-range, high-LET reaction products that cause the upsets [10]. Heavy ion testing runs at linear energy transfer between 60 and 100 MeV cm2/mg at beam energies near 12 MeV/nucleon, and devices sensitive to single event latchup are screened against an LET threshold of 36 MeV cm2/mg [11]. Proton fluence for Orbiter avionics assessment was integrated over 10 to 200 MeV. Large solar energetic particle events deliver of order 2.5e7 protons/cm2 above 10 MeV [12].

Internal, or deep dielectric, charging is a separate mechanism from surface charging. Electrons near 100 keV penetrate 3 mils of aluminum and those near 1 MeV penetrate 80 mils, depositing charge in dielectrics behind the outer shell [9]. The design threshold derived from CRRES data is 2e10 electrons/cm2 accumulated over 10 hours, equivalent to an average flux of 5e5 electrons/cm2/s.

Plasma and surface charging

The ionosphere at ISS altitude has an electron density near 1e11 m-3 and an electron temperature near 990 K, measured simultaneously by the wide and narrow Langmuir probes of the Floating Potential Measurement Unit [8]. Those instruments cover 1e9 to 5e12 m-3 and 500 to 3000 K, with the floating potential probe spanning -180 to +180 V.

Two mechanisms drive the ISS structure negative. The 160 V photovoltaic array strings collect electrons at their exposed high-potential ends, and motion across the geomagnetic field induces a potential of about -10 V on the pressurized truss with a 30 V variation tip to tip across the arrays [7]. Structure potential is clamped by two plasma contactor units emitting below 0.1 A nominally, with a maximum observed discharge current of 100 mA per orbit. With the contactors off, the largest floating potential recorded by the Floating Potential Probe in 2001 was -23 V, at eclipse exit; the effective ion collection area of the ISS structure, the term that sets the electron loss balance, is 35 m2; and above about 160 V, snapover raises the conductivity of exposed array surfaces and increases collection [7].

Charging matters for robotics because an EVA suit, a grapple fixture or a manipulator end effector making first contact with a differently charged surface completes a circuit that was previously open.

Meteoroids and orbital debris

The meteoroid environment model MEM 3 covers particle masses from 1e-6 to 10 g, the range able to damage a spacecraft, and gives an average impact speed in low Earth orbit of 19 km/s [14]. Particles below 1e-6 g are assigned a density of 2 g/cm3. Showers contribute about 1 percent of total flux at the 1 microgram limit and less than 10 percent for particles below 0.5 cm, so the sporadic background dominates design [14].

LDEF recorded 34,336 micrometeoroid and debris impacts over 5.8 years [6]; the comparison of MEM against that record works from craters with lip diameters above 1000 microns, a limiting crater diameter of 750 microns [15]. ISS radiators carry MMOD strikes on all panels, and to date no strike has produced a leak in a radiator flow path [16]. Impacts are not uniformly distributed in time; the LDEF record showed them arriving in temporal bursts [6].

Flux depends on orientation as strongly as atomic oxygen does. A manipulator parked in one attitude for months accumulates a different exposure on each face.

Vacuum tribology

At 1e-4 to 1e-7 torr [6] there is no adsorbed water or oxide-replenishing atmosphere on a freshly exposed metal surface, and boundary lubrication must be supplied by a fluid that does not evaporate. Vapor pressure therefore sets the choice of lubricant before viscosity does [17].

Friction against ambient pressure is not monotonic, so a single vacuum figure does not describe a mechanism that sees a range of pressures. AISI 52100 steel sliding on itself gives a friction coefficient of 0.45 in laboratory air at 760 mm Hg, falls to about 0.2 between 1e-1 and 1e-2 mm Hg, and rises again to 0.375 at 5e-7 mm Hg, all measured on a 3/16 in radius hemispherical rider against a 2.5 in rotating disk at 1000 g load, 390 ft/min and room temperature [18]. The rise at the low end is the residual oxygen supply becoming too thin to re-oxidize metal transferred to the rider, and the whole curve sits six to eight orders of magnitude above the roughly 1e-14 mm Hg the same authors give for space.

Below the oxide the contact welds. Matched copper single crystals contacted at 1e-11 torr under a 50 g load with a 10 second dwell give an adhesion coefficient, breakaway force divided by applied load, of 1.02 before any sliding, and after sliding 0.735 cm at 0.001 cm/s the matched (100) pair exceeds 130 against 10.5 for the matched (111) pair [19]. The specimens are 99.999 percent metals cleaned in place by argon ion bombardment and the upper values are the limit of the apparatus rather than measurements, so the set bounds what a plated, machined and handled surface can do rather than predicting it. The ratio between the static and post-sliding values is why a fretting interface is treated as the cold welding risk in an orbital mechanism and a static clamped one is not.

Flight bears that distinction out. A French experiment on LDEF held aluminum, copper, titanium, stainless steel, copper beryllium, silver alloy and palladium washer pairs in static loaded contact at 25 to 135 MPa under Belleville springs for 69 months in low Earth orbit [21], with no sliding motion. Post-flight metallurgical examination of every flight and control pair found no cold welding event, and the ground controls held under vacuum for the same 69 months found none either; the experimenters attribute the null result to the oxide never being removed, because nothing slid [21]. The same report’s review of prior on-orbit experience found no documented significant cold welding failure on a US spacecraft, every reported seizure tracing to vibration-driven fretting or to galling after loss of lubricant. That review is a literature survey rather than a measurement.

Lubricant	Vapor pressure at 150 C (Pa)	Kinematic viscosity at 40 / 100 C (cSt)	Source
PFPAE 815Z	3.5e-6	not stated	[17]
Trisilahydrocarbon 2-94-96	3.1e-5	133 / 20	[17]
Pentasilahydrocarbon MJD990405	1.9e-5	206 / 31	[17]
Multiply alkylated cyclopentane, Pennzane P2001A	2.3e-3	108 / 15	[17]

Extrapolated to 25 C, the trisilahydrocarbon sits at 1.1e-8 Pa and pentasilahydrocarbon MJD991029 at 3.0e-12 Pa [17]. Film thickness under Hertzian contact depends on the pressure-viscosity coefficient, measured at 16 and 11 GPa-1 at 21 and 40 C for 2-94-96 against 11 and 8.5 GPa-1 for Pennzane P2001A.

Wear testing is run in vacuum because the result changes with ambient chemistry. Spiral orbit tribometer runs at below 1.3e-6 Pa, 200 rpm, 23 C, 1.5 GPa mean Hertzian stress and a 50 microgram oil charge per ball take a friction coefficient of 0.28 as the failure criterion [17]. In a reciprocating tribometer at 250 N, 0.7 GPa peak Hertzian stress, 6 Hz and 100 C, a tetrasilahydrocarbon produced a 0.71 mm2 wear scar against 1.02 mm2 for Pennzane 2001A. A bearing test at 1e-5 Pa, 600 rpm, 89 N axial load, 75 C and 20 mg free oil ran about 1000 hours on a pentasilahydrocarbon without measurable degradation.

Lubricated hardware returned from orbit

One set of lubricated hardware has come back. LDEF carried Castrol Braycote 601, a PTFE filled perfluoropolyether grease, on trailing edge drive shafts, ultraviolet exposed with negligible atomic oxygen [21]. The returned grease had a lower base oil viscosity than its control, read as ultraviolet chain scission of the polyether; infrared showed no new carbonyl and therefore no oxidation, but new peaks at 1100 to 1400 wavenumbers suggest the PTFE filler itself degraded, and new DSC endotherms appeared near 106 C and 211 C, the second unexplained. The grease had darkened to black and the cause was not identified [21].

Bonded dry film did worse than either. Everlube 620, a heat-cured MoS2 film in a proprietary organic binder, was gone after the same 69 months on the same trailing edge: only traces of MoS2 remained in the machining grooves, too little to test, the binder having been decomposed by ultraviolet and evaporated [21]. That is the one clear lubricant loss on LDEF and it happened where atomic oxygen fluence was negligible. Tungsten disulfide dry film on the grapple shafts, by contrast, was intact with no discernible difference between ram and trailing surfaces, but no surface analysis and no friction test was run on it, so it is a survival observation and not a tribological result.

A friction number did survive. DuPont Vespel SP-21, graphite filled polyimide, exposed on a trailing edge tray for 69 months, returned a coefficient of friction unchanged from its control at about 0.15 after run-in, over 10,000 cycles against two unflown controls [21]. The friction test was run on the ground in laboratory air, so the 0.15 carries the same room-temperature-air caveat as any bench figure and is not a vacuum value; what the flight establishes is that 69 months of orbital ultraviolet and vacuum did not change it [21].

References

McPherson, K., Kelly, E. and Keller, J. (2009). Acceleration Environment of the International Space Station . AIAA Aerospace Sciences Meeting including The New Horizons Forum and Aerospace Exposition, 20120012936. Source

BibTeX

@inproceedings{mcpherson2009acceleration,
  title = {Acceleration Environment of the International Space Station},
  author = {McPherson, Kevin and Kelly, Eric and Keller, Jennifer},
  booktitle = {AIAA Aerospace Sciences Meeting including The New Horizons Forum and Aerospace Exposition},
  number = {20120012936},
  institution = {NASA},
  year = {2009},
  doi = {10.2514/6.2009-957},
  abstract = {Measurement of the microgravity acceleration environment on the International Space Station has been accomplished by two accelerometer systems since 2001. The Microgravity Acceleration Measurement System records the quasi-steady microgravity environment, including the influences of aerodynamic drag, vehicle rotation, and venting effects. Measurement of the vibratory/transient regime, comprised of vehicle, crew, and equipment disturbances, has been accomplished by the Space Acceleration Measurement System-II. Until the arrival of the Columbus Orbital Facility and the Japanese Experiment Module, the location of these sensors, and therefore, the measurement of the microgravity acceleration environment, has been limited to within the United States Laboratory. Japanese Aerospace Exploration Agency has developed a vibratory acceleration measurement system called the Microgravity Measurement Apparatus which will be deployed within the Japanese Experiment Module to make distributed measurements of the Japanese Experiment Module's vibratory acceleration environment. Two Space Acceleration Measurement System sensors from the United States Laboratory will be re-deployed to support vibratory acceleration data measurement within the Columbus Orbital Facility. The additional measurement opportunities resulting from the arrival of these new laboratories allows Principal Investigators with facilities located in these International Space Station research laboratories to obtain microgravity acceleration data in support of their sensitive experiments. The Principal Investigator Microgravity Services project, at NASA Glenn Research Center, in Cleveland, Ohio, has supported acceleration measurement systems and the microgravity scientific community through the processing, characterization, distribution, and archival of the microgravity acceleration data obtained from the International Space Station acceleration measurement systems. This paper summarizes the PIMS capabilities available to the International Space Station scientific community, introduces plans for extending microgravity analysis results to the newly arrived scientific laboratories, and provides summary information for known microgravity environment disturbers.}
}

DeLombard, R., McPherson, K., Hrovat, K., Moskowitz, M., Rogers, M. J. B. and Reckart, T. (1997). Microgravity Environment Description Handbook . NASA, NASA-TM-107486. Source

BibTeX

@techreport{delombard1997microgravity,
  title = {Microgravity Environment Description Handbook},
  author = {DeLombard, Richard and McPherson, Kevin and Hrovat, Kenneth and Moskowitz, Milton and Rogers, Melissa J. B. and Reckart, Timothy},
  number = {NASA-TM-107486},
  institution = {NASA},
  year = {1997},
  url = {https://ntrs.nasa.gov/citations/19970022250},
  abstract = {The Microgravity Measurement and Analysis Project (MMAP) at the NASA Lewis Research Center (LeRC) manages the Space Acceleration Measurement System (SAMS) and the Orbital Acceleration Research Experiment (OARE) instruments to measure the microgravity environment on orbiting space laboratories. These laboratories include the Spacelab payloads on the shuttle, the SPACEHAB module on the shuttle, the middeck area of the shuttle, and Russia's Mir space station. Experiments are performed in these laboratories to investigate scientific principles in the near-absence of gravity. The microgravity environment desired for most experiments would have zero acceleration across all frequency bands or a true weightless condition. This is not possible due to the nature of spaceflight where there are numerous factors which introduce accelerations to the environment. This handbook presents an overview of the major microgravity environment disturbances of these laboratories. These disturbances are characterized by their source (where known), their magnitude, frequency and duration, and their effect on the microgravity environment. Each disturbance is characterized on a single page for ease in understanding the effect of a particular disturbance. The handbook also contains a brief description of each laboratory.}
}

Foster, W. (2009). Space Acceleration Measurement System-II . NASA, 20090014818. Source

BibTeX

@techreport{foster2009space,
  title = {Space Acceleration Measurement System-II},
  author = {Foster, William},
  number = {20090014818},
  institution = {NASA},
  year = {2009},
  url = {https://ntrs.nasa.gov/citations/20090014818},
  abstract = {Space Acceleration Measurement System (SAMS-II) is an ongoing study of the small forces (vibrations and accelerations) on the ISS that result from the operation of hardware, crew activities, as well as dockings and maneuvering. Results will be used to generalize the types of vibrations affecting vibration-sensitive experiments. Investigators seek to better understand the vibration environment on the space station to enable future research.}
}

deGroh, K. K., Banks, B. A., McCarthy, C. E., Rucker, R. N., Roberts, L. M. and Berger, L. A. (2006). MISSE PEACE Polymers Atomic Oxygen Erosion Results . MISSE Post-Retrieval Conference, NASA/TM-2006-214482. Source

BibTeX

@inproceedings{degroh2006misse,
  title = {MISSE PEACE Polymers Atomic Oxygen Erosion Results},
  author = {deGroh, Kim K. and Banks, Bruce A. and McCarthy, Catherine E. and Rucker, Rochelle N. and Roberts, Lily M. and Berger, Lauren A.},
  booktitle = {MISSE Post-Retrieval Conference},
  number = {NASA/TM-2006-214482},
  institution = {NASA},
  address = {Orlando, FL},
  year = {2006},
  url = {https://ntrs.nasa.gov/citations/20070002707},
  abstract = {Forty-one different polymer samples, collectively called the Polymer Erosion and Contamination Experiment (PEACE) Polymers, have been exposed to the low Earth orbit (LEO) environment on the exterior of the International Space Station (ISS) for nearly 4 years as part of Materials International Space Station Experiment 2 (MISSE 2). The objective of the PEACE Polymers experiment was to determine the atomic oxygen erosion yield of a wide variety of polymeric materials after long term exposure to the space environment. The polymers range from those commonly used for spacecraft applications, such as Teflon (DuPont) FEP, to more recently developed polymers, such as high temperature polyimide PMR (polymerization of monomer reactants). Additional polymers were included to explore erosion yield dependence upon chemical composition. The MISSE PEACE Polymers experiment was flown in MISSE Passive Experiment Carrier 2 (PEC 2), tray 1, on the exterior of the ISS Quest Airlock and was exposed to atomic oxygen along with solar and charged particle radiation. MISSE 2 was successfully retrieved during a space walk on July 30, 2005, during Discovery s STS-114 Return to Flight mission. Details on the specific polymers flown, flight sample fabrication, pre-flight and post-flight characterization techniques, and atomic oxygen fluence calculations are discussed along with a summary of the atomic oxygen erosion yield results. The MISSE 2 PEACE Polymers experiment is unique because it has the widest variety of polymers flown in LEO for a long duration and provides extremely valuable erosion yield data for spacecraft design purposes.}
}

de Groh, K. K. and Banks, B. A. (2019). Atomic Oxygen Erosion Data from the MISSE 2-8 Missions . NASA, 20190025445. Source

BibTeX

@techreport{degroh2019atomic,
  title = {Atomic Oxygen Erosion Data from the MISSE 2-8 Missions},
  author = {de Groh, Kim K. and Banks, Bruce A.},
  number = {20190025445},
  institution = {NASA},
  year = {2019},
  url = {https://ntrs.nasa.gov/citations/20190025445},
  abstract = {Polymers and other oxidizable materials on the exterior of spacecraft in the low Earth orbit (LEO) space environment can be eroded from reaction with atomic oxygen (AO). Therefore, in order to design durable spacecraft it is important to know the extent of erosion that will occur during a mission. This can be determined by knowing the LEO AO erosion yield, E(sub y) (volume loss per incident oxygen atom), of materials susceptible to AO reaction. In addition, recent flight experiments have shown that the AO E(sub y) can vary with the AO fluence and/or solar exposure. Therefore obtaining AO E(sub y) data for materials flown on various spaceflight missions is important. NASA Glenn Research Center has flown numerous experiments as part of the Materials International Space Station Experiment (MISSE) missions on the exterior of the International Space Station to characterize the LEO E(sub y) of polymers, composites, protective coatings, and other spacecraft materials. This report provides a summary of the erosion data for ram samples from six Glenn polymer experiments flown as part of MISSE 2, 4, 6, 7, and 8. A total of 71 types of materials with 111 E(sub y) values are provided. The E(sub y) values for uncoated polymers range from 3.81×10(exp –27) cu cm/atom for DC 93-500 silicone exposed to an AO fluence of 4.62×10(exp 21) atoms/sq cm on MISSE 8 to 9.14×10(exp –24) cu cm/atom for polyoxymethylene (POM) exposed to an AO fluence of 8.43×10(exp 21) atoms/sq cm on MISSE 2. One polymer, Triton oxygen resistant, low modulus (TOR(TM) LM), experienced mass gain when exposed to an AO fluence of 2.15×10(exp 21) atoms/sq cm on MISSE 4. In many cases the same material was flown on numerous missions so that trends for E(sub y) versus AO fluence and/or solar exposure can be determined, along with temperature effects. }
}

Stein, B. A. (1992). An interim overview of LDEF materials findings . NASA, NASA-TM-107664. Source

BibTeX

@techreport{stein1992interim,
  title = {An interim overview of LDEF materials findings},
  author = {Stein, Brad A.},
  number = {NASA-TM-107664},
  institution = {NASA},
  year = {1992},
  url = {https://ntrs.nasa.gov/citations/19930009140},
  abstract = {The flight and retrieval of the National Aeronautics and Space Administration's Long Duration Exposure Facility (LDEF) provided an opportunity for the study of the low-Earth orbit (LEO) environment and long-duration space environmental effects (SEE) on materials that is unparalleled in the history of the U.S. Space Program. The remarkable flight attitude stability of LDEF enables specific analyses of various individual and combined effects of LEO environmental parameters on identical materials on the same space vehicle. This paper provides an overview of the interim LDEF materials findings of the Principal Investigators and the Materials Special Investigation Group. In general, the LDEF data is remarkably consistent; LDEF will provide a 'benchmark' for materials design data bases for satellites in low-Earth orbit. Some materials were identified to be encouragingly resistant to LEO SEE for 5.8 years; other 'space qualified' materials displayed significant environmental degradation. Molecular contamination was widespread; LDEF offers an unprecedented opportunity to provide a unified perspective of unmanned LEO spacecraft contamination mechanisms. New material development requirements for long-term LEO missions have been identified and current ground simulation testing methods/data for new, durable materials concepts can be validated with LDEF results. LDEF findings are already being integrated into the design of Space Station Freedom.}
}

Barsamian, H., Mikatarian, R. R., Alred, J., Minow, J. and Koontz, S. (2004). ISS Plasma Interaction: Measurements and Modeling . Spacecraft Charging and Technology Conference, 20040111048. Source

BibTeX

@inproceedings{barsamian2004iss,
  title = {ISS Plasma Interaction: Measurements and Modeling},
  author = {Barsamian, H. and Mikatarian, Ronald R. and Alred, J. and Minow, J. and Koontz, S.},
  booktitle = {Spacecraft Charging and Technology Conference},
  number = {20040111048},
  institution = {NASA},
  year = {2004},
  url = {https://ntrs.nasa.gov/citations/20040111048},
  abstract = {Ionospheric plasma interaction effects on the International Space Station are discussed in the following paper. The large structure and high voltage arrays of the ISS represent a complex system interacting with LEO plasma. Discharge current measurements made by the Plasma Contactor Units and potential measurements made by the Floating Potential Probe delineate charging and magnetic induction effects on the ISS. Based on theoretical and physical understanding of the interaction phenomena, a model of ISS plasma interaction has been developed. The model includes magnetic induction effects, interaction of the high voltage solar arrays with ionospheric plasma, and accounts for other conductive areas on the ISS. Based on these phenomena, the Plasma Interaction Model has been developed. Limited verification of the model has been performed by comparison of Floating Potential Probe measurement data to simulations. The ISS plasma interaction model will be further tested and verified as measurements from the Floating Potential Measurement Unit become available, and construction of the ISS continues.}
}

Wright, K. H. J., Swenson, C., Thompson, D., Barjatya, A., Koontz, S. L., Schneider, T., Vaughn, J., Minow, J., Craven, P., Coffey, V., Parker, L. and Bui, T. (2007). Initial Results from the Floating Potential Measurement Unit aboard the International Space Station . Spacecraft Charging and Technology Conference, 20070032723. Source

BibTeX

@inproceedings{wright2007initial,
  title = {Initial Results from the Floating Potential Measurement Unit aboard the International Space Station},
  author = {Wright, Kenneth H., Jr. and Swenson, Charles and Thompson, Don and Barjatya, Aroh and Koontz, Steven L. and Schneider, Todd and Vaughn, Jason and Minow, Joseph and Craven, Paul and Coffey, Victoria and Parker, Linda and Bui, Them},
  booktitle = {Spacecraft Charging and Technology Conference},
  number = {20070032723},
  institution = {NASA},
  address = {Biarritz},
  year = {2007},
  url = {https://ntrs.nasa.gov/citations/20070032723},
  abstract = {The Floating Potential Measurement Unit (FPMU) is a multi-probe package designed to measure the floating potential of the 1nternational Space Station (ISS) as well as the density and temperature of the local ionospheric plasma environment. The role oj the FPMU is to provide direct measurements of ISS spacecraft charging as continuing construction leads to dramatic changes in ISS size and configuration. FPMU data are used for refinement and validation of the ISS spacecraft charging models used to evaluate the severity and frequency of occurrence of ISS charging hazards. The FPMU data and the models are also used to evaluate the effectiveness of proposed hazard controls. The FPMU consists of four probes: a floating potential probe, two Langmuir probes. and a plasma impedance probe. These probes measure the floating potential of the ISS, plasma density, and electron temperature. Redundant measurements using different probes support data validation by inter-probe comparisons. The FPMU was installed by ISS crewmembers, during an ExtraVehicular Activity, on the starboard (Sl) truss of the ISS in early August 2006, when the ISS incorporated only one 160V US photovoltaic (PV) array module. The first data campaign began a few hours after installation and continued for over five days. Additional data campaigns were completed in 2007 after a second 160V US PV array module was added to the ISS. This paper discusses the general performance characteristics of the FPMU as integrated on ISS, the functional performance of each probe, the charging behavior of the ISS before and after the addition of a second 160V US PV array module, and initial results from model comparisons.}
}

Whittlesey, A. and Garrett, H. B. (1996). NASA's Technical Handbook for Avoiding On-Orbit ESD Anomalies Due to Internal Charging Effects . NASA, 20000055759. Source

BibTeX

@techreport{whittlesey1996nasa,
  title = {NASA's Technical Handbook for Avoiding On-Orbit ESD Anomalies Due to Internal Charging Effects},
  author = {Whittlesey, Albert and Garrett, Henry B.},
  number = {20000055759},
  institution = {NASA},
  year = {1996},
  url = {https://ntrs.nasa.gov/citations/20000055759},
  abstract = {This paper describes NASA-HDBK-4002, "Avoiding Problems Caused by Spacecraft On-Orbit Internal Charging Effects". The handbook includes a description of internal charging and why it is of concern to spacecraft designers. It also suggests how to determine when a project needs to consider internal spacecraft charging, it contains an electron penetration depth chart, rationale for a critical electron flux criterion, a worst-case geosynchronous electron plasma spectrum, general design guidelines, quantitative design guidelines, and a typical materials characteristics list. Appendices include a listing of some environment codes, electron transport codes, a discussion of geostationary electron plasma environments, a brief description of electron beam and other materials tests, and transient susceptibility tests. The handbook will be in the web page, hftp://standards.nasa.gov. A prior document, NASA TP2361 "Design Guidelines for Assessing and controlling Spacecraft Charging Effects", 1984, is in use to describe mitigation techniques for the effects of surface charging of satellites in space plasma environments. HDBK-4002 is meant to complement 2361 and together, the pair of documents describe both cause and mitigation designs for problems caused by energetic space plasmas.}
}

Koontz, S. L., Suggs, R. M., Alred, J. W., Worthy, E. S., Boeder, P., Steagall, C. A., Hartman, W. A., Gingras, B. D. and Schmidl, W. D. (2018). The International Space Station Space Radiation Environment: Avionics Systems Performance in Low-Earth Orbit Single Event Effects (SEE) Environments . International Conference on Environmental Systems, ICES-2018-69. Source

BibTeX

@inproceedings{koontz2018international,
  title = {The International Space Station Space Radiation Environment: Avionics Systems Performance in Low-Earth Orbit Single Event Effects (SEE) Environments},
  author = {Koontz, Steven L. and Suggs, Robert M. and Alred, John W. and Worthy, Erica S. and Boeder, Paul and Steagall, Courtney A. and Hartman, William A. and Gingras, Benjamin D. and Schmidl, William D.},
  booktitle = {International Conference on Environmental Systems},
  number = {ICES-2018-69},
  address = {Albuquerque, New Mexico},
  year = {2018},
  url = {https://ttu-ir.tdl.org/items/7fb5d403-ad77-4f6b-9cff-c70af5c04802}
}

Reddell, B. D. (1999). Orbiter Avionics Radiation Handbook . NASA, 20180002644. Source

BibTeX

@techreport{reddell1999orbiter,
  title = {Orbiter Avionics Radiation Handbook},
  author = {Reddell, Brandon D.},
  number = {20180002644},
  institution = {NASA},
  year = {1999},
  url = {https://ntrs.nasa.gov/citations/20180002644},
  abstract = {This handbook was assembled to document he radiation environment for design of Orbiter avionics. It also maps the environment through vehicle shielding and mission usage into discrete requirements such as total dose. Some details of analytical techniques for calculating radiation effects are provided. It is anticipated that appropriate portions of this document will be added to formal program specifications. }
}

Adams, J. H. J. (1988). Current models of the intensely ionizing particle environment in space . NASA, 19890019087. Source

BibTeX

@techreport{adams1988current,
  title = {Current models of the intensely ionizing particle environment in space},
  author = {Adams, James H., Jr.},
  number = {19890019087},
  institution = {NASA},
  year = {1988},
  url = {https://ntrs.nasa.gov/citations/19890019087},
  abstract = {The Cosmic Ray Effects on MicroElectronics (CREME) model that is currently in use to estimate single event effect rates in spacecraft is described. The CREME model provides a description of the radiation environment in interplanetary space near the orbit of the earth that contains no major deficiencies. The accuracy of the galactic cosmic ray model is limited by the uncertainties in solar modulation. The model for solar energetic particles could be improved by making use of all the data that has been collected on solar energetic particle events. There remain major uncertainties about the environment within the earth's magnetosphere, because of the uncertainties over the charge states of the heavy ions in the anomalous component and solar flares, and because of trapped heavy ions. The present CREME model is valid only at 1 AU, but it could be extended to other parts of the heliosphere. There is considerable data on the radiation environment from 0.2 to 35 AU in the ecliptic plane. This data could be used to extend the CREME model.}
}

NASA. (2020). Cross-Program Design Specification for Natural Environments (DSNE), Revision G . NASA Marshall Space Flight Center. Source

BibTeX

@techreport{nasa2020cross,
  title = {Cross-Program Design Specification for Natural Environments (DSNE), Revision G},
  author = {{NASA}},
  institution = {NASA Marshall Space Flight Center},
  year = {2020},
  url = {https://ntrs.nasa.gov/citations/20200000867},
  abstract = {The DSNE completes environment-related specifications for architecture, system-level, and lower-tier documents by specifying the ranges of environmental conditions that must be accounted for by NASA ESD Programs. To assure clarity and consistency, and to prevent requirements documents from becoming cluttered with extensive amounts of technical material, natural environment specifications have been compiled into this document. The intent is to keep a unified specification for natural environments that each Program calls out for appropriate application.}
}

Moorhead, A. V. (2020). NASA Meteoroid Engineering Model (MEM) Version 3 . NASA, NASA/TM-2020-220555. Source

BibTeX

@techreport{moorhead2020nasa,
  title = {NASA Meteoroid Engineering Model (MEM) Version 3},
  author = {Moorhead, A. V.},
  number = {NASA/TM-2020-220555},
  institution = {NASA},
  year = {2020},
  url = {https://ntrs.nasa.gov/citations/20200000563},
  abstract = {The Meteoroid Engineering Model (MEM) version 3 is NASA’s most current and accurate model of the meteoroid environment. MEM 3 supersedes all previous versions of MEM, including MEM Release 2.0 (MEMR2), MEM Release 1.0c (MEMR1c), and previously internally controlled and released versions of MEMCxP v2.0 and LunarMEM v2.0. Earlier versions of MEM superseded older models of the meteoroid environment such as the Grün model and its derivative, Technical Memo 4527 (hereafter abbreviated as TM 4527) [1]. Prior to the establishment of the NASA Meteoroid Environment Office (MEO), NASA’s meteoroid environment models relied on a simple empirical expression derived from [2], as described in [3] and later in [1]. This expression describes the meteoroid flux incident on a flat plate near 1 au. TM 4527 assumes an isotropic environment, making the orientation of the plate irrelevant [4]. The flux was combined with scale factors to account for the reduction in flux occurring when the Earth shields the spacecraft from a portion of the meteoroid environment and the enhancement in flux due to the focusing effect of Earth’s gravitational field. TM 4527 also introduced a crude, piecewise meteoroid speed distribution with an average velocity of 19 km/s for an orbiting spacecraft based on [5]. Finally, TM 4527 assumed a three-step density distribution in which dust particles smaller than 10−6 g have a density of 2 g/cu cm, micrometeoroids between 10−6 g and 0.01 g have a density of 1 g/cu cm, and meteoroids larger than 0.01 g have a density of 0.5 g/cm3. Thus, the meteoroid model presented in TM 4527 was assembled from multiple independent sources. The model of TM 4527 was also used for years in Space Station risk assessments, and is described in Space Station Specification (SSP) 30425.}
}

Ehlert, S., Moorhead, A. and Cooke, W. J. (2017). A Comparison of Results From NASA's Meteoroid Engineering Model to the LDEF Cratering Record . European Conference on Space Debris, 20170004455. Source

BibTeX

@inproceedings{ehlert2017comparison,
  title = {A Comparison of Results From NASA's Meteoroid Engineering Model to the LDEF Cratering Record},
  author = {Ehlert, S. and Moorhead, A. and Cooke, W. J.},
  booktitle = {European Conference on Space Debris},
  number = {20170004455},
  organization = {ESA Space Debris Office},
  address = {Darmstadt},
  year = {2017},
  url = {https://ntrs.nasa.gov/citations/20170004455},
  abstract = {NASA's Long Duration Exposure Facility (LDEF) has provided an extensive record of the meteoroid environment in low Earth orbit. LDEF's combination of fixed orientation, large collecting area, and long lifetime imposes constraints on the absolute flux of potentially hazardous meteoroids. The relative impact rate on each of LDEF's fourteen surfaces arises from the underlying velocity distribution and directionality of the meteoroid environment. For the first time, we model the meteoroid environment encountered by LDEF over its operational lifetime using NASA's Meteoroid Engineering Model Release 2 (MEMR2) and compare the model results with the observed craters of potentially hazardous meteoroids (i.e. crater diameters larger than approximately 0.75 mm). We discuss the extent to which the observations and model agree and how the impact rates across all of the LDEF surfaces may be utilized to help calibrate future versions of MEM.}
}

Rodriguez, A. and Morton, R. (2025). ISS Radiator Face Sheet Anomaly Investigation and Return to Function . International Conference on Environmental Systems, 20260002946. Source

BibTeX

@inproceedings{rodriguez2025iss,
  title = {ISS Radiator Face Sheet Anomaly Investigation and Return to Function},
  author = {Rodriguez, Aaron and Morton, Richard},
  booktitle = {International Conference on Environmental Systems},
  number = {20260002946},
  institution = {NASA},
  year = {2025},
  doi = {10.32865/2346/108957},
  abstract = {Maintaining sufficient heat rejection on the International Space Station (ISS) is critical to the function of Low Earth Orbit station. The External Thermal Control System (EATCS) rejects the heat generated by the US On-orbit Segment (USOS) modules. The system uses ammonia to collect the heat and reject it to 6 radiators (3 on the Starboard side, 3 on the Port side). Each radiator is made up of 8 panels. In September of 2008, imagery of the Starboard Radiators was conducted and showed that one of the panels’ face sheets had peeled up from the internal honeycomb core structure. Out of an abundance of caution, the radiator was isolated from the rest of the EATCS and the ammonia was vented to space. Since the radiator was vented, periodic imagery of all radiator panels has taken place to monitor any changes in the face sheet and inspect for other radiator anomalies. In January 2025, after years of trending and inspections, the radiator will be reintegrated into the system and will reject heat. This paper will document the multi-year investigation that took place to determine root cause and mitigations implemented to reduce risk to the system. A summary of the data since the investigation will show rationale for reintegrating the radiator even with the damaged panel.}
}

Jones, W. R. J., Jansen, M. J., Gschwender, L. J., Snyder, C. E. J., Sharma, S. K., Predmore, R. E. and Dube, M. J. (2001). The Tribological Properties of Several Silahydrocarbons for Use in Space Mechanisms . Journal of Synthetic Lubrication, NASA/TM-2001-211196. Source

BibTeX

@article{jones2001tribological,
  title = {The Tribological Properties of Several Silahydrocarbons for Use in Space Mechanisms},
  author = {Jones, W. R., Jr. and Jansen, M. J. and Gschwender, L. J. and Snyder, C. E., Jr. and Sharma, Shiv K. and Predmore, R. E. and Dube, Michael J.},
  journal = {Journal of Synthetic Lubrication},
  volume = {20},
  number = {NASA/TM-2001-211196},
  pages = {303-315},
  institution = {NASA},
  address = {Liege},
  year = {2001},
  doi = {10.1002/jsl.3000200404},
  abstract = {Abstract Silahydrocarbons are members of a relatively new class of liquid lubricants with great potential for use in space mechanisms. They are unimolecular species consisting of silicon, carbon, and hydrogen. They possess unique wear, viscosity, and volatility properties while retaining the ability to solubilise conventional additives. The tribological properties of several members of this class, including tri‐, tetra‐, and penta‐compounds, are presented. These properties include viscosity‐temperature, viscosity—pressure, vapour pressure, lubricant life, traction, and reciprocating and four‐ball wear rates. Lubricant lifetimes were determined using a vacuum ball bearing simulator, the spiral orbit tribometer. Wear was measured using a Cameron Plint reciprocating tribometer and wear rates with a vacuum four‐ball tribometer. Conventional viscometry was used for viscosity—temperature measurements and a Knudsen cell for vapour pressure. Thermogravimetric analysis was also used for volatility measurements. Pressure—viscosity coefficients (α—values) were estimated from elastohydrodynamic lubrication film thickness measurements. These properties are compared to those of existing state‐of‐the‐art space lubricants.}
}

Buckley, D. H., Swikert, M. and Johnson, R. L. (1962). Friction, Wear, and Evaporation Rates of Various Materials in Vacuum to 1e-7 mm Hg . ASLE Transactions. Source

BibTeX

@article{buckley1962friction,
  title = {Friction, Wear, and Evaporation Rates of Various Materials in Vacuum to 1e-7 mm Hg},
  author = {Buckley, Donald H. and Swikert, Max and Johnson, Robert L.},
  journal = {ASLE Transactions},
  volume = {5},
  pages = {8--23},
  year = {1962},
  url = {https://ntrs.nasa.gov/citations/20150020881},
  abstract = {The requirements for bearings and seals to operate in the environment of space dictate a new area for lubrication research. The low ambient pressures encountered in space can be expected to influence the behavior of oil, grease, and solid-film lubricants. The property of these materials most significantly affected by low ambient pressures is the evaporation rate. Various investigators have therefore measured the evaporation rates of oils and greases in vacuum as one method of establishing their relative merit for space applications (1-3). The results of this work have given some indication as to the oils and greases with the greatest stability at reduced ambient pressures. Only limited experimental work, however, has been reported in the literature for inorganic solids and soft metals which have potential use as solid lubricant films or coatings for hard alloy substrates [e.g. Reference ( 4 )]. In general, the evaporation rates of these materials would be lower than those of oils and greases. These films might therefore be very attractive as lubricants for high vacuum service.}
}

Buckley, D. H. (1968). Influence of Crystal Structure, Orientation, and Solubility on Adhesion and Sliding Friction of Metal Single Crystals in Vacuum . NASA Lewis Research Center, NASA TN D-4954. Source

BibTeX

@techreport{buckley1968influence,
  title = {Influence of Crystal Structure, Orientation, and Solubility on Adhesion and Sliding Friction of Metal Single Crystals in Vacuum},
  author = {Buckley, Donald H.},
  number = {NASA TN D-4954},
  institution = {NASA Lewis Research Center},
  year = {1968},
  url = {https://ntrs.nasa.gov/citations/19680011757},
  abstract = {Influence of crystal structure, orientation, and solubility on adhesion and sliding friction of metal single crystals in vacuum}
}

deGroh, K. K., Banks, B. A., Dever, J. A., Jaworske, D. A., Miller, S. K., Sechkar, E. A. and Panko, S. R. (2008). NASA Glenn Research Center's Materials International Space Station Experiments (MISSE 1-7) . International Symposium on SM/MPAC and SEED Experiments, NASA/TM-2008-215482. Source

BibTeX

@inproceedings{degroh2008nasa,
  title = {NASA Glenn Research Center's Materials International Space Station Experiments (MISSE 1-7)},
  author = {deGroh, Kim K. and Banks, Bruce a. and Dever, Joyce A. and Jaworske, Donald A. and Miller, Sharon K. and Sechkar, Edward A. and Panko, Scott R.},
  booktitle = {International Symposium on SM/MPAC and SEED Experiments},
  number = {NASA/TM-2008-215482},
  institution = {NASA},
  address = {Tsukuba},
  year = {2008},
  url = {https://ntrs.nasa.gov/citations/20090005995},
  abstract = {NASA Glenn Research Center (Glenn) has 39 individual materials flight experiments (>540 samples) flown as part of the Materials International Space Station Experiment (MISSE) to address long duration environmental durability of spacecraft materials in low Earth orbit (LEO). MISSE is a series of materials flight experiments consisting of trays, called Passive Experiment Carriers (PECs) that are exposed to the space environment on the exterior of the International Space Station (ISS). MISSE 1-5 have been successfully flown and retrieved and were exposed to the space environment from one to four years. MISSE 6A & 6B were deployed during the STS-123 shuttle mission in March 2008, and MISSE 7A & 7B are being prepared for launch in 2009. The Glenn MISSE experiments address atomic oxygen (AO) effects such as erosion and undercutting of polymers, AO scattering, stress effects on AO erosion, and in-situ AO fluence monitoring. Experiments also address solar radiation effects such as radiation induced polymer shrinkage, stress effects on radiation degradation of polymers, and radiation degradation of indium tin oxide (ITO) coatings and spacesuit fabrics. Additional experiments address combined AO and solar radiation effects on thermal control films, paints and cermet coatings. Experiments with Orion Crew Exploration Vehicle (CEV) seals and UltraFlex solar array materials are also being flown. Several experiments were designed to provide ground-facility to in-space calibration data thus enabling more accurate in-space performance predictions based on ground-laboratory testing. This paper provides an overview of Glenn s MISSE 1-7 flight experiments along with a summary of results from Glenn s MISSE 1 & 2 experiments.}
}

Dursch, H. W., Spear, W. S., Miller, E. A., Bohnhoff-Hlavacek, G. L. and Edelman, J. (1992). Analysis of systems hardware flown on LDEF. Results of the systems special investigation group . Legacy CDMS, 19920022433. Source

BibTeX

@techreport{dursch1992analysis,
  title = {Analysis of systems hardware flown on LDEF. Results of the systems special investigation group},
  author = {Dursch, Harry W. and Spear, W. Steve and Miller, Emmett A. and Bohnhoff-Hlavacek, Gail L. and Edelman, Joel},
  number = {19920022433},
  institution = {Legacy CDMS},
  year = {1992},
  url = {https://ntrs.nasa.gov/citations/19920022433},
  abstract = {The Long Duration Exposure Facility (LDEF) was retrieved after spending 69 months in low Earth orbit (LEO). LDEF carried a remarkable variety of mechanical, electrical, thermal, and optical systems, subsystems, and components. The Systems Special Investigation Group (Systems SIG) was formed to investigate the effects of the long duration exposure to LEO on systems related hardware and to coordinate and collate all systems analysis of LDEF hardware. Discussed here is the status of the LDEF Systems SIG investigation through the end of 1991.}
}