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At least 307 records · Page 17

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NASA NTRS1976

Testing of typical spacecraft materials in a simulated substorm environment

An investigation to determine spacecraft materials characteristics under charging conditions was conducted at the Lewis Research Center substorm simulation facility. The test specimens were spacecraft paints, silvered Teflon, thermal blankets and solar array segments. The samples, ranging in size from 300 to 1,000 sq cm were exposed to monoenergetic electron energies from 2 to 20 keV at a current density of 1 nA sq cm. The samples generally behaved as capacitors with strong voltage gradients at their edges. The charging characteristics of the silvered Teflon, Kapton, and solar cell covers were controlled by the secondary emission characteristics. Insulators that did not discharge were the spacecraft paints and the quartz fiber cloth thermal blanket sample. All other samples did experience discharges when the surface voltage reached -8 to -16 kV. The discharges were photographed. The breakdown voltage for each sample was determined and the average energy lost in the discharge was computed.

Stephens, N. J.↗

NASA NTRS1976

High-explosive cratering analogs for bowl-shaped, central uplift, and multiring impact craters

The paper describes six experimental explosion craters in terms of their basic morphology, subsurface structural deformation, and surrounding ejecta blanket. These craters exhibit one or more of the following features: bowl shapes with underlying breccia lens, central uplifts, multirings, terraced walls, rim strata, zones of concentric rim deformation, inner continuous ground cover of ejecta blankets formed by overturned flaps, secondary cratering, and fused alluvium. These craters were formed by large shock wave energy transfers at or near zero heights-of-burst, and it is possible that impact craters with analogous morphologic and structural features may have formed under similar surface energy transfer conditions.

Roddy, D. J.↗

NASA NTRS1977

Testing of typical spacecraft materials in a simulated substorm environment

The test specimens were spacecraft paints, silvered Teflon, thermal blankets, and solar array segments. The samples, ranging in size from 300 to 1000 sq cm were exposed to monoenergetic electron energies from 2 to 20 keV at a current density of 1 NA/sq cm. The samples generally behaved as capacitors with strong voltage gradient at their edges. The charging characteristics of the silvered Teflon, Kapton, and solar cell covers were controlled by the secondary emission characteristics. Insulators that did not discharge were the spacecraft paints and the quartz fiber cloth thermal blanket sample. All other samples did experience discharges when the surface voltage reached -8 to -16kV. The discharges were photographed. The breakdown voltage for each sample was determined and the average energy lost in the discharge was computed.

Stevens, N. J.↗

NASA NTRS1978

Design and fabrication of a radiative actively cooled honeycomb sandwich structural panel for a hypersonic aircraft

The panel assembly consisted of an external thermal protection system (metallic heat shields and insulation blankets) and an aluminum honeycomb structure. The structure was cooled to temperature 442K (300 F) by circulating a 60/40 mass solution of ethylene glycol and water through dee shaped coolant tubes nested in the honeycomb and adhesively bonded to the outer skin. Rene'41 heat shields were designed to sustain 5000 cycles of a uniform pressure of + or - 6.89kPa (+ or - 1.0 psi) and aerodynamic heating conditions equivalent to 136 kW sq m (12 Btu sq ft sec) to a 422K (300 F) surface temperature. High temperature flexible insulation blankets were encased in stainless steel foil to protect them from moisture and other potential contaminates. The aluminum actively cooled honeycomb sandwich structural panel was designed to sustain 5000 cycles of cyclic in-plane loading of + or - 210 kN/m (+ or - 1200 lbf/in.) combined with a uniform panel pressure of + or - 6.89 kPa (?1.0 psi).

Ellis, D. A.↗

NASA NTRS1978

Thermal performance of gaseous-helium-purged tank-mounted multilayer insulation system during ground-hold and space-hold thermal cycling and exposure to water vapor

An experimental investigation was conducted to determine (1) the ground-hold and space-hold thermal performance of a multilayer insulation (MLI) system mounted on a spherical, liquid-hydrogen propellant tank and (2) the degradation to the space-hold thermal performance of the insulation system that resulted from both thermal cycling and exposure to moisture. The propellant tank had a diameter of 1.39 meters (4.57ft). The MLI consisted of two blankets of insulation; each blanket contained 15 double-aluminized Mylar radiation shields separated by double silk net spacers. Nineteen tests simulating basic cryogenic spacecraft thermal (environmental) conditions were conducted. These tests typically included initial helium purge, liquid-hydrogen fill and ground-hold, ascent, space-hold, and repressurization. No significant degradation of the space-hold thermal performance due to thermal cycling was noted.

Sumner, I. E.↗

NASA NTRS1978

Alteration of glass as a possible source of clay minerals on Mars

Thermodynamic calculations show that, under present Martian surface conditions, favorable gas-solid weathering products of feldspar glasses should include beidellites (clays of the montmorillonite series) + carbonates + quartz. The gas-solid weathering of mafic silicate glass ( of volcanic or impact origin) may similarly favor the production of metastable Fe-rich montmorillonite clays. Simple mass-balance calculations suggest that gas-solid weathering of Martian proto-regolith containing 10% glass could conceivably produce a global blanket of clays at a rate of at least 0.4 cm/b.y. The production rate should be expected to increase significantly with the glass content and rate of reworking of the proto-regolith and with the availability of water. Complete extraction of altered glass from a lunar-like proto-regolith might yield a global Martian clay blanket about 10-100 cm in thickness.

Gooding, J. L.↗

NASA NTRS1978

Phobos - Photometry and origin of dark markings on crater floors

Viking Orbiter 1 close encounter pictures of Phobos reveal unusual dark patches on the floors of many craters. Photometry of these features indicates that they have a similar normal reflectance but a phase coefficient 30% larger than the average surface. These facts suggest that the 'dark' material has a similar composition but a much rougher texture than the average surface of Phobos. By analogy with terrestrial impact, explosion and laboratory craters, the dark patches are interpreted as impact generated melt pools which remain visible on Phobos due to the low value of g and consequent small amount of fallback ejecta and slumping. Since blanketing by ejecta from subsequent impacts and micrometeoroid erosion should eventually obliterate such features, they should be more conspicuous in the fresher craters. Similar features should be visible on small asteroids and other low g bodies. However, they are not prominent on the surface of Deimos - possibly because most of the craters imaged at high resolution on the outer satellite appear to have been blanketed by several meters of fine-grained material.

Goguen, J.↗

NASA NTRS1978

Ultra-low-mass flexible planar solar arrays using 50-micron-thick solar cells

A conceptual design study has been completed which has shown the feasibility of ultra-low-mass planar solar arrays with specific power of 200 watts/kilogram. The beginning of life (BOL) power output of the array designs would be 10 kW at 1 astronomical unit (AU) and a 55C deg operating temperature. Two designs were studied: a retractable rollout design and a non-retractable fold-out. The designs employed a flexible low-mass blanket and low-mass structures. The blanket utilized 2 x 2 cm high-efficiency (13.5% at 28C deg AM0), ultra-thin (50 micron), silicon solar cells protected by thin (75 micron) plastic encapsulants. The structural design utilized the 'V'-stiffened approach which allows a lower mass boom to be used. In conjunction with the conceptual design, modules using the thin cells and plastic encapsulant were designed and fabricated.

Costogue, E. N.↗

NASA NTRS1979

SEP solar array development testing

This paper describes the test program of a lightweight 25 kW solar array for solar electric propulsion. A full-scale development wing was made of aluminum with the containment box cover of graphite-epoxy, while the flight design array wing uses a graphite-epoxy structure. The full-scale continuous longeron array extension mast was tested for performance on a water table, and the full scale wing was functionally tested to demonstrate automatic containment box unlocking, wing extension, and retraction, blanket tensioning, and automatic application of blanket preload. The wing was then tested to the Shuttle acoustic environment, followed by a thermal/vacuum test in which the wing was extended and retracted at high and low temperature. Finally, the wing was tested in vibration with sine and random vibration environments.

Elms, R. V., Jr.↗

NASA NTRS1979

Asteroidal regoliths

A physical model is developed for the evolution of regoliths on small bodies and applied to the asteroid and meteorite parent bodies. The model considers global deposition of that fraction of cratering ejecta that is not lost to space. In addition, it is applied to cases of both strong, cohesive bodies and to bodies of weak, unconsolidated materials. It is found that large, strong asteroids generate surficial regoliths of a few kilometers depth while strong asteroids smaller than 10-km diameter generate negligible regoliths. In conclusion, it is noted that the theory that substantial regoliths are produced predominantly by blanketing differs from earlier hypotheses that asteroidal regoliths might be thin or absent and that short surface exposure of asteroidal materials is due chiefly to erosion rather than blanketing.

Housen, K. R.↗

NASA NTRS1980

Large solar arrays

A spectrophotovoltaic converter, a thermophotovoltaic converter, a cassegrainian concentrator, a large silicon cell blanket, and a high flux approach are among the concepts being investigated as part of the multihundred kW solar array program for reducing the cost of photovoltaic energy in space. These concepts involve a range of technology risks, the highest risk being represented by the thermophotovoltaics and spectrophotovoltaics approaches which involve manipulation to of the incoming spectrum to enhance system efficiency. The planar array (solar blanket) has no technology risk and a moderate payback. The primary characteristics, components, and technology concerns of each of these concepts are summarized. An orbital power platform mission in the late 1980's is being used to allow a coherent technology advancement program in order to achieve a ten year life with maintenance at a capital recurring cost of $30/watt based on 1978 dollars.

Crabtree, W. L.↗

NASA NTRS1979

Solar array technology developments and design for electric propulsion application

The design, fabrication, and test of foldout and rollout multikilowatt solar array models are discussed. Consideration is given to the development of (1) thin 50-micron solar cells; (2) laminated Kapton plus substrates; (3) thin 50 to 100-micron coverglass; (4) low-mass structure designs; (5) low-mass blanket designs; and (6) solar insolation concentrators to enhance solar array performance. It is noted that the recent achievements in solar cell technology have shown that blanket specific power of 120 W/kg and higher is realizable for solar arrays, thus meeting propulsion stage requirements.

Costogue, E. N.↗

NASA NTRS1980

Martian rampart and pedestal craters' ejecta-emplacement Coprates quadrangle

A study of the Martian rampart and pedestal craters in the Coprates quadrangle is presented. Relationships between crater size and ejecta-blanket areal extents imply a maximum ejecta-blanket thickness or maximum rim height for Martian rampart craters; the limiting thickness is encountered only for craters exceeding 6 km diameter. The larger craters have an additional component of internal flow of the ejecta caused by the greater weight of their thicker ejecta deposits, although smaller rampart craters have ejecta which appears to have undergone flow during emplacement; pedestal craters most likely result from impacts into less volatile-rich substrates which produce a less fluidized ejecta and no flow lobes.

Mutch, P.↗

NASA NTRS1980

Efficient structures for geosynchronous spacecraft solar arrays

Structural concepts for deploying and supporting lightweight solar array blankets for geosynchronous electrical powr are evaluated. First, the requirements for more mass efficient solar arrays is established by describing future needs. Then analytical results are set forth which show that not only must lighter weight blankets be developed but also the supporting structure must be improved proportionately. The SEPS configuration is taken to be the state of the art point of departure for improved structural concepts. Several directions for improvement are indicated.

Hedgepeth, J. M.↗

NASA NTRS1981

Summary of LSST systems analysis and integration task for SPS flight test articles

The structural and equipment requirements for two solar power satellite (SPS) test articles are defined. The first SPS concept uses a hexagonal frame structure to stabilize the array of primary tension cables configured to support a Mills Cross antenna containing 17,925 subarrays composed of dipole radiating elements and solid state power amplifier modules. The second test article consists of a microwave antenna and its power source, a 20 by 200 m array of solar cell blankets, both of which are supported by the solar blanket array support structure. The test article structure, a ladder, is comprised of two longitudinal beams (215 m long) spaced 10 m apart and interconnected by six lateral beams. The system control module structure and bridge fitting provide bending and torsional stiffness, and supplement the in plane Vierendeel structure behavior. Mission descriptions, construction, and structure interfaces are addressed.

Greenberg, H. S.↗

NASA NTRS1981

Electron penetration of spacecraft thermal insulation

The external thermal blanket with 13 mils of polyethylene which has the known range and stopping power as a function of electron energy is investiated. The most recent omnidirectional peak Jovian electron flux at 5 Jupiter radii is applied, the electron current penetrating the thermal blanket is calculated and allowed to impinge on a typical 20 mil polyethylene insulator surrounding a wire. The radiation dose rate to the insulator is then calculated and the electrical conductivity found. The results demonstrate that the increased electronic mobility is sufficient to keep the maximum induced electric field two orders of magnitude below the critical breakdown strength.

Powers, W. L.↗

NASA NTRS1981

Efficient structures for geosynchronous spacecraft solar arrays. Phase 1, 2 and 3

Structural concepts for deploying and supporting lightweight solar-array blankets for geosynchronous electrical power are evaluated. It is recommended that the STACBEAM solar-array system should be the object of further study and detailed evaluation. The STACBEAM system provides high stiffness at low mass, and with the use of a low mass deployment mechanism, full structural properties can be maintained throughout deployment. The stowed volume of the STACBEAM is acceptably small, and its linear deployment characteristic allows periodic attachments to the solar-array blanket to be established in the stowed configuration and maintained during deployment.

Adams, L. R.↗

NASA NTRS1983

More-Uniform Heat Curing for Structural Repairs

Copper lining helps to distribute heat under heating blanket. Bond during setup employs heating blanket over copper sheet within vacuum bag. Copper sheet smoothes out temperature distribution under hotspots in heater. Method applicable to uniformly heating such low-thermal-conductivity metals as titanium or stainless steel.

Bauer, P. E.↗