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In blanket of insulating material, layers of aluminized Mylar polyethylene terephthalate separated by spacers cause mutliple reflections of thermal radiation and impede passage of radiation through blanket. Insulating quality of blanket proportional to number of functional aluminized radiation-shield surfaces in assembly.
Results of research performed for the purpose of studying the sensitivity of multi-layer insulation blanket performance caused by penetrations through the blanket are presented. The work described in this paper presents the experimental data obtained from thermal vacuum tests of various penetration geometries similar to those present on the Hubble Space Telescope. The data obtained from these tests is presented in terms of electrical power required sensitivity factors referenced to a multi-layer blanket without a penetration. The results of these experiments indicate that a significant increase in electrical power is required to overcome the radiation heat losses in the vicinity of the penetrations.
Design of vent reduces radiative transfer of heat. Venting system allows gases to escape while minimizing loss of heat by radiation. Prevents excessive buildup of internal pressure on thermal blanket. Polyimide box serves as modular vent. Many modules combined to suit gas-flow requirements of equipment. Indirect flow paths, external thermal control film, and thermal blankets reduce radiative loss of heat from interior. Developed to permit outgassing from thermal-blanket-covered spacecraft during passage from atmosphere of Earth to vacuum of space, venting approach adaptable to thermal insulation in laboratory vacuum systems.
During the photovoltaic array environmental protection program, a coating material, application technique, and design approach intended to protect flexible solar array blankets during a nominal fifteen-year operating lifetime were developed. Numerous thin-film coatings for protecting the Kapton polyimide material used in the construction of the Space Station Freedom flexible solar array blanket were evaluated. The critical solar array design features and protection measures are discussed with special emphasis on the effects of solar array fabrication and flexible printed circuit manufacturing processes on coating durability. The results of the mechanical and environmental test evaluation, including oxygen plasma, neutral oxygen beam, and UV/charged-particle combined exposure, are discussed. These results led to the selection of a silicon dioxide thin-film coating to protect the solar array blanket from the low-earth-orbit atomic oxygen environment.
Polysiloxane-polyimide films are of interest as a replacement for polyimide Kapton in the Space Station Freedom solar array blanket. The blanket provides the structural support for the solar cells as well as providing transport of heat away from the back of the cells. Polyimide Kapton would be an ideal material to use; however, its high rate of degradation due to attack by atomic oxygen in low Earth orbit, at the altitudes Space Station Freedom will fly, is of such magnitude that if left unprotected, the blanket will undergo structural failure in much less than the desired 15 year operating life. Polysiloxane-polyimide is of interest as a replacement material because it should from its own protective silicon dioxide coating upon exposure to atomic oxygen. Mass, optical, and photomicrographic data obtained in the evaluation of the durability of polysiloxane-polyimide to an atomic oxygen environment are presented.
Pieces of silvered Teflon thermal control blanket from seventeen locations on the Long Duration Exposure Facility (LDEF) were studied. Results of the measurements of surface texturing, and mechanical, electrical and thermal properties are presented. Photos of the blankets, scanning electron microscope pictures, infrared spectra, and surface analysis results are presented. Contamination issues relative to performance of the blankets are discussed. The variation in measured properties are reported as a function of exposure conditions.
The results are presented of a study conducted to obtain experimental heat transfer data on a liquid hydrogen tank insulated with 34 layers of MLI (multilayer insulation) for warm side boundary temperatures of 630, 530, and 150 R. The MLI system consisted of two blankets, each blanket made up of alternate layers of double silk net (16 layers) and double aluminized Mylar radiation shields (15 layers) contained between two cover sheets of Dacron scrim reinforced Mylar. The insulation system was designed for and installed on a 87.6 in diameter liquid hydrogen tank. Nominal layer density of the insulation blankets is 45 layers/in. The insulation system contained penetrations for structural support, plumbing, and electrical wiring that would be representative of a cryogenic spacecraft. The total steady state heat transfer rates into the test tank for shroud temperatures of 630, 530, 152 R were 164.4, 95.8, and 15.9 BTU/hr respectively. The noninsulation heat leaks into the tank (12 fiberglass support struts, tank plumbing, and instrumentation lines) represent between 13 to 17 pct. of the total heat input. The heat input values would translate to liquid H2 losses of 2.3, 1.3, and 0.2 pct/day, with the tank held at atmospheric pressure.
The results are presented of a study conducted to obtain experimental heat transfer data on a liquid hydrogen tank insulated with 34 layers of MLI (multilayer insulation) for warm side boundary temperatures of 630, 530, and 150 R. The MLI system consisted of two blankets, each blanket made up of alternate layers of double silk net (16 layers) and double aluminized Mylar radiation shields (15 layers) contained between two cover sheets of Dacron scrim reinforced Mylar. The insulation system was designed for and installed on an 87.6 in. diameter liquid hydrogen tank. Nominal layer density of the insulation blankets is 45 layers/in. The insulation system contained penetrations for structural support, plumbing, and electrical wiring that would be representative of a cryogenic spacecraft. The total steady state heat transfer rates into the test tank for shroud temperatures of 630, 530, 152 R were 164.4, 95.8, and 15.9 BTU/hr, respectively. The noninsulation heat leaks into the tank (12 fiberglass support struts, tank plumbing, and instrumentation lines) represent between 13 to 17 pct. of the total heat input. The heat input values would translate to liquid H2 losses of 2.3, 1.3, and 0.2 pct/day, with the tank held at atmospheric pressure.
The photovoltaic (PV) power system for Space Station Freedom (SSF) uses solar array blankets which provide structural support for the solar cells and house the electrical interconnections. In the low Earth orbital (LEO) environment where SSF will be located, surfaces will be exposed to potentially damaging environmental conditions including solar ultraviolet (UV) radiation, thermal cycling, and atomic oxygen. It is necessary to use ground based tests to determine how these environmental conditions would affect the mass loss and optical properties of candidate SSF blanket materials. Silicone containing, silicone coated, and SiO(x) coated polyimide film materials were exposed to simulated LEO environmental conditions to determine there durability and whether the environmental conditions of UV, thermal cycling and oxygen atoms act synergistically on these materials. A candidate PV blanket material called AOR Kapton, a polysiloxane polyimide cast from a solution mixture, shows an improvement in durability to oxygen atoms erosion after exposure to UV radiation or thermal cycling combined with UV radiation. This may indicate that the environmental conditions react synergistically with this material, and the damage predicted by exposure to atomic oxygen alone is more severe than that which would occur in LEO where atomic oxygen, thermal cycling and UV radiation are present together.
Laboratory impact experiments are needed to understand the relationship between a measured penetration hole diameter and associated projectile dimension in the thermal blankets of experiment A0178, which occupied some 16 sq. m. These blankets are composed of 125 micron thick Teflon that has an Ag/enconel second mirror surface, backed by organic binder and Chemglaze paint for a total thickness of some 170 microns. While dedicated experiments are required to understand the penetration behavior of this compound target in detail, we report here on impact simulations sponsored by other projects into pure Teflon and aluminum targets. These experiments will allow first order interpretations of impact features on the Long Duration Exposure Facility (LDEF), and they will serve as guides for dedicated experiments that employ the real LDEF blankets, both exposed and unexposed, for a refined understanding of the LDEF's collisional environment. We employed a light gas gun to launch soda-lime glass spheres from 50 to 3200 microns in diameter that impacted targets of variable thickness. Penetration measurements are given.
The photovoltaic (PV) power system for Space Station Freedom (SSF) uses solar array blankets which provide structural support for the solar cells and house the electrical interconnections. In the low earth orbital (LEO) environment where SSF will be located, surfaces will be exposed to potentially damaging environmental conditions including solar ultraviolet (UV) radiation, thermal cycling, and atomic oxygen. It is necessary to use ground based tests to determine how these environmental conditions would affect the mass loss and optical properties of candidate SSF blanket materials. Silicone containing, silicone coated, and SiO(x) coated polyimide film materials were exposed to simulated LEO environmental conditions to determine their durability and whether the environmental conditions of UV, thermal cycling and oxygen atoms act synergistically on these materials. A candidate PV blanket material called AOR Kapton, a polysiloxane polyimide cast from a solution mixture, shows an improvement in durability to oxygen atoms erosion after exposure to UV radiation or thermal cycling combined with UV radiation. This may indicate that the environmental conditions react synergistically with this material, and the damage predicted by exposure to atomic oxygen alone is more severe than that which would occur in LEO where atomic oxygen, thermal cycling and UV radiation are present together.
Sputter-deposited SiO(X) (where X between 1.9 and 2.0) thin film coatings have been found to be durable to atomic oxygen. Such coatings will be used to protect polyimide Kapton photovoltaic array blankets from atomic oxygen attack in low earth orbit (LEO) on the Space Station Freedom (SSF). Monte Carlo modeling of atomic oxygen attack at defect sites in protected Kapton exposed in laboratory RF plasma ashers and on solar tracking photovoltaic arrays in space has been conducted to enable understanding of degradation processes relevant to the durability of the SSF solar array blanket. SiO(X) protective coating performance data from RF plasma asher tests will be presented, along with Monte Carlo modeling considerations, to enable the projection of in-space durability of the SSF solar array blankets based on ground laboratory test results.
A study was performed to design a self-deploying tent shaped PV (photovoltaic) array and optimize the design for maximum specific power. Each structural component of the design was analyzed to determine the size necessary to withstand the various forces it would be subjected to. Through this analysis the component weights were determined. An optimization was performed to determine the array dimensions and blanket geometry which produce the maximum specific power for a given PV blanket. This optimization was performed for both Lunar and Martian environmental conditions. The performance specifications for the array at both locations and with various PV blankets were determined.
Experiments are described which were conducted at the JPL Advanced Teleoperator Lab to demonstrate and evaluate the effectiveness of various teleoperator control modes in the performance of a simulated Solar Max Satellite Repair (SMSR) task. THe SMSR was selected as a test because it is very rich in performance capability requirements and it actually has been performed by two EVA astronauts in the Space Shuttle Bay in 1984. The main subtasks are: thermal blanket removal; installation of a hinge attachment for electrical panel opening; opening of electrical panel; removal of electrical connectors; relining of cable bundles; replacement of electrical panel; securing parts and cables; re-mate electrical connectors; closing of electrical panel; and reinstating thermal blanket. The current performance experiments are limited to thermal blanket cutting, electrical panel unbolting and handling electrical bundles and connectors. In one formal experiment even different control modes were applied to the unbolting and reinsertion of electrical panel screws subtasks. The seven control modes are alternative combinations of manual position and rate control with force feedback and remote compliance referenced to force-torque sensor information. Force-torque sensor and end effector position data and task completion times were recorded for analysis and quantification of operator performance.
Among the solar cell technologies to be tested in space as part of the Solar Array Module Plasma Interactions Experiment (SAMPIE) will be the Advanced Photovoltaic Solar Array (APSA). Several prototype twelve cell coupons were built for NASA using different blanket materials and mounting techniques. The first conforms to the baseline design for APSA which calls for the cells to be mounted on a carbon loaded Kapton blanket to control charging in GEO. When deployed, this design has a flexible blanket supported around the edges. A second coupon was built with the cells mounted on Kapton-H, which was in turn cemented to a solid aluminum substrate. A final coupon was identical to the latter but used germanium coated Kapton to control atomic oxygen attack in LEO. Ground testing of these coupons in a plasma chamber showed considerable differences in plasma current collection. The Kapton-H coupon demonstrated current collection consistent with exposed interconnects and some degree of cell snapover. The other two coupons experienced anomalously large collection currents. This behavior is believed to a consequence of enhanced plasma sheaths supported by the weakly conducting carbon and germanium used in these coupons. The results reported here are the first experimental evidence that the use of such materials can result in power losses to high voltage space power systems.
In this research, we aim to determine to what extent the solar irradiance changes measured through the 1980's from orbiting vehicles are accompanied by spectroscopic irradiance changes observable from the ground. We describe fractional changes in line absorption as 'blanketing'. In section 2, we briefly review results obtained in an earlier project and which have been published. Section 3 describes the data of this investigation; section 4 describes the data reduction; section 5 describes the observational results in terms of blanketing; and section 6 describes the interpretation of the measured changes. Section 7 contains an outline of possible uses for Doppler-shift data that emerges with the measurements of the blanketing variations. Section 8 is an outline for future research suggested by our results in this project.
Utilizing the Termoskan data set of the Phobos '88 mission we have recognized a new feature on Mars: Ejecta blanket Distinct In the THermal infrared (EDITH). Virtually all of the more than one hundred of these features discovered in the Termoskan data are located on the plains near Valles Manneris. EDITH's have a startlingly clear dependence upon terrains of Hesperian age, implying a spatial or temporal dependence on Hesperian terrains. Almost no thermally distinct ejecta blankets are associated with any of the thousands of craters within the data set that occur on the older Noachian units. EDITH's also do not appear on the portions of the younger Tharsis Amazonian units seen in the data. The Hesperian terrain dependence cannot be explained by either atmospheric or impactor variations; Noachian and Hesperian terrains must have experienced identical atmospheric and impactor conditions during Hesperian times. Thermally distinct eject a blankets therefore reflect target material differences and/or secondary modification processes. A further discussion of EIDTH's is presented.
A test was conducted to determine the venting characteristics of the multiple-layer insulation (MLI) to be installed on the Space Station Freedom (SSF). A full MLI blanket with inter-blanket joints was installed onto a model of a section of the SSF pressure wall, support structure, and debris shield. Data were taken from this test and were used to predict the venting of the actual Space Station pressure-wall/MLI/debris-shield assemply during launch and possible re-entry. It was found that the pressure differences across the debris shields and MLI blankets were well within the specified limits in all cases.